Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion Systems

Esteban Valencia, Víctor Alulema and Darío Rodríguez

### Abstract

The inspection of wetlands in the Ecuadorian highlands has gained importance due to the environmental issues linked to the growth of human activities and the expansion of the agricultural and livestock frontiers. In this sense, unmanned aerial vehicles (UAVs) have been amply used in monitoring activities such as the supervision of threatened ecosystems, where cyclic measurements and high-resolution imagery are needed. However, the harsh operating conditions in the Andean highlands and sensitive ecosystem restrictions demand efficient propulsion configurations with low environmental impact. Electrical distributed propulsion (EDP) systems have surged as a forefront alternative since they offer benefits in both the propulsive and aerodynamic performance of fixed-wing UAVs. In this chapter, an EDP system is sized for a design point at the Andean operating conditions. Thereafter, two propulsion configurations were established based on off-the-shelf components, and their performance was characterized through analytical approaches. These results highlight the trends in power consumption and performance when the number of propulsors is increased. A significant contribution of this work is to exhibit important patterns in the performance of electric propulsion by using commercial components, and to set the operating limitations that can be further explored for analogous configurations in larger UAVs.

Keywords: unmanned aerial vehicles, distributed propulsion, electrical propulsion, blended wing body, wetland monitoring

#### 1. Introduction

The Andean region, which comprises paramos<sup>1</sup> [1] and wetlands, is considered a biodiversity hotspot that contains about one-sixth of earth's plant life [2, 3]. This extension of land is of great importance, since it represents the main water reservoir for major cities in Colombia, Ecuador, and Peru [4, 5]. Both wetlands and paramos

<sup>1</sup> A paramo is a Neotropical high mountain biome with a vegetation composed mainly of giant rosette plants, shrubs, and grasses.

are endangered ecosystems, and, hence, efficient and suitable monitoring solutions are urgently needed. In this way, different monitoring techniques including satellite imagery and the use of high-resolution cameras mounted on manned airplanes have been utilized. Nonetheless, the aforesaid methods are not commonly affordable because they are costly and require long setup times.

accomplish the aforesaid requirements thanks to its high-efficiency, moderate cost, and eco-friendly essence. Nonetheless, the capacity of commercial batteries remains to be an issue because of their lower energy density compared with their counter-

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

The distributed propulsion is a revolutionary technology that seeks to reduce the

noise and weight of an aircraft by means of replacing large propulsors with a moderate amount of small ones along the airframe [16] as depicted in Figure 1. This offers the possibility of increasing the propulsive efficiency because a larger propulsive area is considered, which in turn, implies a lower jet velocity. Its application on small fixed-wing UAVs has not been formally studied<sup>3</sup> [19], and, consequently, the present chapter aims to assess the performance of small UAV configurations with electric distributed propulsion. The study will focus mainly on power consumption and performance improvements to demonstrate the feasibility of employing this technology in small UAVs. Propulsive efficiency has not been considered as a figure of merit in the present study, because of the low operating speeds and the electrical propulsion system, where the use of this parameter does not capture well the improvement in the aircraft performance, as it does for turbofan engines. It is important to note that distributed propulsion may offer other numerous benefits such as the elimination of the aircraft control surfaces (thrust vectoring), flexible maintenance, decrease in noise, and reduction in aircraft weight through inlet-wing integration [20]. The study of these advantages is beyond the scope of the present work, since this chapter is aimed at setting the basic conceptual configurations and assessing their suitability for the case study. Nonetheless, these various features will be implemented in further research, where the selected conceptual

configurations will be assessed using a more holistic perspective.

of the UAV systems are further explained.

2.2 Initial sizing

Figure 1.

assessments.

53

For the assessment of suitable UAV configurations for wetland monitoring, parametric sizing and aerodynamic assessment approaches were implemented into the conceptual design stage. Then, a brief insight about the influence of electric distributed propulsion into the performance of the UAV configuration using a semiempirical approach is exposed. In the next sections, initial sizing and modelling

The design procedure starts by defining the mission requirements such as flight altitude, velocities, and payload sensors. In this sense, a precise study of wetlands and paramos demands the usage of special sensors applied in monitoring activities

such as crop scouting, precision agriculture, surveillance, and air quality

<sup>3</sup> Distributed propulsion configurations have been explored by hobbyists using mainly empirical

Difference between distributed and non-distributed propulsion in a BWB.

part, the fossil fuels [17, 18].

DOI: http://dx.doi.org/10.5772/intechopen.84402

The advent of unmanned aerial vehicles (UAVs) has encouraged periodical and low-cost management of threatened ecosystems through real-time data acquisition. The incursion of aerial platforms into forestry remote sensing [6] has had a positive impact thanks to the usage of high-resolution sensors [7, 8] to gather data regarding flora health, species inventory, or mapping in a periodic way. In this respect, multicopters have been seen as the first option for monitoring; however, their low autonomy limits the area covered per flight. Conversely, fixed-wing UAVs have been introduced to overfly larger areas. The imagery provided by these tools has been collected using different payloads, ranging from basic RGB cameras to sophisticated radars. Nevertheless, the time employed for a specific mission profile is higher when operating at the Andean highlands because of the harsh atmospheric conditions, which constrain the UAV autonomy and performance.

Commercial UAVs usually perform under sea level conditions with low wind gusts (lower than 16 m/s) and higher air density. This denotes that an improvement in some UAV subsystems is required to tailor them for high-altitude monitoring applications [9, 10]. Among the different characteristics that need to be upgraded to enhance the UAV performance, the following can be summarized: robust flight control system able to withstand the strong wind gusts (18 m/s), aerodynamic and high volumetric fuselage to store the avionics and payload, and high-efficient and eco-friendly propulsion system, which reduces energy consumption. The two latter options are linked, and, thus, their implementation into the conceptual design requires the assessment of their suitability to explore synergies for a more efficient UAV configuration.

The purpose of the present chapter is to investigate the performance of an electricpowered blended wing body (BWB)2 UAV deployed on the aforesaid ecosystems. This baseline configuration has been selected based on previous research [11], where it has been found that BWB configurations offer high volumetric efficiency while providing good aerodynamic performance [12] as a result of the elliptical lift distribution improvement over the whole airframe [13–15]. Furthermore, the BWB model facilitates the integration of different propulsion architectures, which results in a broader spectrum of configurations for distributed propulsion [16].

Regarding the power source, the electric option has been seen attractive because of the reduction of polluting gas emissions, moderate cost, lighter weight, and high reliability. In the next section, a deeper explanation about the propulsion configuration for this conceptual design is described.

#### 2. UAV design methodology

#### 2.1 Distributed electric propulsion

Reconnaissance and surveillance of endangered environments through overflight missions require short setup time, versatility, and noise mitigation. In this sense, electric propulsion has emerged as a reliable and potential solution to

<sup>2</sup> A BWB is a tailless aircraft design that integrates the fuselage and the wings through blended cross sections in a single body.

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion… DOI: http://dx.doi.org/10.5772/intechopen.84402

accomplish the aforesaid requirements thanks to its high-efficiency, moderate cost, and eco-friendly essence. Nonetheless, the capacity of commercial batteries remains to be an issue because of their lower energy density compared with their counterpart, the fossil fuels [17, 18].

The distributed propulsion is a revolutionary technology that seeks to reduce the noise and weight of an aircraft by means of replacing large propulsors with a moderate amount of small ones along the airframe [16] as depicted in Figure 1. This offers the possibility of increasing the propulsive efficiency because a larger propulsive area is considered, which in turn, implies a lower jet velocity. Its application on small fixed-wing UAVs has not been formally studied<sup>3</sup> [19], and, consequently, the present chapter aims to assess the performance of small UAV configurations with electric distributed propulsion. The study will focus mainly on power consumption and performance improvements to demonstrate the feasibility of employing this technology in small UAVs. Propulsive efficiency has not been considered as a figure of merit in the present study, because of the low operating speeds and the electrical propulsion system, where the use of this parameter does not capture well the improvement in the aircraft performance, as it does for turbofan engines.

It is important to note that distributed propulsion may offer other numerous benefits such as the elimination of the aircraft control surfaces (thrust vectoring), flexible maintenance, decrease in noise, and reduction in aircraft weight through inlet-wing integration [20]. The study of these advantages is beyond the scope of the present work, since this chapter is aimed at setting the basic conceptual configurations and assessing their suitability for the case study. Nonetheless, these various features will be implemented in further research, where the selected conceptual configurations will be assessed using a more holistic perspective.

For the assessment of suitable UAV configurations for wetland monitoring, parametric sizing and aerodynamic assessment approaches were implemented into the conceptual design stage. Then, a brief insight about the influence of electric distributed propulsion into the performance of the UAV configuration using a semiempirical approach is exposed. In the next sections, initial sizing and modelling of the UAV systems are further explained.

#### 2.2 Initial sizing

are endangered ecosystems, and, hence, efficient and suitable monitoring solutions are urgently needed. In this way, different monitoring techniques including satellite imagery and the use of high-resolution cameras mounted on manned airplanes have been utilized. Nonetheless, the aforesaid methods are not commonly affordable

The advent of unmanned aerial vehicles (UAVs) has encouraged periodical and low-cost management of threatened ecosystems through real-time data acquisition. The incursion of aerial platforms into forestry remote sensing [6] has had a positive impact thanks to the usage of high-resolution sensors [7, 8] to gather data regarding flora health, species inventory, or mapping in a periodic way. In this respect, multicopters have been seen as the first option for monitoring; however, their low autonomy limits the area covered per flight. Conversely, fixed-wing UAVs have been introduced to overfly larger areas. The imagery provided by these tools has been collected using different payloads, ranging from basic RGB cameras to sophisticated radars. Nevertheless, the time employed for a specific mission profile is higher when operating at the Andean highlands because of the harsh atmospheric

Commercial UAVs usually perform under sea level conditions with low wind gusts (lower than 16 m/s) and higher air density. This denotes that an improvement in some UAV subsystems is required to tailor them for high-altitude monitoring applications [9, 10]. Among the different characteristics that need to be upgraded to enhance the UAV performance, the following can be summarized: robust flight control system able to withstand the strong wind gusts (18 m/s), aerodynamic and high volumetric fuselage to store the avionics and payload, and high-efficient and eco-friendly propulsion system, which reduces energy consumption. The two latter options are linked, and, thus, their implementation into the conceptual design requires the assessment of their suitability to explore synergies for a more efficient

The purpose of the present chapter is to investigate the performance of an electricpowered blended wing body (BWB)2 UAV deployed on the aforesaid ecosystems. This baseline configuration has been selected based on previous research [11], where it has been found that BWB configurations offer high volumetric efficiency while providing good aerodynamic performance [12] as a result of the elliptical lift distribution improvement over the whole airframe [13–15]. Furthermore, the BWB model facilitates the integration of different propulsion architectures, which results in a

Regarding the power source, the electric option has been seen attractive because of the reduction of polluting gas emissions, moderate cost, lighter weight, and high reliability. In the next section, a deeper explanation about the propulsion configu-

Reconnaissance and surveillance of endangered environments through overflight missions require short setup time, versatility, and noise mitigation. In this sense, electric propulsion has emerged as a reliable and potential solution to

<sup>2</sup> A BWB is a tailless aircraft design that integrates the fuselage and the wings through blended cross

because they are costly and require long setup times.

UAV configuration.

Propulsion Systems

conditions, which constrain the UAV autonomy and performance.

broader spectrum of configurations for distributed propulsion [16].

ration for this conceptual design is described.

2. UAV design methodology

sections in a single body.

52

2.1 Distributed electric propulsion

The design procedure starts by defining the mission requirements such as flight altitude, velocities, and payload sensors. In this sense, a precise study of wetlands and paramos demands the usage of special sensors applied in monitoring activities such as crop scouting, precision agriculture, surveillance, and air quality

<sup>3</sup> Distributed propulsion configurations have been explored by hobbyists using mainly empirical assessments.


#### Table 1.

Monitoring sensors applied for wetlands and forestry [21–25].

monitoring. Some sensors used in these monitoring tasks are listed in Table 1. In this work, the payload's weight was assumed to be 1 kg for practical purposes

Next, the UAV layout is carried out, and main aircraft characteristics such as preliminary weight (WTOp), wing planform area (S), and preliminary power required (PRp) are delineated through the constraint analysis technique [26, 27]. This method consists of a matching plot that allows defining the design space of the aircraft depending on their performance requirements such as stall speed, maximum speed, takeoff run, and ceiling altitude [28]. The outcomes of this design stage represent the general characteristics of a preliminary aircraft architecture and will be employed to size other parts.

Afterwards, the wing shape is outlined by defining its geometrical parameters and sectional airfoils [26, 29]. The wing geometry was set according to technical and semiempirical correlations [18], and then, the obtained results were contrasted with corresponding data of commercial UAVs with similar characteristics [6]. In this way, the aerodynamic assessment was accomplished through the employment of the open-source Athena Vortex Lattice (AVL) software which incorporates the vortex lattice method (VLM) [30]. On the other hand, due to the lack of suitable analytical methods to calculate the weight of small UAV configurations, the preliminary weight was estimated as a function of the internal volume and the structural material's density of the aircraft [26].

At the end of this stage, the external shape of a conceptual model is obtained. Thereafter, it is necessary to define a proper propulsion system through the match of the thrust required and the thrust available. Finally, the weight of the resulting architecture is assessed through a refined model that takes into account the propulsion system and the power source weight. Figure 2 depicts the road map of the methodology employed to generate and characterize a conceptual BWB UAV model. It is worth to mention that all the symbols utilized along the chapter are reported in the nomenclature section at the end.

study, the aforesaid parameters were obtained from an experimental database of low Reynolds propellers [32] by using the advance ratio (J) as the key driver for the selection routine. This latter parameter relates the freestream velocity, the propeller diameter (ϕprop), and its rotational speed (ω). For this work, the freestream velocity was set according to the desired cruise speed. In this way, the proposed method involves an iterative scheme that consists of estimating the thrust and power generated by a preselected propeller through the variation of ϕprop and ω for the desired freestream velocity (Vc). The iterative loop stops when the thrust and power required are met by a certain configuration. Finally, these results were used to select appropriate electric motors and batteries which can adapt well to the design requirements. This semiempirical scheme was preferred since most of the available techniques [34, 35] are focused on large propeller assessment and, hence, they present limitations for their implementation into small aerial platforms.

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

Figure 2.

55

General methodology of initial aircraft sizing.

DOI: http://dx.doi.org/10.5772/intechopen.84402

Brushless electric motors are commonly employed for small UAVs considering their simple design, potential to downsize, little maintenance, and independent performance of the flight altitude. In addition, their purely inductive nature and their outrunner configuration (rotor with magnets that surrounds the fixed coils of the stator) enable them to generate high torque at a low rotational speed, eliminating the need of a gearbox and facilitating their integration and test at early stages of UAV design [18, 36]. In this context, appropriate motors were outlined based solely on basic parameters provided by manufacturers like rotational speed and torque. The selected motor must be able to generate the torque required by the propeller for its adequate functioning at a certain rotational speed [37]. Once a motor has been

#### 2.3 Propulsion modeling

The main function of aircraft propulsion systems is to generate enough thrust to overcome the drag and maintain a steady flight. For this work, firstly, suitable propellers were selected based on operating conditions and performance requirements of the UAV model. Then, the rest of propulsion elements (motor, electronic speed control and battery) was outlined based on propeller's characteristics. Finally, the established propulsion set is evaluated to verify that both the power and the thrust available satisfy the requirements for cruise condition. It is important to mention that, at this conceptual stage, the distortion and momentum drag reduction of the incoming flow to the propeller has not been considered and will be studied in future work.

The propellers are commonly characterized by the thrust (CT) and power (CP) coefficients through semiempirical models at early stages of design [31]. For this

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion… DOI: http://dx.doi.org/10.5772/intechopen.84402

Figure 2. General methodology of initial aircraft sizing.

study, the aforesaid parameters were obtained from an experimental database of low Reynolds propellers [32] by using the advance ratio (J) as the key driver for the selection routine. This latter parameter relates the freestream velocity, the propeller diameter (ϕprop), and its rotational speed (ω). For this work, the freestream velocity was set according to the desired cruise speed. In this way, the proposed method involves an iterative scheme that consists of estimating the thrust and power generated by a preselected propeller through the variation of ϕprop and ω for the desired freestream velocity (Vc). The iterative loop stops when the thrust and power required are met by a certain configuration. Finally, these results were used to select appropriate electric motors and batteries which can adapt well to the design requirements. This semiempirical scheme was preferred since most of the available techniques [34, 35] are focused on large propeller assessment and, hence, they present limitations for their implementation into small aerial platforms.

Brushless electric motors are commonly employed for small UAVs considering their simple design, potential to downsize, little maintenance, and independent performance of the flight altitude. In addition, their purely inductive nature and their outrunner configuration (rotor with magnets that surrounds the fixed coils of the stator) enable them to generate high torque at a low rotational speed, eliminating the need of a gearbox and facilitating their integration and test at early stages of UAV design [18, 36]. In this context, appropriate motors were outlined based solely on basic parameters provided by manufacturers like rotational speed and torque. The selected motor must be able to generate the torque required by the propeller for its adequate functioning at a certain rotational speed [37]. Once a motor has been

monitoring. Some sensors used in these monitoring tasks are listed in Table 1. In this work, the payload's weight was assumed to be 1 kg for practical purposes Next, the UAV layout is carried out, and main aircraft characteristics such as preliminary weight (WTOp), wing planform area (S), and preliminary power required (PRp) are delineated through the constraint analysis technique [26, 27]. This method consists of a matching plot that allows defining the design space of the aircraft depending on their performance requirements such as stall speed, maximum speed, takeoff run, and ceiling altitude [28]. The outcomes of this design stage represent the general characteristics of a preliminary aircraft architecture and

Main camera Mass [g] Resolution [MP]

Logitech C510 225 8 Canon PowerShot S60 230 5 Kodak Professional DCS Pro Back 770 16 Sony DSC-R1 929 10.3

Afterwards, the wing shape is outlined by defining its geometrical parameters and sectional airfoils [26, 29]. The wing geometry was set according to technical and semiempirical correlations [18], and then, the obtained results were contrasted with corresponding data of commercial UAVs with similar characteristics [6]. In this way, the aerodynamic assessment was accomplished through the employment of the open-source Athena Vortex Lattice (AVL) software which incorporates the vortex lattice method (VLM) [30]. On the other hand, due to the lack of suitable analytical methods to calculate the weight of small UAV configurations, the preliminary weight was estimated as a function of the internal volume and the struc-

At the end of this stage, the external shape of a conceptual model is obtained. Thereafter, it is necessary to define a proper propulsion system through the match of the thrust required and the thrust available. Finally, the weight of the resulting architecture is assessed through a refined model that takes into account the propulsion system and the power source weight. Figure 2 depicts the road map of the methodology employed to generate and characterize a conceptual BWB UAV model. It is worth to mention that all the symbols utilized along the chapter are

The main function of aircraft propulsion systems is to generate enough thrust to overcome the drag and maintain a steady flight. For this work, firstly, suitable propellers were selected based on operating conditions and performance requirements of the UAV model. Then, the rest of propulsion elements (motor, electronic speed control and battery) was outlined based on propeller's characteristics. Finally, the established propulsion set is evaluated to verify that both the power and the thrust available satisfy the requirements for cruise condition. It is important to mention that, at this conceptual stage, the distortion and momentum drag reduction of the incoming flow to the propeller has not been considered and will be studied in future work. The propellers are commonly characterized by the thrust (CT) and power (CP) coefficients through semiempirical models at early stages of design [31]. For this

will be employed to size other parts.

Monitoring sensors applied for wetlands and forestry [21–25].

Table 1.

Propulsion Systems

tural material's density of the aircraft [26].

reported in the nomenclature section at the end.

2.3 Propulsion modeling

54

selected, various operating parameters like no-load current, voltage constant, and internal resistance, together with torque and rotational speed that were taken from the datasheets, were employed to estimate the required voltage (Um) and current (Im) of the motor [36].

2.4 Performance evaluation

DOI: http://dx.doi.org/10.5772/intechopen.84402

propulsion system and its power source [35].

PR ¼

ffiffiffiffiffiffiffiffiffiffiffiffi 2 ρalt ∙ S

WTO <sup>3</sup>=<sup>2</sup> � � CD

The power available (PA), which depends of the propeller, motor, and battery characteristics, was estimated through analytical relationships regarding nondimensional coefficients (CT, CP, and J) [35]. The computed value of PA was verified to be greater or equal to PR in order to guarantee that the aircraft reaches the absolute ceiling altitude<sup>4</sup> at the desired rate of climb as explained before. Note that PR will be less for the cruise condition because the airplane is not climbing anymore and, thus, the excess of power is zero. Finally, for the distributed propulsion case, the total power available is estimated by multiplying the number of propellers by

It is important to highlight that the battery must provide a greater power than PA to account for energy losses as shown in Eq. (4), where ηprop represents the propeller efficiency, η<sup>e</sup> is the efficiency of the electric set (motor, electronic speed driver

The endurance and range are the key performance parameters because they reflect the time and distance that the aircraft is able to fly without recharging. Their

For this case, a simplified but good enough approach has been employed to estimate the endurance and range [36]. This method assumes that the voltage

<sup>4</sup> This term is referred to the absolute maximum altitude that the aircraft can ever maintain level flight,

estimation for electric-powered airplanes through analytical models has not received special attention because the devices for the efficient energy storage are still under development and research. Nevertheless, few authors have introduced distinct and elaborated methods to predict the aforesaid parameters from aerody-

C<sup>3</sup>=<sup>2</sup> L

PA ¼ PR þ RC∙WTO ½ � W (3)

PA ¼ ηpropηePbat ½ � W (4)

½ � W (2)

!

s

and battery), and Pbat is the power supplied by the battery:

namic characteristics and battery working conditions [38, 41].

climb (RC) as Eq. (3) shows [35]:

their generated power, respectively.

i.e., the RC is zero [26].

57

The performance analysis is an engineering discipline that relies on inputs from aerodynamic and propulsion assessments. In this sense, the performance evaluation aims to verify if a propulsion set (battery, motor and propeller) meets the mission requirements such as endurance and range. For this purpose, both the power required (PR) and the power available (PA) are determined. The former depends of the weight and the aerodynamic efficiency of an aircraft, while PA depends on the

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

The power required (PR) is calculated by means of Eq. (2) [34, 35], where ρalt is the air density at a desired altitude, S is the planform wing area, WTO is the takeoff gross weight, and CD and CL are the drag and lift aerodynamic coefficients. Note that WTO and S were previously defined in the initial sizing phase through the constraint analysis and the aerodynamic coefficients were estimated through the employment of the AVL software and parametric characterization [40]. This term represents the power required for steady cruise condition. However, the target power available (PA) must be greater than PR to consider a more demanding flight condition such as the takeoff phase. This excess of power is linked to the rate of

The motor current (Im) is then employed to select a proper electronic speed control (ESC) device and a lithium-polymer (LiPo) battery. It is important to highlight that batteries for small UAVs are almost exclusively lithium-based because they offer high capacity, low weight, and high discharge rates [38]. For the battery selection, two different scenarios were explored in this work. The first consisted of defining a nominal battery capacity based on commercial off-the-shelf devices to estimate the flight endurance. The second scenario aims to determine a suitable battery by giving a target endurance. This latter approach was employed to assess the maximum endurance that could be achieved by the UAV, without the constraints of off-the-shelf electronic components. Figure 3 illustrates the road map to establish the electric propulsion system during the conceptual design stage.

Once the propulsion system and the aircraft external shape have been framed, it is necessary to estimate the UAV total weight in a more refined and accurate way. This value is then contrasted with the admissible weight stated in the design requirements. Since typical procedures are focused on civil aviation, their application cannot be extended to small aerial platforms. Instead, this work proposes a method that individually accounts for the airframe, propulsion system, battery, and payload weights and then adds each contribution to obtain the total weight as stated in Eq. (1).

The structural weight of the airframe was calculated with respect to the fuselage internal volume and the material's density. The former was estimated through the convex hull method [39], and high-density foam was assumed as the major airframe material [26]. The weight of remaining components from Eq. (1) was readily obtained from manufacturer's datasheets.

$$\mathbf{W}\_{\rm TO} = \mathbf{W}\_{\rm air}\mathbf{r}\mathbf{r}\mathbf{e} + \mathbf{W}\_{\rm propulsion} + \mathbf{W}\_{\rm payload} + \mathbf{W}\_{\rm battery}\mathbf{[N]}\tag{1}$$

Figure 3. Electric propulsion definition and performance assessment methodology.

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion… DOI: http://dx.doi.org/10.5772/intechopen.84402

#### 2.4 Performance evaluation

selected, various operating parameters like no-load current, voltage constant, and internal resistance, together with torque and rotational speed that were taken from the datasheets, were employed to estimate the required voltage (Um) and current

The motor current (Im) is then employed to select a proper electronic speed control (ESC) device and a lithium-polymer (LiPo) battery. It is important to highlight that batteries for small UAVs are almost exclusively lithium-based because they offer high capacity, low weight, and high discharge rates [38]. For the battery selection, two different scenarios were explored in this work. The first consisted of defining a nominal battery capacity based on commercial off-the-shelf devices to estimate the flight endurance. The second scenario aims to determine

a suitable battery by giving a target endurance. This latter approach was

employed to assess the maximum endurance that could be achieved by the UAV, without the constraints of off-the-shelf electronic components. Figure 3 illustrates the road map to establish the electric propulsion system during the conceptual

Once the propulsion system and the aircraft external shape have been framed, it is necessary to estimate the UAV total weight in a more refined and accurate way. This value is then contrasted with the admissible weight stated in the design requirements. Since typical procedures are focused on civil aviation, their application cannot be extended to small aerial platforms. Instead, this work proposes a method that individually accounts for the airframe, propulsion system, battery, and payload weights and then adds each contribution to obtain the total weight as stated

The structural weight of the airframe was calculated with respect to the fuselage internal volume and the material's density. The former was estimated through the convex hull method [39], and high-density foam was assumed as the major airframe material [26]. The weight of remaining components from Eq. (1) was readily

WTO ¼ Wairframe þ Wpropulsion þ Wpayload þ Wbattery½ � N (1)

(Im) of the motor [36].

Propulsion Systems

design stage.

in Eq. (1).

Figure 3.

56

obtained from manufacturer's datasheets.

Electric propulsion definition and performance assessment methodology.

The performance analysis is an engineering discipline that relies on inputs from aerodynamic and propulsion assessments. In this sense, the performance evaluation aims to verify if a propulsion set (battery, motor and propeller) meets the mission requirements such as endurance and range. For this purpose, both the power required (PR) and the power available (PA) are determined. The former depends of the weight and the aerodynamic efficiency of an aircraft, while PA depends on the propulsion system and its power source [35].

The power required (PR) is calculated by means of Eq. (2) [34, 35], where ρalt is the air density at a desired altitude, S is the planform wing area, WTO is the takeoff gross weight, and CD and CL are the drag and lift aerodynamic coefficients. Note that WTO and S were previously defined in the initial sizing phase through the constraint analysis and the aerodynamic coefficients were estimated through the employment of the AVL software and parametric characterization [40]. This term represents the power required for steady cruise condition. However, the target power available (PA) must be greater than PR to consider a more demanding flight condition such as the takeoff phase. This excess of power is linked to the rate of climb (RC) as Eq. (3) shows [35]:

$$P\_R = \sqrt{\frac{2}{\rho\_{alt} \bullet \mathcal{S}}} \left( W\_{TO}^{-3/2} \right) \left( \frac{\mathcal{C}\_D}{\mathcal{C}\_L^{3/2}} \right) \left[ W \right] \tag{2}$$

$$P\_A = P\_R + RC \bullet W\_{TO} \text{ [W]} \tag{3}$$

The power available (PA), which depends of the propeller, motor, and battery characteristics, was estimated through analytical relationships regarding nondimensional coefficients (CT, CP, and J) [35]. The computed value of PA was verified to be greater or equal to PR in order to guarantee that the aircraft reaches the absolute ceiling altitude<sup>4</sup> at the desired rate of climb as explained before. Note that PR will be less for the cruise condition because the airplane is not climbing anymore and, thus, the excess of power is zero. Finally, for the distributed propulsion case, the total power available is estimated by multiplying the number of propellers by their generated power, respectively.

It is important to highlight that the battery must provide a greater power than PA to account for energy losses as shown in Eq. (4), where ηprop represents the propeller efficiency, η<sup>e</sup> is the efficiency of the electric set (motor, electronic speed driver and battery), and Pbat is the power supplied by the battery:

$$P\_A = \eta\_{prop} \eta\_e P\_{bat} \text{ [W]} \tag{4}$$

The endurance and range are the key performance parameters because they reflect the time and distance that the aircraft is able to fly without recharging. Their estimation for electric-powered airplanes through analytical models has not received special attention because the devices for the efficient energy storage are still under development and research. Nevertheless, few authors have introduced distinct and elaborated methods to predict the aforesaid parameters from aerodynamic characteristics and battery working conditions [38, 41].

For this case, a simplified but good enough approach has been employed to estimate the endurance and range [36]. This method assumes that the voltage

<sup>4</sup> This term is referred to the absolute maximum altitude that the aircraft can ever maintain level flight, i.e., the RC is zero [26].

remains constant and the battery capacity is decreased linearly. In this sense, Eq. (5) was used to calculate the endurance at cruise condition, where Cmin represents the battery minimum capacity that can be reached in a safety margin and Ib is the battery current. The former was supposed to be 20% of the total battery capacity because lithium-based batteries can be damaged if discharged more than 80% [42]. On the other hand, Ib is a function of the motor current, avionics current, and internal resistance. Its calculation is further explained in Ref. [36], and, hence, it will not be addressed in this work. The numeric value (0.06) in Eq. (5) represents a unit conversion factor because the capacity of batteries is commonly given in milliamperes-hour, Ib in amperes, and the computed time is given in minutes. It is important to highlight that only a single battery device was considered and its number of cells was determined based on the voltage required by the motor. The range was calculated, by using the cruise speed the endurance through the assumption of a rectilinear displacement:

$$E = 0.06 \left( \frac{C\_b - C\_{min}}{I\_b} \right) \text{[min]} \tag{5}$$

3.2 Conceptual model

Constraint analysis to determine multiple design points.

DOI: http://dx.doi.org/10.5772/intechopen.84402

Figure 4.

59

The constraint analysis illustrated in Figure 4 shows the variation of the weightto-power ratio (W/P) with respect to the wing loading (W/S) for various performance parameters such as the maximum speed (Vmax), stall speed (Vs), and ceiling altitude (hc) [26, 28]. The intersection of these curves enables to define an acceptable region of design and hence, establish a starting point for the aircraft initial sizing. It is worth to mention that for a propeller-driven aircraft, the acceptable region is located below the set of the aforesaid curves. A higher relation W/P is beneficial because this yields the smallest electric motor in terms of power required [26]; however, this is limited by the stall speed for each case. For this work, as the maximum velocity (Vmax) function depends on the air density, a curve for each flight altitude analyzed (Table 2) was drawn. Thus, a total of four plots were sketched, involving two values of Vmax at two different flight altitudes (hc).

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

Two design points were defined within a permissible range of wing loading for small UAVs [27, 44] while maintaining a weight-to-power ratio as high as possible. The W/S range was established by considering the strong linkage between the structural and aerodynamic behavior of the airplane. These design points will allow to investigate the variation of wing planform area and propulsion power for different design requirements and operating conditions (Table 2) to contrast distinct scenarios. Note that the propulsion power computed through the constraint analysis (Table 3) is a preliminary estimate. A more accurate value is calculated once the

Afterwards, the entire airframe (wing and fuselage) was outlined through classical methods of fixed-wing aircraft design [26]. Major geometrical parameters were obtained with respect to the preliminary weight (WTOp) and wing planform

geometrical parameters and the aerodynamic coefficients are defined.

area (S), previously computed in the constraint analysis. Moreover, nondimensional parameters such as wing aspect ratio (AR) and taper ratio<sup>5</sup> (λ) were initially established with regard to similar UAV architectures [12] and permissible ranges for this application [26]. The airfoil selected was the NACA 64A210 because

<sup>5</sup> The taper ratio (λ) is defined as the ration between the tip chord and the root chord of a wing.

#### 3. Results and discussion

In this section, the methodology previously explained is implemented to set the propulsion configurations for wetland monitoring at the Andean highlands. The obtained results from the UAV conceptual design, aerodynamic assessment, electric propulsion evaluation, and performance analysis are presented.

#### 3.1 Case study

The case study was carried out for wetlands located between 3500 and 4500 masl in the Antisana volcano region from Ecuador. Table 2 enlists some of the key operating conditions and design requirements that the aerial platform needs to fulfill in order to ensure a successful performance. Both cruise and stall speeds were set to ensure good-quality data gathering [43]. The payload mass was set not to exceed 1 kg after making a brief survey of common sensors (Table 1) that UAVs employ for forest monitoring [6]. These data will serve to determine the starting design point by means of the constraint analysis. It is important to mention that the maximum speed was set to be 1.25 times the cruise speed [26].


#### Table 2.

Operating conditions and design requirements.

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion… DOI: http://dx.doi.org/10.5772/intechopen.84402

Figure 4.

remains constant and the battery capacity is decreased linearly. In this sense, Eq. (5) was used to calculate the endurance at cruise condition, where Cmin represents the battery minimum capacity that can be reached in a safety margin and Ib is the battery current. The former was supposed to be 20% of the total battery capacity because lithium-based batteries can be damaged if discharged more than 80% [42]. On the other hand, Ib is a function of the motor current, avionics current, and internal resistance. Its calculation is further explained in Ref. [36], and, hence, it will not be addressed in this work. The numeric value (0.06) in Eq. (5) represents a unit conversion factor because the capacity of batteries is commonly given in milliamperes-hour, Ib in amperes, and the computed time is given in minutes. It is important to highlight that only a single battery device was considered and its number of cells was determined based on the voltage required by the motor. The range was calculated, by using the cruise speed the endurance through the assump-

<sup>E</sup> <sup>¼</sup> <sup>0</sup>:<sup>06</sup> Cb � Cmin

The case study was carried out for wetlands located between 3500 and 4500 masl in the Antisana volcano region from Ecuador. Table 2 enlists some of the key operating conditions and design requirements that the aerial platform needs to fulfill in order to ensure a successful performance. Both cruise and stall speeds were set to ensure good-quality data gathering [43]. The payload mass was set not to exceed 1 kg after making a brief survey of common sensors (Table 1) that UAVs employ for forest monitoring [6]. These data will serve to determine the starting design point by means of the constraint analysis. It is important to mention that the

Parameter Lower bound Upper bound Cruise speed [m/s] 18 22 Maximum speed [m/s] 22 26 Stall speed [m/s] 10 15 Absolute ceiling [m] 3500 4500 Aspect ratio\* 4.5 5.5 Payload mass [kg] 1 1 \*The aspect ratio (AR) is defined as the ratio between the wing span and its mean aerodynamic chord [33].

propulsion evaluation, and performance analysis are presented.

maximum speed was set to be 1.25 times the cruise speed [26].

Ib 

In this section, the methodology previously explained is implemented to set the propulsion configurations for wetland monitoring at the Andean highlands. The obtained results from the UAV conceptual design, aerodynamic assessment, electric

½ � min (5)

tion of a rectilinear displacement:

3. Results and discussion

3.1 Case study

Propulsion Systems

Table 2.

58

Operating conditions and design requirements.

#### Constraint analysis to determine multiple design points.

#### 3.2 Conceptual model

The constraint analysis illustrated in Figure 4 shows the variation of the weightto-power ratio (W/P) with respect to the wing loading (W/S) for various performance parameters such as the maximum speed (Vmax), stall speed (Vs), and ceiling altitude (hc) [26, 28]. The intersection of these curves enables to define an acceptable region of design and hence, establish a starting point for the aircraft initial sizing. It is worth to mention that for a propeller-driven aircraft, the acceptable region is located below the set of the aforesaid curves. A higher relation W/P is beneficial because this yields the smallest electric motor in terms of power required [26]; however, this is limited by the stall speed for each case. For this work, as the maximum velocity (Vmax) function depends on the air density, a curve for each flight altitude analyzed (Table 2) was drawn. Thus, a total of four plots were sketched, involving two values of Vmax at two different flight altitudes (hc).

Two design points were defined within a permissible range of wing loading for small UAVs [27, 44] while maintaining a weight-to-power ratio as high as possible. The W/S range was established by considering the strong linkage between the structural and aerodynamic behavior of the airplane. These design points will allow to investigate the variation of wing planform area and propulsion power for different design requirements and operating conditions (Table 2) to contrast distinct scenarios. Note that the propulsion power computed through the constraint analysis (Table 3) is a preliminary estimate. A more accurate value is calculated once the geometrical parameters and the aerodynamic coefficients are defined.

Afterwards, the entire airframe (wing and fuselage) was outlined through classical methods of fixed-wing aircraft design [26]. Major geometrical parameters were obtained with respect to the preliminary weight (WTOp) and wing planform area (S), previously computed in the constraint analysis. Moreover, nondimensional parameters such as wing aspect ratio (AR) and taper ratio<sup>5</sup> (λ) were initially established with regard to similar UAV architectures [12] and permissible ranges for this application [26]. The airfoil selected was the NACA 64A210 because

<sup>5</sup> The taper ratio (λ) is defined as the ration between the tip chord and the root chord of a wing.


Table 3.

Initial sizing parameters obtained from the constraint analysis.

it has proven to be suitable for small BWB models [45]. No twist angle was considered for this study. Main parameters from the initial sizing stage are summarized in Table 3.

The geometrical parameters from Table 3 were employed to generate a threedimensional shape of the proposed airframes as depicted in Figure 5. This resulted in two different models whose primary difference is the size.

#### 3.3 Aerodynamic assessment

The aerodynamic coefficients from both design concepts were estimated using the AVL open-source software. For this aim, these coefficients were obtained regarding the variation of the attack (α), that is the angle between the freestream velocity vector and the flight path as shown in Figure 6.

The drag polar obtained from the aerodynamic assessment of both conceptual UAV models (Table 3 and Figure 5) is depicted in Figure 7. In this sense, the left plot illustrates the variation of both lift (CL) and drag (CD) coefficients with respect to the angle of attack. It is important to highlight that the AVL software employs the Vortex Lattice Theory to predict the aerodynamic coefficients and, thus, it does not predict the stall behavior at high angles of attack. In addition, this code does not calculate the zero-lift drag coefficient, and, therefore, the total drag coefficient was estimated through semiempirical methods [40, 46].

On the other hand, the right plot of Figure 7 depicts the aerodynamic relation lift-to-drag (L/D) as a function of the angle of attack. This relation is a common way of reflecting the aerodynamic efficiency of an aircraft as a function of geometrical configuration and flight conditions. Moreover, this relation is highly important because it directly impacts on the endurance and range. The maximum L/D ratios obtained for cases A and B are 15.5 and 15.8, respectively. For this case, the minimum drag condition has been set for the cruise flight regime [35]. Finally, as seen in Figure 7, an L/D ratio of 8 at cruise condition is expected for both cases, which is in

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

On the other hand, the target power available, estimated by means of Eq. (3), was 302 W for both cases. This value is the same for cases A and B because their weight, aerodynamic coefficients, and operating conditions are almost equal, and hence an equivalent behavior is expected. Note that the estimated value through Eq. (3) is greater than the preliminary power required computed in the constraint analysis (Table 3) because during the initial sizing stage, several parameters were assumed based on historical and statistical data. Therefore, the power required calculated with Eq. (3) is more accurate because it considers the real performance

To investigate the integration of electric distributed propulsion into small UAV

concepts, the available space on the trailing edge of the fuselage, as shown in

accordance with commercial UAV models [6].

Angle of attack between the freestream and a half wing.

DOI: http://dx.doi.org/10.5772/intechopen.84402

and geometrical characteristics of the conceptual UAV.

3.4 Propulsion modelling

61

Figure 6.

Figure 7.

Drag polars of UAV conceptual models.

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion… DOI: http://dx.doi.org/10.5772/intechopen.84402

Figure 6.

it has proven to be suitable for small BWB models [45]. No twist angle was considered for this study. Main parameters from the initial sizing stage are summarized in

Parameter Design point 1 case A Design point 2 case B

Power loading [N/W] 0.145 0.141

Wingspan [m] 1.60 0.99 Aspect ratio 4.96 5.22 Preliminary WTO [N] 33 20 Preliminary power required [W] 224 244

] 87.97 165.37

] 0.52 0.19

The geometrical parameters from Table 3 were employed to generate a threedimensional shape of the proposed airframes as depicted in Figure 5. This resulted

The aerodynamic coefficients from both design concepts were estimated using

The drag polar obtained from the aerodynamic assessment of both conceptual UAV models (Table 3 and Figure 5) is depicted in Figure 7. In this sense, the left plot illustrates the variation of both lift (CL) and drag (CD) coefficients with respect to the angle of attack. It is important to highlight that the AVL software employs the Vortex Lattice Theory to predict the aerodynamic coefficients and, thus, it does not predict the stall behavior at high angles of attack. In addition, this code does not calculate the zero-lift drag coefficient, and, therefore, the total drag coefficient was

the AVL open-source software. For this aim, these coefficients were obtained regarding the variation of the attack (α), that is the angle between the freestream

in two different models whose primary difference is the size.

Initial sizing parameters obtained from the constraint analysis.

velocity vector and the flight path as shown in Figure 6.

estimated through semiempirical methods [40, 46].

Graphical representation of BWB airframe for two design cases (dimension in mm).

Table 3.

Figure 5.

60

Table 3.

Wing loading [N/m<sup>2</sup>

Propulsion Systems

Reference area [m<sup>2</sup>

3.3 Aerodynamic assessment

Angle of attack between the freestream and a half wing.

#### Figure 7.

Drag polars of UAV conceptual models.

On the other hand, the right plot of Figure 7 depicts the aerodynamic relation lift-to-drag (L/D) as a function of the angle of attack. This relation is a common way of reflecting the aerodynamic efficiency of an aircraft as a function of geometrical configuration and flight conditions. Moreover, this relation is highly important because it directly impacts on the endurance and range. The maximum L/D ratios obtained for cases A and B are 15.5 and 15.8, respectively. For this case, the minimum drag condition has been set for the cruise flight regime [35]. Finally, as seen in Figure 7, an L/D ratio of 8 at cruise condition is expected for both cases, which is in accordance with commercial UAV models [6].

On the other hand, the target power available, estimated by means of Eq. (3), was 302 W for both cases. This value is the same for cases A and B because their weight, aerodynamic coefficients, and operating conditions are almost equal, and hence an equivalent behavior is expected. Note that the estimated value through Eq. (3) is greater than the preliminary power required computed in the constraint analysis (Table 3) because during the initial sizing stage, several parameters were assumed based on historical and statistical data. Therefore, the power required calculated with Eq. (3) is more accurate because it considers the real performance and geometrical characteristics of the conceptual UAV.

#### 3.4 Propulsion modelling

To investigate the integration of electric distributed propulsion into small UAV concepts, the available space on the trailing edge of the fuselage, as shown in

#### Figure 8.

Main components of propeller geometry.


#### Table 4.

Thrust required for distributed propulsion.

Figures 1 and 6 has been considered. This allows estimating a suitable diameter for the propeller Φprop (Figure 8) and setting the quantity of motors that fits adequately the available space and generate the power needed.

The thrust that each propeller must generate, in conjunction with their diameter, is presented in Table 4 for both cases (A and B) with three distinct configurations: single and distributed propulsion with double and triple propulsors. Note that when the propulsion system possesses more than one propulsor, the thrust required per propeller (Treq,i) is obtained by dividing the total thrust required (Treq) by the number of installed propellers. Besides, it is important to mention that the term Treq was obtained by dividing the power required for the desired value for cruise speed (15 m/s for both cases).

#### 3.4.1 Propeller selection

Tables 5 and 6 summarize the performance characteristics of selected propellers for cases A and B, respectively. The suitable models were established regarding the freestream velocity, their diameter, and the rotational speed. Special attention was paid to guarantee that the considered propeller arrangements fit the available space and, at the same time, generate the thrust required. The cruise speed was set for minimum drag condition [35].

number of propellers implemented. Note that the shaft power and the battery power will be the same for a single propeller configuration because only one motor is considered and the electrical efficiency losses were neglected for practical

Np 1 23

DOI: http://dx.doi.org/10.5772/intechopen.84402

Np 1 2

Propeller 12 6.5 11 7 6 4 6 3 RPM 103 <sup>9</sup> <sup>10</sup> <sup>25</sup> <sup>23</sup> J 0.326 0.33 0.23 0.26 CT 0.0979 0.1036 0.0949 0.1166 CP 0.0544 0.0657 0.0473 0.0627 η<sup>P</sup> 0.58 0.52 0.46 0.48 Pshaft [W] 589.1 634.5 344.5 355.7 PA,i [W] 335.8 329.9 158.5 170.7 PA [W] 335.8 329.9 316.9 341.5 Q [Nm] 0.627 0.605 0.131 0.148 Ti [N] 23.1 21.5 10.8 10.9 T [N] 23.1 21.5 21.6 21.8 Pbat [kW] 0.589 0.634 0.689 0.711

Propeller 17 10 18 10 8 6 8 10 5 3 5 5 RPM 103 <sup>5</sup> <sup>5</sup> <sup>13</sup> <sup>12</sup> <sup>27</sup> <sup>25</sup> J 0.44 0.42 0.36 0.4 0.26 0.3 CT 0.069 0.061 0.1078 0.1195 0.1145 0.1213 CP 0.0415 0.0353 0.068 0.095 0.0642 0.0851 η<sup>P</sup> 0.73 0.72 0.57 0.5 0.47 0.42 Pshaft [W] 432.5 499.6 294.5 320.6 238.1 249.1 PA,i [W] 315.7 359.7 167.9 160.3 111.9 104.6 PA [W] 315.7 359.7 335.8 320.6 335.7 313.8 Q [Nm] 0.831 0.954 0.216 0.257 0.083 0.095 Ti [N] 20.2 22.7 10.5 10.1 7.4 6.7 T [N] 20.2 22.7 21 20.2 22.2 20.1 Pbat [kW] 0.432 0.499 0.589 0.641 0.714 0.747

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

Note that, in Tables 5 and 6, the propeller efficiency is dramatically affected as its diameter is reduced. As observed, the lesser the propeller's size, the higher the rotational speed is needed to generate the required thrust. For instance, smaller models (e.g., three propellers 5 3) must operate at a high RPM (27,000) to produce the thrust required (Table 4). In contrast, a single larger propeller (e.g., a propeller 17 10) operates at lower RPM (5000) to generate the same thrust. These

purposes.

63

Table 6.

Propeller selection—case B.

Table 5.

Propeller selection—case A.

For each propulsion set, several parameters from propellers, such as: nondimensional coefficients (J, CT, and CP), efficiency, torque, shaft power, and thrust generated were extracted from the experimental database at desired operating conditions. Notice that the power available of individual propellers (PA,i) was obtained by multiplying the power shaft for its corresponding propeller efficiency (ηp). Likewise, the total power available (PA) is a function of the number or propellers (Np) implemented, and it was obtained by multiplying Np for its corresponding individual power (PA,i). Finally, the power that battery must supply to the propulsion system was obtained by multiplying the power shaft for the

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion… DOI: http://dx.doi.org/10.5772/intechopen.84402


#### Table 5.

Figures 1 and 6 has been considered. This allows estimating a suitable diameter for the propeller Φprop (Figure 8) and setting the quantity of motors that fits

Cases Np

Case A Treq,T = 20.1 [N] @ 15 [m/s] Treq,i [N] 20.10 10.05 6.70

Case B Treq,T = 20.1 [N] @ 15 [m/s] Treq,i [N] 20.10 10.05 6.70

The thrust that each propeller must generate, in conjunction with their diameter, is presented in Table 4 for both cases (A and B) with three distinct configurations: single and distributed propulsion with double and triple propulsors. Note that when the propulsion system possesses more than one propulsor, the thrust required per propeller (Treq,i) is obtained by dividing the total thrust required (Treq) by the number of installed propellers. Besides, it is important to mention that the term Treq was obtained by dividing the power required for the desired value for cruise speed

1 23

Φprop [in] 18.00 8.50 5.50

Φprop [in] 13.00 6.00 3.75

Tables 5 and 6 summarize the performance characteristics of selected propellers for cases A and B, respectively. The suitable models were established regarding the freestream velocity, their diameter, and the rotational speed. Special attention was paid to guarantee that the considered propeller arrangements fit the available space and, at the same time, generate the thrust required. The cruise speed was set for

For each propulsion set, several parameters from propellers, such as: nondimensional coefficients (J, CT, and CP), efficiency, torque, shaft power, and thrust generated were extracted from the experimental database at desired operating conditions. Notice that the power available of individual propellers (PA,i) was obtained by multiplying the power shaft for its corresponding propeller efficiency (ηp). Likewise, the total power available (PA) is a function of the number or propellers (Np) implemented, and it was obtained by multiplying Np for its

corresponding individual power (PA,i). Finally, the power that battery must supply to the propulsion system was obtained by multiplying the power shaft for the

adequately the available space and generate the power needed.

(15 m/s for both cases).

Figure 8.

Propulsion Systems

Table 4.

Main components of propeller geometry.

Thrust required for distributed propulsion.

3.4.1 Propeller selection

62

minimum drag condition [35].

Propeller selection—case A.


#### Table 6.

Propeller selection—case B.

number of propellers implemented. Note that the shaft power and the battery power will be the same for a single propeller configuration because only one motor is considered and the electrical efficiency losses were neglected for practical purposes.

Note that, in Tables 5 and 6, the propeller efficiency is dramatically affected as its diameter is reduced. As observed, the lesser the propeller's size, the higher the rotational speed is needed to generate the required thrust. For instance, smaller models (e.g., three propellers 5 3) must operate at a high RPM (27,000) to produce the thrust required (Table 4). In contrast, a single larger propeller (e.g., a propeller 17 10) operates at lower RPM (5000) to generate the same thrust. These aspects also reflect the variation of propeller efficiency due to its direct linkage to the cruise speed, rotational speed (ω), and propeller diameter (ϕprop) [47, 48]. In this sense, the decrease of propeller efficiency (as its diameter is lower) is attributed to the increase of the induced velocity at the propeller tips, which in turn increments the tip losses due to drag [47].

3.5 Performance assessment

DOI: http://dx.doi.org/10.5772/intechopen.84402

Table 9.

> CA-1

> CA-2

Table 10.

65

Parameter for battery selection—cases A and B.

Wpropeller [kg]

WESC [kg]

[kg]

Weight assessment—cases A and B.

Case Np Wmotor

As mentioned, the performance of an electric aerial vehicle mainly depends on the battery capacity and the total current draw supplied to the propulsion system. For this work, a single battery device was considered for each configuration, even in the case of distributed propulsion. In this way, the battery current (IB) was computed as a function of several motor parameters such as voltage constant (Kv), no-load current (Im0), no-load voltage (Umo), and motor resistance (Tables 7 and 8). Another parameter employed for the battery selection was its nominal

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

For case A, when a single propeller is considered, the battery voltage is almost twice the voltage of the distributed propulsion scenario. This is because a larger propeller demands higher torque to work properly, and hence a larger motor that works with a higher voltage is employed. Nonetheless, it is also observed that the current supplied by the battery to the motors (IB) for the distributed propulsion configuration is twice the value of the single propeller set. This is because two motors that work in a parallel are employed in the first case. Moreover, it is seen that the motor of case B-1 requires a lower voltage than the case A-1, because of the difference in size of the propellers employed (Tables 5 and 6). The number of cells of the battery for each configuration was set according to the manufacturer suggestion. Table 9 compiles the results of general battery characteristics such as

Regarding the takeoff gross weight estimation, it can be appreciated that the higher the nominal voltage, the higher number of cells from the battery and hence the aircraft weight increases. In this sense, two different scenarios were investigated to select a battery. The first one consisted of selecting a desired capacity from commercial datasheets, and, from this, the vehicle performance was assessed. For this approach, Table 10 and Figure 9 present the UAV weight estimation and its breakdown for the three configurations once the battery characteristics have been outlined. Meanwhile, Table 11 shows the estimated range and endurance under these battery characteristics. It is interesting to note that for case A, the takeoff gross weight (WTO) of the aircraft is lower when the distributed propulsion system is implemented; however, the weight fractions of the components are altered. For instance, the propulsion

Case Np Cells UB [V] Ueo [V] Im [A] IB [A] CA-1 1 6 22.2 19.1 31.9 28.4 CA-2 2 3 11.1 10.7 32.6 63.8 CB-1 1 3 11.1 10.8 75.4 74.2

> Wbattery [kg]

1 0.275 0.063 0.05 2.024 0.538 1 3.95

2 0.190 0.014 0.05 1.055 0.538 1 3.10

CB-1 1 0.177 0.041 0.06 1.055 0.118 1 2.45

Wfuselage [kg]

Wpayload [kg]

WTO [kg]

voltage (UB), which is directly linked to the number of internal cells.

the number of cells and total current that battery supplies.

In addition, the lower propeller efficiency, the higher power that needs to be delivered by the motor (Pshaft). As observed in Tables 5 and 6, the large drop in efficiency (around 20% for the increment of one propeller) produces an increment in total power consumed. Nonetheless, the total power available (PA) for all the cases remains almost constant, and it agrees the target power available (302 W).

Table 6 shows similar results as Table 5; however, it is worth to highlight that propeller efficiencies are even lower due to the smaller propeller's diameters for this case, which results in higher values of power shaft compared to case A. Additionally, note that the three-propeller configuration for this case was not assessed due to the unsuitability of allocating more than two propellers within the airframe trailing edge.

#### 3.4.2 Motor selection

The selection of an electric motor for each set of propellers was accomplished through a catalogue-search of different manufacturers. The key parameters for motor selection and propeller-motor matching were the revolutions per minute (RPM), power shaft, and voltage constant. The latter represents a motor constant which correlates the RPMs and the operating voltage. Finally, when the set of the propeller and electric components is found, recommended values given by manufacturers for propeller pairing were used to check if the established arrangement meets the requirements.

Tables 7 and 8 present the motor devices for the propulsion configurations of cases A and B, respectively. As observed, for the distributed propulsion systems with three propellers of case A (Table 7) and with two propellers of case B (Table 8), it was not possible to find any off-the-shelf electrical motor that fulfills the propeller requirements, and, hence, they were not used in further analysis. For all favorable cases (first and second arrangements of case A and first set of case B), adequate motor models were decided, and their operating parameters were employed to size and select a suitable battery to perform a given mission.


Table 7. Motor selection—case A.


Table 8. Motor selection—case B.

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion… DOI: http://dx.doi.org/10.5772/intechopen.84402

#### 3.5 Performance assessment

aspects also reflect the variation of propeller efficiency due to its direct linkage to the cruise speed, rotational speed (ω), and propeller diameter (ϕprop) [47, 48]. In this sense, the decrease of propeller efficiency (as its diameter is lower) is attributed to the increase of the induced velocity at the propeller tips, which in turn incre-

In addition, the lower propeller efficiency, the higher power that needs to be delivered by the motor (Pshaft). As observed in Tables 5 and 6, the large drop in efficiency (around 20% for the increment of one propeller) produces an increment in total power consumed. Nonetheless, the total power available (PA) for all the cases remains almost constant, and it agrees the target power available (302 W). Table 6 shows similar results as Table 5; however, it is worth to highlight that propeller efficiencies are even lower due to the smaller propeller's diameters for this case, which results in higher values of power shaft compared to case A. Additionally, note that the three-propeller configuration for this case was not assessed due to the unsuitability of allocating more than two propellers within the airframe trailing edge.

The selection of an electric motor for each set of propellers was accomplished through a catalogue-search of different manufacturers. The key parameters for motor selection and propeller-motor matching were the revolutions per minute (RPM), power shaft, and voltage constant. The latter represents a motor constant which correlates the RPMs and the operating voltage. Finally, when the set of the propeller and electric components is found, recommended values given by manufacturers for propeller pairing were used to check if the established arrangement

Tables 7 and 8 present the motor devices for the propulsion configurations of cases A and B, respectively. As observed, for the distributed propulsion systems with three propellers of case A (Table 7) and with two propellers of case B (Table 8), it was not possible to find any off-the-shelf electrical motor that fulfills the propeller requirements, and, hence, they were not used in further analysis. For all favorable cases (first and second arrangements of case A and first set of case B),

Np Propeller Motor RPM Battery cells KV Im0 [A] Um0 [V] Rm [Ohm] 1 17 10 A40-14L V4 14-Pole 6000 6S 355 0.85 8.4 0.050 2 8 6 A40-12S V4 8-Pole 12,820 3S 1350 1.94 8.4 0.018 3 5 3 No motor matching – – –– – –

Np Propeller Motor RPM Battery cells KV Im0 [A] Um0 [V] Rm [Ohm] 1 12 6.5 A30 8 XL V4 9500 3S 1100 2.8 8.4 0.015 2 6 3 No motor matching – – –– – –

adequate motor models were decided, and their operating parameters were employed to size and select a suitable battery to perform a given mission.

ments the tip losses due to drag [47].

3.4.2 Motor selection

Propulsion Systems

meets the requirements.

Table 7.

Table 8.

64

Motor selection—case A.

Motor selection—case B.

As mentioned, the performance of an electric aerial vehicle mainly depends on the battery capacity and the total current draw supplied to the propulsion system. For this work, a single battery device was considered for each configuration, even in the case of distributed propulsion. In this way, the battery current (IB) was computed as a function of several motor parameters such as voltage constant (Kv), no-load current (Im0), no-load voltage (Umo), and motor resistance (Tables 7 and 8). Another parameter employed for the battery selection was its nominal voltage (UB), which is directly linked to the number of internal cells.

For case A, when a single propeller is considered, the battery voltage is almost twice the voltage of the distributed propulsion scenario. This is because a larger propeller demands higher torque to work properly, and hence a larger motor that works with a higher voltage is employed. Nonetheless, it is also observed that the current supplied by the battery to the motors (IB) for the distributed propulsion configuration is twice the value of the single propeller set. This is because two motors that work in a parallel are employed in the first case. Moreover, it is seen that the motor of case B-1 requires a lower voltage than the case A-1, because of the difference in size of the propellers employed (Tables 5 and 6). The number of cells of the battery for each configuration was set according to the manufacturer suggestion. Table 9 compiles the results of general battery characteristics such as the number of cells and total current that battery supplies.

Regarding the takeoff gross weight estimation, it can be appreciated that the higher the nominal voltage, the higher number of cells from the battery and hence the aircraft weight increases. In this sense, two different scenarios were investigated to select a battery. The first one consisted of selecting a desired capacity from commercial datasheets, and, from this, the vehicle performance was assessed. For this approach, Table 10 and Figure 9 present the UAV weight estimation and its breakdown for the three configurations once the battery characteristics have been outlined. Meanwhile, Table 11 shows the estimated range and endurance under these battery characteristics.

It is interesting to note that for case A, the takeoff gross weight (WTO) of the aircraft is lower when the distributed propulsion system is implemented; however, the weight fractions of the components are altered. For instance, the propulsion


Table 9.

Parameter for battery selection—cases A and B.


Table 10. Weight assessment—cases A and B.

same mission flight is different for all three cases because of the notable increment

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

Wetland monitoring at the Andean highlands presents imperious needs in terms

of more efficient and environmentally friendly UAV designs. The challenges imposed by the hard-operating requirements make commercial electrical small UAVs not suitable for monitoring tasks. However, their lower environmental impact and low costs encourage their improvements through the implementation of different technologies. In this context, this work has assessed the performance of two different configurations using electrical distributed propulsion system. Since for this particular application the payload required is small, the search for architectures with a large number of propulsors was not suitable, and hence the maximum number of propellers that could be allocated was set to three. The results showed

that the propeller's size reduction affected dramatically its performance,

outweighing the benefits in weight and propulsive efficiency. The results from the performance analysis show that endurance, range, and weight decrease for distributed propulsion configurations. The reduction on total aircraft weight is beneficial; however, this was outweighed by the lower thermal performance of the propellers, which reflected on higher total power consumption for the distributed propulsion cases. It is important to highlight that distributed propulsion may be a good option when the propeller's size does not decrease so dramatically as in the configurations studied. This aspect is a key variable in order to assess a recommendable range for distributed propulsor size, where the benefits obtained from weight, reliability, control, and aerodynamic aspects are not to be affected by the thermal propulsors'

Another aspect to highlight in this study is that conversely to other aviation sectors, the small UAV category allows to implement or design an important variety of electronic propulsion components, which can be easily tailored for the operating requirements. This aspect opens up the door to generate more flexible designs which incorporates forefront technology such as thrust vectoring, boundary layer ingestion (BLI), and propulsion airframe embedded (PAE) designs, among others in exchange of moderate costs. In this sense, it is important to implement a versatile and flexible optimization methodology, which enables the assessment of different geometrical configurations accounting for the aerodynamic and propulsion integration at system engineering perspective. These routines will enable to evaluate a larger range of aircraft configurations, and it will contribute to establish ideal

Finally, since this study was implemented at conceptual level, the distortion features and momentum reduction produced by the ingestion of the boundary layer have not been considered. Nonetheless, further studies to determine the compromise between size and propeller performance incorporating BLI aspects need to be

The authors would like to thank to Corporación Ecuatoriana para el Desarrollo de la Investigación y Academia (CEDIA) for the financial support given to the present research, development, and innovation work through its CEPRA program, especially for the CEPRA XII-2018-12 fund. Furthermore, the authors gratefully

models that fulfill the requirements from a synergetic standpoint.

of the current draw (Table 9).

DOI: http://dx.doi.org/10.5772/intechopen.84402

4. Summary

performance.

carried out.

67

Acknowledgements

Figure 9.

UAV—weight breakdown. (a) Case A-1 propeller, (b) case A-2 propeller, and (c) case B-1 propeller.


Table 11. Battery selection—cases A and B.

system mass fraction increments from 9.8 to 16.4%, and, in similar way, the fuselage mass fraction increases from 25.3 to 32.3%. Although the lessening of weight that resulted from the implementation of the distributed propulsion system, this is overshadowed by the lessening of the endurance and range of the aircraft because the power consumed by the motors in the distributed propulsion scenario is higher, as showed in Table 11.

At last, comparing the single propeller configurations (Figure 9a and c) indicates that fuselage mass fraction is decreased for case B because of its smaller geometry. Note that the battery mass fraction is lower for case B because a 3-cell battery was employed for this configuration; meanwhile, for case A, a 6-cell configuration was utilized, which results in a weight increment.

The second performance approach determined the minimum battery capacity required for a target endurance and range. Table 11 (Method 2) shows the results for case A (one and two-propeller configuration) and case B (one-propeller configuration). As can be seen, for the same target endurance, the battery capacity for the distributed propulsion configuration is dramatically raised compared with the single propeller setting. In other words, the capacity lasts less due to the augment of current draw that the battery must provide as consequence of the loss in propeller efficiency. When comparing the one-propeller configuration of both cases (A and B), it is observable that the endurance reduces for case B because of the aerodynamic and propulsion drawbacks arising from a lesser wetted area and lower efficiency of the propeller.

The results from the second method of performance assessment showed that the required battery capacity to accomplish a specific mission increases for distributed propulsion arrangements. As observed in Table 11, the capacity to perform the

same mission flight is different for all three cases because of the notable increment of the current draw (Table 9).

#### 4. Summary

Wetland monitoring at the Andean highlands presents imperious needs in terms of more efficient and environmentally friendly UAV designs. The challenges imposed by the hard-operating requirements make commercial electrical small UAVs not suitable for monitoring tasks. However, their lower environmental impact and low costs encourage their improvements through the implementation of different technologies. In this context, this work has assessed the performance of two different configurations using electrical distributed propulsion system. Since for this particular application the payload required is small, the search for architectures with a large number of propulsors was not suitable, and hence the maximum number of propellers that could be allocated was set to three. The results showed that the propeller's size reduction affected dramatically its performance, outweighing the benefits in weight and propulsive efficiency. The results from the performance analysis show that endurance, range, and weight decrease for distributed propulsion configurations. The reduction on total aircraft weight is beneficial; however, this was outweighed by the lower thermal performance of the propellers, which reflected on higher total power consumption for the distributed propulsion cases. It is important to highlight that distributed propulsion may be a good option when the propeller's size does not decrease so dramatically as in the configurations studied. This aspect is a key variable in order to assess a recommendable range for distributed propulsor size, where the benefits obtained from weight, reliability, control, and aerodynamic aspects are not to be affected by the thermal propulsors' performance.

Another aspect to highlight in this study is that conversely to other aviation sectors, the small UAV category allows to implement or design an important variety of electronic propulsion components, which can be easily tailored for the operating requirements. This aspect opens up the door to generate more flexible designs which incorporates forefront technology such as thrust vectoring, boundary layer ingestion (BLI), and propulsion airframe embedded (PAE) designs, among others in exchange of moderate costs. In this sense, it is important to implement a versatile and flexible optimization methodology, which enables the assessment of different geometrical configurations accounting for the aerodynamic and propulsion integration at system engineering perspective. These routines will enable to evaluate a larger range of aircraft configurations, and it will contribute to establish ideal models that fulfill the requirements from a synergetic standpoint.

Finally, since this study was implemented at conceptual level, the distortion features and momentum reduction produced by the ingestion of the boundary layer have not been considered. Nonetheless, further studies to determine the compromise between size and propeller performance incorporating BLI aspects need to be carried out.

#### Acknowledgements

The authors would like to thank to Corporación Ecuatoriana para el Desarrollo de la Investigación y Academia (CEDIA) for the financial support given to the present research, development, and innovation work through its CEPRA program, especially for the CEPRA XII-2018-12 fund. Furthermore, the authors gratefully

system mass fraction increments from 9.8 to 16.4%, and, in similar way, the fuselage mass fraction increases from 25.3 to 32.3%. Although the lessening of weight that resulted from the implementation of the distributed propulsion system, this is overshadowed by the lessening of the endurance and range of the aircraft because the power consumed by the motors in the distributed propulsion scenario is higher,

CA-1 16,000 27.1 24.39 30 27 17731 CA-2 16,000 12.0 10.8 30 27 39,892 CB-1 16,000 10.3 9.27 30 27 46,378

UAV—weight breakdown. (a) Case A-1 propeller, (b) case A-2 propeller, and (c) case B-1 propeller.

Input Output Input Output CB [mAh] E [min] R [km] E [min] R [km] CB [mAh]

Cases Method 1 Method 2

At last, comparing the single propeller configurations (Figure 9a and c) indicates that fuselage mass fraction is decreased for case B because of its smaller geometry. Note that the battery mass fraction is lower for case B because a 3-cell battery was employed for this configuration; meanwhile, for case A, a 6-cell con-

The second performance approach determined the minimum battery capacity required for a target endurance and range. Table 11 (Method 2) shows the results for case A (one and two-propeller configuration) and case B (one-propeller configuration). As can be seen, for the same target endurance, the battery capacity for the distributed propulsion configuration is dramatically raised compared with the single propeller setting. In other words, the capacity lasts less due to the augment of current draw that the battery must provide as consequence of the loss in propeller efficiency. When comparing the one-propeller configuration of both cases (A and B), it is observable that the endurance reduces for case B because of the aerodynamic and propulsion drawbacks arising from a lesser wetted area and lower efficiency of the

The results from the second method of performance assessment showed that the required battery capacity to accomplish a specific mission increases for distributed propulsion arrangements. As observed in Table 11, the capacity to perform the

figuration was utilized, which results in a weight increment.

as showed in Table 11.

Battery selection—cases A and B.

Figure 9.

Propulsion Systems

Table 11.

propeller.

66

acknowledge the financial support provided by Escuela Politécnica Nacional for the development of the internal projects: PIJ 15-11 and PIS 16-20.

### Abbreviations and nomenclature


Author details

Ecuador

69

Esteban Valencia\*, Víctor Alulema and Darío Rodríguez

provided the original work is properly cited.

\*Address all correspondence to: esteban.valencia@epn.edu.ec

Department of Mechanical Engineering, National Polytechnic School, Quito,

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion…

DOI: http://dx.doi.org/10.5772/intechopen.84402

© 2019 The Author(s). Licensee IntechOpen. This chapter is distributed under the terms of the Creative Commons Attribution License (http://creativecommons.org/licenses/ by/3.0), which permits unrestricted use, distribution, and reproduction in any medium,

Wetland Monitoring Using Unmanned Aerial Vehicles with Electrical Distributed Propulsion… DOI: http://dx.doi.org/10.5772/intechopen.84402

### Author details

acknowledge the financial support provided by Escuela Politécnica Nacional for the

]

development of the internal projects: PIJ 15-11 and PIS 16-20.

Abbreviations and nomenclature

Propulsion Systems

ϕprop propeller diameter, [in]

CB battery capacity [mAh]

KV voltage constant [RPM/V] Np number of propellers RC rate of climb [m/s] S planform wing area [m2

PA power available [W] PR power required [W]

Pbat power of battery [W]

UAV unmanned aerial vehicle UB nominal battery voltage [V]

Vc cruise speed [m/s] Vs stall speed [m/s] W/S wing loading [N/m<sup>2</sup>

Wbattery battery weight [N]

Wfuselage fuselage weight [N] Wmotor motor weight [N] Wpayload payload weight [N] Wpropeller propeller weight [N] WTO takeoff gross weight [N]

68

Q torque [Nm] T thrust [N]

E endurance [min] IB total current [A] Im motor current [A]

Cmin minimum battery capacity [mAh]

Im0 nominal no-load motor current [A]

PRp preliminary power required [W] Pshaft power shaft of the motor [W]

Um0 nominal no-load voltage of motor [V]

]

Ueo equivalent battery voltage [V] Vmax maximum flight speed [m/s]

W/P weight to power ratio [N/W]

WESC electronic speed control weight [N]

WTOp preliminary takeoff gross weight [N]

]

ηprop propeller efficiency BWB blended wing body CD drag coefficient CL lift coefficient

ρalt air density at desired altitude [kg/m<sup>3</sup>

ω motor/propeller rotational speed [RPM]

η<sup>e</sup> electric set (motor, battery, ESC) efficiency

Esteban Valencia\*, Víctor Alulema and Darío Rodríguez Department of Mechanical Engineering, National Polytechnic School, Quito, Ecuador

\*Address all correspondence to: esteban.valencia@epn.edu.ec

© 2019 The Author(s). Licensee IntechOpen. This chapter is distributed under the terms of the Creative Commons Attribution License (http://creativecommons.org/licenses/ by/3.0), which permits unrestricted use, distribution, and reproduction in any medium, provided the original work is properly cited.

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[12] Panagiotou P, Fotiadis-Karras S, Yakinthos K. Conceptual design of a blended wing body MALE UAV. Aerospace Science and Technology. 2018;73:32-47

[13] Lehmkuehler K, Wong KC, Verstraete D. Design and test of a UAV blended wing body configuration. 28th Congr Int. Counc. Aeronaut. Sci. 2012, ICAS 2012, vol. 1. April. 2012. pp. 432-442

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[15] Dehpanah P, Nejat A. The aerodynamic design evaluation of a blended-wing-body configuration. Aerospace Science and Technology. 2015;43:96-110

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with distributed propulsion. Aerospace Science and Technology. 2013;25(1): 16-28

References

Propulsion Systems

[1] Baruch Z. Ordination and classification of vegetation along an altitudinal gradient in the Venezuelan páramos. Vegetatio. 1984;55(2):115-126

[2] Schoolmeester T et al. Outlook on climate change adaptation in the Tropical Andes mountains. 2016

[9] E. 38 Unmanned Systems. The E384 Mapping Drone – Event 38 Unmanned Systems. 2018. [Online]. Available from: https://event38.com/fixed-wing/e 384-mapping-drone/ [Accessed: 18-

[10] Parrot. Parrot DISCO FPV | Official Parrot® Site. [Online]. Available: https://www.parrot.com/us/drones/ parrot-disco-fpv#-parrot-disco-fpv

[11] Valencia E, Saá JM, Alulema V, Hidalgo V. Parametric study of

coupled Blended Wing Body

2018. pp. 1-15

2018;73:32-47

432-442

aerodynamic integration issues in highly

configurations implemented in UAVs.

[12] Panagiotou P, Fotiadis-Karras S, Yakinthos K. Conceptual design of a blended wing body MALE UAV. Aerospace Science and Technology.

[13] Lehmkuehler K, Wong KC,

[14] Panagiotou P, Yakinthos K. Parametric aerodynamic study of Blended-Wing-Body platforms at low subsonic speeds for UAV applications. In: 35th AIAA Appl. Aerodyn. Conf., no.

[15] Dehpanah P, Nejat A. The aerodynamic design evaluation of a blended-wing-body configuration. Aerospace Science and Technology.

[16] Leifsson L, Ko A, Mason WH, Schetz JA, Grossman B, Haftka RT. Multidisciplinary design optimization of blended-wing-body transport aircraft

June. 2017. pp. 1-19

2015;43:96-110

Verstraete D. Design and test of a UAV blended wing body configuration. 28th Congr Int. Counc. Aeronaut. Sci. 2012, ICAS 2012, vol. 1. April. 2012. pp.

[Accessed: 12-Sep-2018]

June-2018]

[3] da GABF, Norman Myers JK, Mittermeier RA, Mittermeier CG. Biodiversity Hotspots for conservation

[4] Josse CE, Cuesta F, Navarro G, Barrena V, Chacón-Moreno ESJ, Ferreira W, et al. Ecosistemas de los Andes del Norte Centro. Bolivia, Colombia, Ecuador, Perú y Venezuela.

priorities. Nature. 2000;403

Lima: Secretaría General de la

ECOBONA-Intercooperation,

SRL; 2009

(1W4):249-256

70

Comunidad Andina, Programa Regional

CONDESAN-Proyecto Páramo Andino, Programa BioAndes, EcoCiencia, NatureServe, IAvH, LTA-UNALM, ICAE-ULA, CDC-UNALM, RUMBOL

[5] BuytaertW et al. Human impact on the hydrology of the Andean páramos. Earth-Science Reviews. 2006;79(1–2):53-72

[6] Gundlach J. Civil and Commercial

Washington, DC: American Institute of Aeronautics and Astronautics, Inc; 2016

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[8] Shahbazi M, Théau J, Ménard P. Recent applications of unmanned aerial

imagery in natural resource management. GIScience Remote Sensing. 2014;51(4):339-365

Unmanned Aircraft Systems.

[17] Hepperle M. Electric Flight– Potential and Limitations. 2012. pp. 1-30

[18] Gundlach J. Designing Unmanned Aircraft Systems: A Comprehensive Approach. Manassas, Virginia: American Institute of Aeronautics and Astronautics, Inc; 2011

[19] Gohardani AS. A synergistic glance at the prospects of distributed propulsion technology and the electric aircraft concept for future unmanned air vehicles and commercial/military aviation. Progress in Aerospace Science. 2013;57:25-70

[20] Kim HD, Perry AT, Ansell PJ. A review of distributed electric propulsion concepts for air vehicle technology. In: 2018 AIAA/IEEE Electr. Aircr. Technol. Symp. 2018

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vineyards with a small UAV. Biosystems Engineering. 2013;108(4):49-61

[25] Macjke DC Jr. Systems and image database resources for UAV search and rescue applications [masters theses]. 2013. p. 115

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Chapter 4

Koichi Mori

reaches at altitude higher than 10 km.

power transmission

1. Introduction

73

Abstract

Beamed Launch Propulsion

An advanced concept of launch system from ground to orbit, called laser launch

system, has been discussed. As a 100-kW-class fiber laser has been developed today, the laser propulsion is now a realistic option for launching microsatellites frequently at very low cost. In this chapter, we shall discuss several unresolved technical problems such as propulsion design and laser beam transmission through atmosphere. It is proved theoretically that high specific impulse higher than 900 seconds is possible in a new conceptual design. On the other hand, the laser beam may be suffered by the atmospheric turbulence when the launch vehicle

Keywords: launch vehicle, laser propulsion, laser, rocket propulsion, wireless

In this chapter, we consider the unknown system called laser propulsion. Designing the system is a mixture of various engineering fields, including propulsion engineering, laser engineering, electromagnetic wave engineering, flight dynamics, and control engineering, which are interconnected to each other. The laser propulsion has been studied in the field of the propulsion engineering for more than 50 years. Moreover, the laser propulsion appeared as a gadget in several sci-fi works [1, 2]. (Strangely, in all those works, the laser propulsion is introduced as a technology of aliens rather than an earth-oriented technology. It would reflect that this technology is full of mysterious images.) However, in view of the practical application, few have achieved so far. This is mainly because no laser facilities whose continuous power is sufficiently high have been available. However, in 2010, the world has changed. A 100 kW fiber laser has been commercialized by IPG Photonics Inc. A 100 kW fiber laser facility has been delivered to the NADEX laser R&D Center in Fukui, Japan. We come to the place where we can do genuine experiments of laser propulsion. Surrounding situations are being prepared today.

As shown in Figure 1, in the laser launch system, a vehicle is propelled by transmitting the propulsive energy via laser beam from the ground. Gaining the specific impulse higher than the practical chemical propulsions, the laser propulsion is of the same class of the electric propulsion for spacecraft. At the same time, by leaving the heavy part of the energy source on the ground, the lightweight vehicle

In the following, at first, we are going to consider why the laser propulsion is necessary. For this, we need to consider what kind of launch systems will be required in near future. The conventional rocket technology is in the period of

Laser propulsion is one step short of practical application.

on the basis of simple propulsion energy system is realized.

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[44] Sóbester A, Forrester AIJ. Aircraft Aerodynamic Design: Geometry and Optimization. 2014

[45] Shim HJ, Park SO. Low-speed windtunnel test results of a BWB-UCAV model. Procedia Engineering. 2013;67: 50-58

[46] Gur O, Mason WH, Schetz JA. Fullconfiguration drag estimation. Journal of Aircraft. 2010;47(4):1356-1367

[47] Filippone A. Advanced Aircraft Flight Performance. 2010

[48] Torenbeek E. Synthesis of Subsonic Airplane Design. Dordrecht: Springer Netherlands, 1982

## Chapter 4 Beamed Launch Propulsion

Koichi Mori

### Abstract

[37] Drela M. DC Motor/Propeller Matching - Lab 5 Lecture Notes. 2005;6:6

Propulsion Systems

[38] Avanzini G, De Angelis EL, Giulietti F. Optimal performance and sizing of a battery-powered aircraft. Aerospace Science and Technology. 2016;59:132-144

[39] Elekes G. A geometric inequality and the complexity of computing volume. Discrete & Computational Geometry. Dec. 1986;1(4):289-292

[40] Valencia EA, Hidalgo V, Rodriguez

aerodynamic assessment of a fixed wing UAV implemented for Site Specific Management. In: 2018 AIAA Inf. Syst. Infotech @ Aerosp., no. January. 2018.

[41] Traub LW. Range and endurance estimates for battery-powered aircraft. Journal of Aircraft. 2011;48(2):703-707

[42] Scarpino M. Motors for Makers A Guide to Steppers, Servos, and Other

[43] Kontogiannis S, Ekaterianaris J. Design, performance evaluation and optimization of a UAV. Aerospace Science and Technology. 2013;January:

[44] Sóbester A, Forrester AIJ. Aircraft Aerodynamic Design: Geometry and

[45] Shim HJ, Park SO. Low-speed windtunnel test results of a BWB-UCAV model. Procedia Engineering. 2013;67:

[46] Gur O, Mason WH, Schetz JA. Fullconfiguration drag estimation. Journal of Aircraft. 2010;47(4):1356-1367

[47] Filippone A. Advanced Aircraft

[48] Torenbeek E. Synthesis of Subsonic Airplane Design. Dordrecht: Springer

Flight Performance. 2010

Netherlands, 1982

Electrical Machines; 2016

D. Parametric modelling for

pp. 1–17

339-350

50-58

72

Optimization. 2014

An advanced concept of launch system from ground to orbit, called laser launch system, has been discussed. As a 100-kW-class fiber laser has been developed today, the laser propulsion is now a realistic option for launching microsatellites frequently at very low cost. In this chapter, we shall discuss several unresolved technical problems such as propulsion design and laser beam transmission through atmosphere. It is proved theoretically that high specific impulse higher than 900 seconds is possible in a new conceptual design. On the other hand, the laser beam may be suffered by the atmospheric turbulence when the launch vehicle reaches at altitude higher than 10 km.

Keywords: launch vehicle, laser propulsion, laser, rocket propulsion, wireless power transmission

#### 1. Introduction

In this chapter, we consider the unknown system called laser propulsion. Designing the system is a mixture of various engineering fields, including propulsion engineering, laser engineering, electromagnetic wave engineering, flight dynamics, and control engineering, which are interconnected to each other. The laser propulsion has been studied in the field of the propulsion engineering for more than 50 years. Moreover, the laser propulsion appeared as a gadget in several sci-fi works [1, 2]. (Strangely, in all those works, the laser propulsion is introduced as a technology of aliens rather than an earth-oriented technology. It would reflect that this technology is full of mysterious images.) However, in view of the practical application, few have achieved so far. This is mainly because no laser facilities whose continuous power is sufficiently high have been available. However, in 2010, the world has changed. A 100 kW fiber laser has been commercialized by IPG Photonics Inc. A 100 kW fiber laser facility has been delivered to the NADEX laser R&D Center in Fukui, Japan. We come to the place where we can do genuine experiments of laser propulsion. Surrounding situations are being prepared today. Laser propulsion is one step short of practical application.

As shown in Figure 1, in the laser launch system, a vehicle is propelled by transmitting the propulsive energy via laser beam from the ground. Gaining the specific impulse higher than the practical chemical propulsions, the laser propulsion is of the same class of the electric propulsion for spacecraft. At the same time, by leaving the heavy part of the energy source on the ground, the lightweight vehicle on the basis of simple propulsion energy system is realized.

In the following, at first, we are going to consider why the laser propulsion is necessary. For this, we need to consider what kind of launch systems will be required in near future. The conventional rocket technology is in the period of

laser beam may be a critical factor in the feasibility study of the LTP. A laser beam is transmitted across turbulent atmosphere for a long distance up to 100 km from

There are two kinds of the laser propulsion: repetitive pulse (RP) laser type and continuous wave (CW) laser type. The concept of laser propulsion is first proposed by Kantrowitz in 1971 [7]. His concept was to irradiate the laser beam on the ablator installed on the bottom surface of the vehicle as propellant. At that time, it was unknown how much the momentum can be generated for a certain laser power. The research team led by Kantrowitz first investigated the momentum coupling performances of RP laser propulsion and its physical mechanisms. As a result, the impulse generation mechanisms of laser-supported detonation waves and laser-supported combustion waves have been developed. At the beginning of 2000s, Myrabo invented a new vehicle design called a Lightcraft, which is illustrated in Figure 2, to perform first launch demonstrations using a 10-kW-class CO2 RP laser facility of US Air force [8]. He succeeded in the independent flight of the vehicle without any external guide or support except for the laser beam for the first time in the world. The world record of the flight altitude was 71 m. Through the development of the Lightcraft, Myrabo developed the concept of "beam-riding." The vehicle must be kept irradiated to generate the momentum all the time of flight. Lightcraft was designed to keep its trajectory along a fixed laser beam, while this is the meaning of the term of beam-riding. When the vehicle position is deviated from the laser beam, the recovery side force is generated to the vehicle to keep the trajectory. His consideration was epoch-making because no previous studies in the laser propulsion have considered the flight dynamics of the vehicle. In the same periods, Sasoh invented the in-tube laser propulsion, which is illustrated in Figure 3, and investigated the concept experimentally using a 1-kW-class RP CO2 laser. A projectile could be accelerated in a tube efficiently due to the confinement effect of the tube wall. In the times of early 2000s, several different types of RP laser propulsion have been invented and investigated experimentally. The concepts of laser propulsion that have been proposed so far are reviewed in the two review papers in detail [9, 10]. From the research team of the author, new laser launch system using "donut-beam," whose power density has hollow distribution on the cross-sectional plane, higher at peripheral of the cross section than at the center, and a spherical vehicle for stable acceleration has been proposed. This concept is studied in experiments [11] and numerical simulations [12]. Because the concept uses the atmospheric air as propellant, the acceleration performance of vehicle is determined by the aerodynamic drag and the atmospheric air density to be propelled using the laser power. Each concept of

ground to space.

Beamed Launch Propulsion

DOI: http://dx.doi.org/10.5772/intechopen.82236

Figure 2. Myrabo's Lightcraft.

75

#### Figure 1. Laser launch system.

maturity, while the fetal movement for new launch system of high specific impulse that is not limited by the chemical reaction has been ignited. Then, we need to investigate what kind of the laser propulsion system is technologically feasible. Ensuring the foregoing studies of the laser propulsion, we shall consider the practical laser launch system. The technological problems to be solved are numerous. The laser transmission through turbulent atmosphere gives critical problems to the feasibility in the laser launch system although they have been half ignored so far. Solving the numerous technological problems that are necessary to realize the laser launch system is related to the advanced technology such as Starshot project [3], space-based solar power [4], and laser communication in space [5]. Naturally, the space will become familiar by realizing the laser propulsion. Laser propulsion is the technology that is worth challenging.

### 2. Principles of laser propulsions

Laser propulsion is a variation of the wireless power transfer (WPT) technology, which transfers power remotely using electromagnetic (EM) waves such as microwaves or laser beams. When electric power is necessary at the receiving side, the power of EM waves is transformed to dc current using semiconductors. For the laser propulsion, the power of EM waves is transformed directly into the enthalpy of a working fluid to generate momentum via thermal propulsion mechanisms. This can be called the laser thermal propulsion (LTP). Similar idea would be the laser electric propulsion (LEP) that converts the EM wave power to the dc power to store in the battery once, and then the dc power is used to generate thrust via electric propulsion mechanisms. The LEP is a new idea that allows the storage of the energy on board. However, the heavy weight of the battery would be the bottleneck for the feasibility of this system. Moreover, the LTP will be more energy efficient than LEP because the LEP requires multiple energy conversion processes, which lose the power at each stage. On the other hand, for LTP, it is necessary that the laser beam is always irradiated on the vehicle so that the control mechanisms keep the linkage between a laser source on ground and a fast-moving vehicle. Moreover, it is necessary to keep the vehicle in sight of the laser source so that available flight trajectory of the vehicle is largely limited. Furthermore, the atmospheric perturbation to the

#### Beamed Launch Propulsion DOI: http://dx.doi.org/10.5772/intechopen.82236

laser beam may be a critical factor in the feasibility study of the LTP. A laser beam is transmitted across turbulent atmosphere for a long distance up to 100 km from ground to space.

There are two kinds of the laser propulsion: repetitive pulse (RP) laser type and continuous wave (CW) laser type. The concept of laser propulsion is first proposed by Kantrowitz in 1971 [7]. His concept was to irradiate the laser beam on the ablator installed on the bottom surface of the vehicle as propellant. At that time, it was unknown how much the momentum can be generated for a certain laser power. The research team led by Kantrowitz first investigated the momentum coupling performances of RP laser propulsion and its physical mechanisms. As a result, the impulse generation mechanisms of laser-supported detonation waves and laser-supported combustion waves have been developed. At the beginning of 2000s, Myrabo invented a new vehicle design called a Lightcraft, which is illustrated in Figure 2, to perform first launch demonstrations using a 10-kW-class CO2 RP laser facility of US Air force [8]. He succeeded in the independent flight of the vehicle without any external guide or support except for the laser beam for the first time in the world. The world record of the flight altitude was 71 m. Through the development of the Lightcraft, Myrabo developed the concept of "beam-riding." The vehicle must be kept irradiated to generate the momentum all the time of flight. Lightcraft was designed to keep its trajectory along a fixed laser beam, while this is the meaning of the term of beam-riding. When the vehicle position is deviated from the laser beam, the recovery side force is generated to the vehicle to keep the trajectory. His consideration was epoch-making because no previous studies in the laser propulsion have considered the flight dynamics of the vehicle. In the same periods, Sasoh invented the in-tube laser propulsion, which is illustrated in Figure 3, and investigated the concept experimentally using a 1-kW-class RP CO2 laser. A projectile could be accelerated in a tube efficiently due to the confinement effect of the tube wall. In the times of early 2000s, several different types of RP laser propulsion have been invented and investigated experimentally. The concepts of laser propulsion that have been proposed so far are reviewed in the two review papers in detail [9, 10]. From the research team of the author, new laser launch system using "donut-beam," whose power density has hollow distribution on the cross-sectional plane, higher at peripheral of the cross section than at the center, and a spherical vehicle for stable acceleration has been proposed. This concept is studied in experiments [11] and numerical simulations [12]. Because the concept uses the atmospheric air as propellant, the acceleration performance of vehicle is determined by the aerodynamic drag and the atmospheric air density to be propelled using the laser power. Each concept of

Figure 2. Myrabo's Lightcraft.

maturity, while the fetal movement for new launch system of high specific impulse that is not limited by the chemical reaction has been ignited. Then, we need to investigate what kind of the laser propulsion system is technologically feasible. Ensuring the foregoing studies of the laser propulsion, we shall consider the practical laser launch system. The technological problems to be solved are numerous. The laser transmission through turbulent atmosphere gives critical problems to the feasibility in the laser launch system although they have been half ignored so far. Solving the numerous technological problems that are necessary to realize the laser launch system is related to the advanced technology such as Starshot project [3], space-based solar power [4], and laser communication in space [5]. Naturally, the space will become familiar by realizing the laser propulsion. Laser propulsion is the

Laser propulsion is a variation of the wireless power transfer (WPT) technology, which transfers power remotely using electromagnetic (EM) waves such as microwaves or laser beams. When electric power is necessary at the receiving side, the power of EM waves is transformed to dc current using semiconductors. For the laser propulsion, the power of EM waves is transformed directly into the enthalpy of a working fluid to generate momentum via thermal propulsion mechanisms. This can be called the laser thermal propulsion (LTP). Similar idea would be the laser electric propulsion (LEP) that converts the EM wave power to the dc power to store in the battery once, and then the dc power is used to generate thrust via electric propulsion mechanisms. The LEP is a new idea that allows the storage of the energy on board. However, the heavy weight of the battery would be the bottleneck for the feasibility of this system. Moreover, the LTP will be more energy efficient than LEP because the LEP requires multiple energy conversion processes, which lose the power at each stage. On the other hand, for LTP, it is necessary that the laser beam is always irradiated on the vehicle so that the control mechanisms keep the linkage between a laser source on ground and a fast-moving vehicle. Moreover, it is necessary to keep the vehicle in sight of the laser source so that available flight trajectory of the vehicle is largely limited. Furthermore, the atmospheric perturbation to the

technology that is worth challenging.

Figure 1. Laser launch system.

Propulsion Systems

74

2. Principles of laser propulsions

exchanger (HX) rocket proposed by Kare, which is illustrated in Figure 6 [14]. The laser power is converted to the propellant enthalpy via solid heat exchanger so that the specific impulse is limited up to 900 seconds due to the allowable maximum temperature for the heat exchanger. However, the strict optical alignment is not necessary for its operation, and this type is robust to the perturbation to the laser power density and the distribution that is expected during the flight. In the CW laser propulsion, the propellant is heated through the isobaric process, and a propellant pump for an additional compression process is necessary. Hence, the engine for CW laser propulsion is more complicated and then heavier than the RP laser propulsions. This is why no launch test has been accomplished until today.

High-power laser is the first priority to realize the laser launch system (LLS). In all the previous studies, except for Myrabo's campaign, the time-average power of the laser was around a few kilowatt is too low for the practical experiment. As mentioned above, 1 MW laser power is necessary to launch 1 kg payload. Hence, if you want to launch 1-ton payload as in the case of the conventional chemical rockets, you need to prepare 1 GW laser facility. This power is 10<sup>6</sup> times as high as in the previous experiments, and it would be natural to think that even the basic phenomena should be different for such high-power laser in the foregoing experiments. In the previous experiments, it is impossible to generate and maintain continuously a laser-supported detonation wave using CW laser, while it will become possible if one uses the MW or GW-class laser. Once the heating

Figure 5.

Figure 6.

77

Heat exchanger rocket.

LSP type CW laser propulsion.

Beamed Launch Propulsion

DOI: http://dx.doi.org/10.5772/intechopen.82236

Figure 4. Laser-supported detonation wave.

RP laser propulsion uses only gaseous propellant or only solid propellant, called the laser ablative propulsion, or the both. The gaseous propellant mostly used is the air atmosphere, and air-breathing propulsion concepts have been studied by many researchers. By focusing an intense laser pulse in the air, a laser-supported detonation wave is generated instantaneously to generate a blast wave around the optical focal point, as illustrated in Figure 4. The impulsive thrust is generated as recoil of the blast wave reflection in the nozzle. This method is a variation of the pulse detonation engine (PDE), which generates the thrust via isochoric heating and unsteady gas expansion. The source of the gas expansion, the LSD wave can be generated even in the hypersonic flow, and it can be applied in the air-breathing engines that can operate in hypersonic speeds.

For the CW laser propulsion, only gaseous propellants have been used, because strong momentum coupling from laser ablation requires intense and short laser pulse, and the power density from the CW lasers is too small to be used for the laser ablative propulsion. Moreover, no air-breathing engines have been studied for CW laser propulsion. Two different kinds of rocket were proposed. The laser-sustained plasma (LSP) engines, illustrated in Figure 5, use the plasma kept by laser absorption to heat the gaseous propellant running through it [13]. The strong point of this type is high specific impulse more than 1000 seconds because of the high temperature of plasma more than 10,000 K. Mystery remains in plasma stability to the perturbation of the laser power density and its distribution relative to the flows of propellant. Strict optical alignment is necessary for operation. Another kind is heat

#### Beamed Launch Propulsion DOI: http://dx.doi.org/10.5772/intechopen.82236

exchanger (HX) rocket proposed by Kare, which is illustrated in Figure 6 [14]. The laser power is converted to the propellant enthalpy via solid heat exchanger so that the specific impulse is limited up to 900 seconds due to the allowable maximum temperature for the heat exchanger. However, the strict optical alignment is not necessary for its operation, and this type is robust to the perturbation to the laser power density and the distribution that is expected during the flight. In the CW laser propulsion, the propellant is heated through the isobaric process, and a propellant pump for an additional compression process is necessary. Hence, the engine for CW laser propulsion is more complicated and then heavier than the RP laser propulsions. This is why no launch test has been accomplished until today.

High-power laser is the first priority to realize the laser launch system (LLS). In all the previous studies, except for Myrabo's campaign, the time-average power of the laser was around a few kilowatt is too low for the practical experiment. As mentioned above, 1 MW laser power is necessary to launch 1 kg payload. Hence, if you want to launch 1-ton payload as in the case of the conventional chemical rockets, you need to prepare 1 GW laser facility. This power is 10<sup>6</sup> times as high as in the previous experiments, and it would be natural to think that even the basic phenomena should be different for such high-power laser in the foregoing experiments. In the previous experiments, it is impossible to generate and maintain continuously a laser-supported detonation wave using CW laser, while it will become possible if one uses the MW or GW-class laser. Once the heating

Figure 5. LSP type CW laser propulsion.

RP laser propulsion uses only gaseous propellant or only solid propellant, called the laser ablative propulsion, or the both. The gaseous propellant mostly used is the air atmosphere, and air-breathing propulsion concepts have been studied by many researchers. By focusing an intense laser pulse in the air, a laser-supported detonation wave is generated instantaneously to generate a blast wave around the optical focal point, as illustrated in Figure 4. The impulsive thrust is generated as recoil of the blast wave reflection in the nozzle. This method is a variation of the pulse detonation engine (PDE), which generates the thrust via isochoric heating and unsteady gas expansion. The source of the gas expansion, the LSD wave can be generated even in the hypersonic flow, and it can be applied in the air-breathing engines that can

For the CW laser propulsion, only gaseous propellants have been used, because strong momentum coupling from laser ablation requires intense and short laser pulse, and the power density from the CW lasers is too small to be used for the laser ablative propulsion. Moreover, no air-breathing engines have been studied for CW laser propulsion. Two different kinds of rocket were proposed. The laser-sustained plasma (LSP) engines, illustrated in Figure 5, use the plasma kept by laser absorption to heat the gaseous propellant running through it [13]. The strong point of this type is high specific impulse more than 1000 seconds because of the high temperature of plasma more than 10,000 K. Mystery remains in plasma stability to the perturbation of the laser power density and its distribution relative to the flows of propellant. Strict optical alignment is necessary for operation. Another kind is heat

operate in hypersonic speeds.

Laser-supported detonation wave.

Figure 3. Sasoh's LITA.

Propulsion Systems

Figure 4.

76

Figure 6. Heat exchanger rocket.

mechanism is changed, the propulsion design will change. Evolution of super highpower CW laser will induce a new research domain of "high-power CW laser engineering."

On the other hand, it should be reasonable to assume that the minimum weight of launch vehicle should be heavier than 100 g. As the vehicle becomes smaller, the structural mass ratio is expected to increase to assure the structural strength. Hence, the minimum CW laser power for launch demonstrations should be around 100 kW. Of course, this estimate is quite rough, and the structural design of vehicle, propulsion performance, and trajectory plans should be considered for more detailed feasibility study. The launch system of laser power at 100 kW–1 MW is possible soon. There exists 100 kW fiber laser. It is technically possible to build a 1 MW laser by increasing the number of bundled fibers, and it is the matter of budget. Recently, high-power, energy-efficient, and compact fiber laser has been evolving. CW laser is easier to attain high power rather than RP laser especially in the case of fiber laser. If we could construct the LLS powered by CW laser, the launch demonstration will be completed at early times. As noted above, the design of LLS is the art of the integration in a broad area of engineering field. The relevant field includes propulsion, laser, beam transmission, flight dynamics, and control engineering.

Even after a high-power laser becomes available, a number of the engineering problems are remained. If we could have a proper propulsion system, we should determine the trajectory of a vehicle. The basic guidance law of conventional launch vehicles is the bilinear tangent law. Typical trajectory is illustrated in Figure 7. However, for LLS, we need to concern the special issue that a vehicle must stay on the laser beam. Kantrowitz assumed a circular trajectory at which center the laser source is located. It is illustrated in Figure 8. For such a trajectory, a good point is that the laser beam can always be irradiated onto the side surface of the vehicle fuselage. Similar trajectory was considered by Escape Dynamics Inc. Possible bad point is the unknown effects of the atmospheric turbulence on the laser beam transmission. Because the atmospheric turbulence is especially strong near the ground, the laser beam should be distorted drastically. This effect gives the degradation of the energy efficiency of the laser beam transmission and the engineering problem to keep a "laser link" between the vehicle and the ground laser facility. Katsurayama et al. proposed a zenith trajectory for LLS as illustrated in Figure 9

[15]. The zenith trajectory gives the minimum influence of the turbulence on the beam transmission. A vehicle is transferred via apogee-kick efficiently to the orbit around the earth. They consider using air-breathing engines, which can produce higher velocity increments on horizontal flight, and the optimum trajectory will be possibly determined from the trade-off between the effect of the air turbulence on the laser beam transmission and the air-breathing engine performance. Phipps et al.

solved the optimum trajectory for RP laser ablative rocket, without any air-

system design in this aspect is important for the feasibility of LLS.

Figure 8.

Figure 9.

79

A circular trajectory concept for laser launch.

Beamed Launch Propulsion

DOI: http://dx.doi.org/10.5772/intechopen.82236

A zenith trajectory concept for laser launch G.

breathing engine, from the ground to the orbit [16]. They commented simply to the effect of the air turbulence on the laser beam propagation. They concluded that the effect of the air turbulence is ignorable to the laser beam whose cross-sectional diameter is smaller than the "seeing size," which will be explained below; typical value is around 10 cm for the wavelength at 530 nm. They also suggested how to correct the pointing error due to the wave front tilt. More detailed estimation and

Moreover, the control mechanisms to keep the laser link between the ground and vehicle are indispensable to maintain the continuous energy supply via laser beam for the operation of the propulsion system. Sasoh's LITA and Myrabo's

Figure 7. Typical trajectory of conventional launch vehicles.

Beamed Launch Propulsion DOI: http://dx.doi.org/10.5772/intechopen.82236

mechanism is changed, the propulsion design will change. Evolution of super highpower CW laser will induce a new research domain of "high-power CW laser

On the other hand, it should be reasonable to assume that the minimum weight of launch vehicle should be heavier than 100 g. As the vehicle becomes smaller, the structural mass ratio is expected to increase to assure the structural strength. Hence,

100 kW. Of course, this estimate is quite rough, and the structural design of vehicle, propulsion performance, and trajectory plans should be considered for more detailed feasibility study. The launch system of laser power at 100 kW–1 MW is possible soon. There exists 100 kW fiber laser. It is technically possible to build a 1 MW laser by increasing the number of bundled fibers, and it is the matter of budget. Recently, high-power, energy-efficient, and compact fiber laser has been evolving. CW laser is easier to attain high power rather than RP laser especially in the case of fiber laser. If we could construct the LLS powered by CW laser, the launch demonstration will be completed at early times. As noted above, the design of LLS is the art of the integration in a broad area of engineering field. The relevant field includes propulsion, laser, beam transmission, flight dynamics, and control

Even after a high-power laser becomes available, a number of the engineering problems are remained. If we could have a proper propulsion system, we should determine the trajectory of a vehicle. The basic guidance law of conventional launch vehicles is the bilinear tangent law. Typical trajectory is illustrated in Figure 7. However, for LLS, we need to concern the special issue that a vehicle must stay on the laser beam. Kantrowitz assumed a circular trajectory at which center the laser source is located. It is illustrated in Figure 8. For such a trajectory, a good point is that the laser beam can always be irradiated onto the side surface of the vehicle fuselage. Similar trajectory was considered by Escape Dynamics Inc. Possible bad point is the unknown effects of the atmospheric turbulence on the laser beam transmission. Because the atmospheric turbulence is especially strong near the ground, the laser beam should be distorted drastically. This effect gives the degradation of the energy efficiency of the laser beam transmission and the engineering problem to keep a "laser link" between the vehicle and the ground laser facility. Katsurayama et al. proposed a zenith trajectory for LLS as illustrated in Figure 9

the minimum CW laser power for launch demonstrations should be around

engineering."

Propulsion Systems

engineering.

Figure 7.

78

Typical trajectory of conventional launch vehicles.

Figure 8. A circular trajectory concept for laser launch.

Figure 9. A zenith trajectory concept for laser launch G.

[15]. The zenith trajectory gives the minimum influence of the turbulence on the beam transmission. A vehicle is transferred via apogee-kick efficiently to the orbit around the earth. They consider using air-breathing engines, which can produce higher velocity increments on horizontal flight, and the optimum trajectory will be possibly determined from the trade-off between the effect of the air turbulence on the laser beam transmission and the air-breathing engine performance. Phipps et al. solved the optimum trajectory for RP laser ablative rocket, without any airbreathing engine, from the ground to the orbit [16]. They commented simply to the effect of the air turbulence on the laser beam propagation. They concluded that the effect of the air turbulence is ignorable to the laser beam whose cross-sectional diameter is smaller than the "seeing size," which will be explained below; typical value is around 10 cm for the wavelength at 530 nm. They also suggested how to correct the pointing error due to the wave front tilt. More detailed estimation and system design in this aspect is important for the feasibility of LLS.

Moreover, the control mechanisms to keep the laser link between the ground and vehicle are indispensable to maintain the continuous energy supply via laser beam for the operation of the propulsion system. Sasoh's LITA and Myrabo's

Lightcrafts are the concepts of the passive way to maintain the laser link. On the other hand, Phipps considered an active control of the optics onboard of the vehicle of the ablative launch system for the first time. Finally, the cooperative control between vehicle and beam pointing will be the natural solution for this issue. For the beam pointing control, the information of the beam position and vehicle position should be resolved precisely. The spatial resolution should be around 1 cm. We need to innovate a high-resolution method to measure vehicle and beam position. After considering these problems, it is clear that the propulsion design and the air turbulence effect on the beam transportation are the root problem.

be more complicated than solid rockets partly because the liquid propellant needs to be pressurized and pumped to the combustion chamber of rocket engines. The final price would be determined by the balance between structural complexity and

This can be explained in more quantitatively as follows. The launch cost of ELV mainly consists of the production cost of the launch vehicle and the propellant cost, if we could ignore the development of the system and the infrastructure maintenance of launch site. We shall start from the Tsiolkovsky rocket equation:

Here, ΔV is the velocity increments required to reach the orbit, Isp is the specific

<sup>1</sup> � exp <sup>Δ</sup><sup>V</sup>

exp <sup>Δ</sup><sup>V</sup> gIsp h i � <sup>ϵ</sup>

From this equation, it is proved that the launch cost for a unit payload mass decreases monotonously with Isp, while it decreases with ε. It is effective way to reduce the launch cost by increasing Isp and reducing the vehicle mass. This theoretical result is consistent to the VSLV designs by the ventures. Moreover, in order to ensure the structural strength of VSLV, ε increases inevitably as the vehicle size is reduced, increasing the launch cost. Here, we assume the single-stage launch, while the launch cost increases with the number of stages. Unfortunately, even when liquid propellants are used, it is quite difficult to realize the single-stage launch vehicle due to the limitation of chemical rocket Isp less than 460 seconds. Extreme reduction of launch cost can be attained by higher jump of Isp with extremely

This will be attained by using laser propulsion. By using only hydrogen as propellant, Isp can reach 900 seconds, which is limited by the allowable temperature limit of the engine materials, and the hydrogen temperature cannot exceed around 3000 K. The vehicle can be extremely simplified as the energy source is left on the ground. Laser propulsion can launch the payload of around 1 kg with a 1 MW laser facility. The maximum power of the available laser facility is 100 kW today. In principle, it is possible to develop a MW-class laser by bundling the fibers with the price of several tens million US dollars. Once it is developed, massive materials, though just 1 kg at a time, can be launched continuously and on demand to the orbit. The price of the 1 MW-class laser facility is almost the same level of a single launch of H-IIA rocket vehicle. Once a VSLV on demand is realized, the induction effect is expected for the technical breakthrough and the market expansion of small satellites. On the other hand, a GW-class laser facility is necessary to launch a payload of several tons at a time, which is typical launch capability of conventional launch systems. The development of such a huge laser facility will require an extremely large budget, and it would not be easy to realize in near future.

gIsp h i

impulse, mst is the structural mass, mprop is the propellant mass, and mpay is the payload mass. We shall define empty mass as mempty = mst + mprop. Structural mass ratio ε is defined as mst/mempty. Furthermore, we shall assume simply that the production cost of the vehicle is proportional to mst and the propellant cost is proportional to mprop. Then, the launch cost, C, is proportional to mempty. The constant of proportionality is defined as α. After several mathematical steps of

¼ exp

ΔV

gIsp � � (1)

(2)

mpay þ mst mpay þ mst þ mprop

Eq. (1), the launch cost for an unit payload mass is formulated as

C mpay

simplified structure of the vehicle, at a single stage.

81

¼ α

increase in the specific impulse.

DOI: http://dx.doi.org/10.5772/intechopen.82236

Beamed Launch Propulsion

#### 3. Motivation of laser propulsion

In order to expand the human activities in space, it is indispensable to drastically reduce the cost of transportation from ground to orbit. For an example of today's launch cost, the launch cost for a unit payload mass of the actual rocket of Japan, H-II, to launch a 6 ton payload to geo transfer orbit is around 20,000 \$/kg. Today the price competition is intense so that much of the same is the launch cost value in the USA and EU. Drastic reduction of the launch cost for the unit payload mass has been at stake for a long period. The space shuttle is the first attempt to reduce the launch cost. Shuttle was the first partly reusable launch vehicle. The orbiter was designed reusable to reduce the launch cost by using the orbiter repetitively at a high frequency. However, as it is well-known today, the space shuttle launch system was too huge and complex to reduce the launch cost due to the expensive maintenance. The Space Shuttle Project left the severe lessons for the engineers who still dream to develop a new reusable launch vehicle. Today, the expandable launch vehicles (ELVs) are still major way to the orbit, and the engineers are reducing the cost mainly by the standardization and the simplification. For an example, the next H-III rocket of Japan is claimed to halve the launch cost.

Falcon Heavy produced by SpaceX is a huge rocket that can deliver 26.7 ton to GTO, which is four times as heavy as H-IIA, reducing the launch cost for an unit payload mass around to 6000 \$/kg, which is around one-third of the H-IIA. This is not surprising. On the basis of the statistical data of the ELV developed so far, the launch cost of the ELV has a trend to decrease inversely with the vehicle size [6]. The launch cost for a unit payload mass draws a unique curve decreasing with the payload mass for the same launch mission. Falcon Heavy owes its low price to its large size. It is unclear if the price would continue to decrease with increasing the rocket size, like a huge launch vehicle for the interplanetary transport system designed by SpaceX. For the drastic cost cut, revolutionary breakthrough is necessary to compete the launch market.

To the opposite direction, the unit launch cost naturally increases with the decreasing the payload weight. Recently, R&D of small satellites is very active, and the nanosats (lighter than 10 kg) and picosats (<1 kg) will become in practice soon. Then, the demand for very small launch vehicle (VSLV) at reasonable cost is increasing. Several teams are now developing VLSV using liquid propellant. The Vector Space System Inc. is launching small-sat launch vehicle, which can deliver a 65 kg payload to LEO using liquid propellant rocket using propylene and LOX. The Interstellar Technologies Inc. is launching a gas pressure-pumped liquid propellant rocket called MOMO, while the Rocket Lab Inc. developed electrically pumped liquid rocket engines. In general, liquid propellant offers the specific impulse higher the solid propellant rockets. On the other hand, the liquid propellant rockets tend to Lightcrafts are the concepts of the passive way to maintain the laser link. On the other hand, Phipps considered an active control of the optics onboard of the vehicle of the ablative launch system for the first time. Finally, the cooperative control between vehicle and beam pointing will be the natural solution for this issue. For the beam pointing control, the information of the beam position and vehicle position should be resolved precisely. The spatial resolution should be around 1 cm. We need to innovate a high-resolution method to measure vehicle and beam position. After considering these problems, it is clear that the propulsion design and the air

In order to expand the human activities in space, it is indispensable to drastically reduce the cost of transportation from ground to orbit. For an example of today's launch cost, the launch cost for a unit payload mass of the actual rocket of Japan, H-II, to launch a 6 ton payload to geo transfer orbit is around 20,000 \$/kg. Today the price competition is intense so that much of the same is the launch cost value in the USA and EU. Drastic reduction of the launch cost for the unit payload mass has been at stake for a long period. The space shuttle is the first attempt to reduce the launch cost. Shuttle was the first partly reusable launch vehicle. The orbiter was designed reusable to reduce the launch cost by using the orbiter repetitively at a high frequency. However, as it is well-known today, the space shuttle launch system was too huge and complex to reduce the launch cost due to the expensive maintenance. The Space Shuttle Project left the severe lessons for the engineers who still dream to develop a new reusable launch vehicle. Today, the expandable launch vehicles (ELVs) are still major way to the orbit, and the engineers are reducing the cost mainly by the standardization and the simplification. For an example, the next

Falcon Heavy produced by SpaceX is a huge rocket that can deliver 26.7 ton to GTO, which is four times as heavy as H-IIA, reducing the launch cost for an unit payload mass around to 6000 \$/kg, which is around one-third of the H-IIA. This is not surprising. On the basis of the statistical data of the ELV developed so far, the launch cost of the ELV has a trend to decrease inversely with the vehicle size [6]. The launch cost for a unit payload mass draws a unique curve decreasing with the payload mass for the same launch mission. Falcon Heavy owes its low price to its large size. It is unclear if the price would continue to decrease with increasing the rocket size, like a huge launch vehicle for the interplanetary transport system designed by SpaceX. For the drastic cost cut, revolutionary breakthrough is neces-

To the opposite direction, the unit launch cost naturally increases with the decreasing the payload weight. Recently, R&D of small satellites is very active, and the nanosats (lighter than 10 kg) and picosats (<1 kg) will become in practice soon. Then, the demand for very small launch vehicle (VSLV) at reasonable cost is increasing. Several teams are now developing VLSV using liquid propellant. The Vector Space System Inc. is launching small-sat launch vehicle, which can deliver a 65 kg payload to LEO using liquid propellant rocket using propylene and LOX. The Interstellar Technologies Inc. is launching a gas pressure-pumped liquid propellant rocket called MOMO, while the Rocket Lab Inc. developed electrically pumped liquid rocket engines. In general, liquid propellant offers the specific impulse higher the solid propellant rockets. On the other hand, the liquid propellant rockets tend to

turbulence effect on the beam transportation are the root problem.

H-III rocket of Japan is claimed to halve the launch cost.

sary to compete the launch market.

80

3. Motivation of laser propulsion

Propulsion Systems

be more complicated than solid rockets partly because the liquid propellant needs to be pressurized and pumped to the combustion chamber of rocket engines. The final price would be determined by the balance between structural complexity and increase in the specific impulse.

This can be explained in more quantitatively as follows. The launch cost of ELV mainly consists of the production cost of the launch vehicle and the propellant cost, if we could ignore the development of the system and the infrastructure maintenance of launch site. We shall start from the Tsiolkovsky rocket equation:

$$\frac{m\_{\rm pay} + m\_{\rm st}}{m\_{\rm pay} + m\_{\rm st} + m\_{\rm prop}} = \exp\left[\frac{\Delta V}{gI\text{sp}}\right] \tag{1}$$

Here, ΔV is the velocity increments required to reach the orbit, Isp is the specific impulse, mst is the structural mass, mprop is the propellant mass, and mpay is the payload mass. We shall define empty mass as mempty = mst + mprop. Structural mass ratio ε is defined as mst/mempty. Furthermore, we shall assume simply that the production cost of the vehicle is proportional to mst and the propellant cost is proportional to mprop. Then, the launch cost, C, is proportional to mempty. The constant of proportionality is defined as α. After several mathematical steps of Eq. (1), the launch cost for an unit payload mass is formulated as

$$\frac{C}{m\_{\text{pay}}} = \alpha \frac{1 - \exp\left[\frac{\Delta V}{g^{I\text{sp}}}\right]}{\exp\left[\frac{\Delta V}{g^{I\text{sp}}}\right] - \epsilon} \tag{2}$$

From this equation, it is proved that the launch cost for a unit payload mass decreases monotonously with Isp, while it decreases with ε. It is effective way to reduce the launch cost by increasing Isp and reducing the vehicle mass. This theoretical result is consistent to the VSLV designs by the ventures. Moreover, in order to ensure the structural strength of VSLV, ε increases inevitably as the vehicle size is reduced, increasing the launch cost. Here, we assume the single-stage launch, while the launch cost increases with the number of stages. Unfortunately, even when liquid propellants are used, it is quite difficult to realize the single-stage launch vehicle due to the limitation of chemical rocket Isp less than 460 seconds. Extreme reduction of launch cost can be attained by higher jump of Isp with extremely simplified structure of the vehicle, at a single stage.

This will be attained by using laser propulsion. By using only hydrogen as propellant, Isp can reach 900 seconds, which is limited by the allowable temperature limit of the engine materials, and the hydrogen temperature cannot exceed around 3000 K. The vehicle can be extremely simplified as the energy source is left on the ground. Laser propulsion can launch the payload of around 1 kg with a 1 MW laser facility. The maximum power of the available laser facility is 100 kW today. In principle, it is possible to develop a MW-class laser by bundling the fibers with the price of several tens million US dollars. Once it is developed, massive materials, though just 1 kg at a time, can be launched continuously and on demand to the orbit. The price of the 1 MW-class laser facility is almost the same level of a single launch of H-IIA rocket vehicle. Once a VSLV on demand is realized, the induction effect is expected for the technical breakthrough and the market expansion of small satellites. On the other hand, a GW-class laser facility is necessary to launch a payload of several tons at a time, which is typical launch capability of conventional launch systems. The development of such a huge laser facility will require an extremely large budget, and it would not be easy to realize in near future.

#### 4. A new design of laser propulsion

Then, assuming that a 100-kW-class CW fiber laser is available, we shall consider how to build the LLS. At first, we need to expect the fluctuations of the laser power, the laser beam incident angle, and the cross-sectional distribution. For this reason, the heat exchanger type requires no precise optical alignment of the incident laser beam on the vehicle. Kare's concept of heat exchanger rocket is illustrated in Figure 6. This is similar to the microwave rocket concept illustrated by the Escape Dynamics Inc. in 2015. In principle, the nuclear thermal rocket, like NERVA, is a kind of heat exchanger rocket. The specific impulse of a thermal rocket is maximized by using propellant of minimum molecular weight, hydrogen. If the propellant temperature could reach at 3000 K, the specific impulse in vacuum becomes 900 seconds [17, 18]. Nevertheless, Kare's estimation of the properspecific impulse is around 600 seconds. In this concept, the laser beam is irradiated and absorbed on the side surface of the vehicle fuselage, where the temperature is highest, and the propellant is heated through the heat convection on the inner surface of the fuselage. The temperature of the propellant never exceeds the temperature of the outer surface. Because the maximum temperature of the outer surface of the heat exchanger limits the maximum temperature of the propellant, the specific impulse is limited by the thermal resistance of the materials for the outer surface. Because the atmosphere includes the oxygen, the oxidation resistance is also an important issue for the outer materials. Moreover, a large amount of black-body radiation is emitted from the outer surface and is dissipated in the air as a significant factor of the energy loss. Furthermore, this type of the heat exchanger requires quite narrow flow channel to assure high heat transfer rate, and this causes the significant pressure loss of the propellant in the heat exchanger.

An alternative design of heat exchanger rocket is illustrated in Figure 10. We shall consider the zenith angle launch similar to Katsurayama's concept. As mentioned earlier, the zenith angle launch minimizes the effect of the atmospheric turbulence on the laser beam transmission and the complexity of the guidance and control. The vehicle introduces the laser beam of high-power density from the bottom surface of vehicle. Due to the atmospheric turbulence, the laser beam is expanded and deflected. In addition to the propulsion system of high performance (efficient and high specific impulse), precise beam pointing to a small window on the bottom surface of the vehicle is another key issue. A thrust-vectoring gimbal is assumed here for the attitude control. The pointing control and guidance are performed in a cooperative system consisting the propulsion thrust, gimbal, and laser beam optical system on the ground. The vehicle position is always informed onto the ground station. This will be done by the GPS signal from the vehicle or the optical tracking system on the ground or on orbit. On the basis of the information, the beam direction is controlled on the ground. At the same time, the exact position of the laser spot is detected on the vehicle and informed to the ground station. The beam position is adjusted from the ground, while vehicle adjusts its position transverse to the beam direction. Although more detailed analysis and design are necessary, we shall leave this issue for the future work. Before this issue, it is critical to investigate the effect of the atmospheric turbulence on the laser beam propagation through the atmosphere. As mentioned later, the high-frequency fluctuation of the laser beam direction (scintillation) should be critical when the vehicle attains the altitude higher than 10 km. Unfortunately, there is no control technique real-time correction of the scintillation today. The astronomers are taking pictures of stars at short exposure time, by catching the instantaneous image. The real-time correction of the beam direction should be based on the adaptive optics, for which we need to

detect the atmospheric turbulence on the ray line between the ground station and the vehicle by some means. We need to develop anti-scintillation techniques in the near future. However, for the development of the launch system, we need to move forward step by step. It should be better to start from the aim to a 10 km altitude along the zenith angle trajectory. Myrabo reached at 71 m. The 10 km is a well worthful challenge. For the control and guidance techniques, before going to the cooperative control, it is realistic to accomplish more simple method of the active

For the propulsion, we shall investigate an externally heated rocket, similar to Kare's concept, using liquid hydrogen as propellant. For the externally heated rocket, the specific impulse is limited by the thermal and oxidation resistance of the high-temperature materials, while the propellant can be selected per request. The propulsion system is illustrated in Figure 11. In this new concept, the inside of the engine is separated via glass window from the outside. Porous material is filed inside of the engine. The porous material absorbs the laser power introduced across the window and coverts the laser energy to the enthalpy of the hydrogen gas, which passes through the porous material. One feature of the porous heat exchanger is high rate of heat convection and low-pressure loss when using high porosity and high heat-resistance porous material like DONACARBO Felt© Osaka Gas Chemical Inc. shown on the right of Figure 4. Because the diameter of carbon fiber is around 10 mm, large surface area and high rate of heat convection are attainable even at high-porosity condition, which leads to the low-pressure loss. Because the heat exchanger is in a close cell filled with the hydrogen, the maximum temperature of

bream-riding flight along a vertical trajectory.

Figure 10. Alternative LLS.

Beamed Launch Propulsion

DOI: http://dx.doi.org/10.5772/intechopen.82236

83

Beamed Launch Propulsion DOI: http://dx.doi.org/10.5772/intechopen.82236

4. A new design of laser propulsion

Propulsion Systems

Then, assuming that a 100-kW-class CW fiber laser is available, we shall consider how to build the LLS. At first, we need to expect the fluctuations of the laser power, the laser beam incident angle, and the cross-sectional distribution. For this reason, the heat exchanger type requires no precise optical alignment of the incident laser beam on the vehicle. Kare's concept of heat exchanger rocket is illustrated in Figure 6. This is similar to the microwave rocket concept illustrated by the Escape Dynamics Inc. in 2015. In principle, the nuclear thermal rocket, like NERVA, is a kind of heat exchanger rocket. The specific impulse of a thermal rocket is maximized by using propellant of minimum molecular weight, hydrogen. If the propellant temperature could reach at 3000 K, the specific impulse in vacuum becomes 900 seconds [17, 18]. Nevertheless, Kare's estimation of the properspecific impulse is around 600 seconds. In this concept, the laser beam is irradiated and absorbed on the side surface of the vehicle fuselage, where the temperature is highest, and the propellant is heated through the heat convection on the inner surface of the fuselage. The temperature of the propellant never exceeds the temperature of the outer surface. Because the maximum temperature of the outer surface of the heat exchanger limits the maximum temperature of the propellant, the specific impulse is limited by the thermal resistance of the materials for the outer surface. Because the atmosphere includes the oxygen, the oxidation resistance is also an important issue for the outer materials. Moreover, a large amount of black-body radiation is emitted from the outer surface and is dissipated in the air as a significant factor of the energy loss. Furthermore, this type of the heat exchanger requires quite narrow flow channel to assure high heat transfer rate, and this causes

the significant pressure loss of the propellant in the heat exchanger.

82

An alternative design of heat exchanger rocket is illustrated in Figure 10. We shall consider the zenith angle launch similar to Katsurayama's concept. As mentioned earlier, the zenith angle launch minimizes the effect of the atmospheric turbulence on the laser beam transmission and the complexity of the guidance and control. The vehicle introduces the laser beam of high-power density from the bottom surface of vehicle. Due to the atmospheric turbulence, the laser beam is expanded and deflected. In addition to the propulsion system of high performance (efficient and high specific impulse), precise beam pointing to a small window on the bottom surface of the vehicle is another key issue. A thrust-vectoring gimbal is assumed here for the attitude control. The pointing control and guidance are performed in a cooperative system consisting the propulsion thrust, gimbal, and laser beam optical system on the ground. The vehicle position is always informed onto the ground station. This will be done by the GPS signal from the vehicle or the optical tracking system on the ground or on orbit. On the basis of the information, the beam direction is controlled on the ground. At the same time, the exact position of the laser spot is detected on the vehicle and informed to the ground station. The beam position is adjusted from the ground, while vehicle adjusts its position transverse to the beam direction. Although more detailed analysis and design are necessary, we shall leave this issue for the future work. Before this issue, it is critical to investigate the effect of the atmospheric turbulence on the laser beam propagation through the atmosphere. As mentioned later, the high-frequency fluctuation of the laser beam direction (scintillation) should be critical when the vehicle attains the altitude higher than 10 km. Unfortunately, there is no control technique real-time correction of the scintillation today. The astronomers are taking pictures of stars at short exposure time, by catching the instantaneous image. The real-time correction of the beam direction should be based on the adaptive optics, for which we need to

detect the atmospheric turbulence on the ray line between the ground station and the vehicle by some means. We need to develop anti-scintillation techniques in the near future. However, for the development of the launch system, we need to move forward step by step. It should be better to start from the aim to a 10 km altitude along the zenith angle trajectory. Myrabo reached at 71 m. The 10 km is a well worthful challenge. For the control and guidance techniques, before going to the cooperative control, it is realistic to accomplish more simple method of the active bream-riding flight along a vertical trajectory.

For the propulsion, we shall investigate an externally heated rocket, similar to Kare's concept, using liquid hydrogen as propellant. For the externally heated rocket, the specific impulse is limited by the thermal and oxidation resistance of the high-temperature materials, while the propellant can be selected per request. The propulsion system is illustrated in Figure 11. In this new concept, the inside of the engine is separated via glass window from the outside. Porous material is filed inside of the engine. The porous material absorbs the laser power introduced across the window and coverts the laser energy to the enthalpy of the hydrogen gas, which passes through the porous material. One feature of the porous heat exchanger is high rate of heat convection and low-pressure loss when using high porosity and high heat-resistance porous material like DONACARBO Felt© Osaka Gas Chemical Inc. shown on the right of Figure 4. Because the diameter of carbon fiber is around 10 mm, large surface area and high rate of heat convection are attainable even at high-porosity condition, which leads to the low-pressure loss. Because the heat exchanger is in a close cell filled with the hydrogen, the maximum temperature of

ρuCp, <sup>g</sup>

Beamed Launch Propulsion

DOI: http://dx.doi.org/10.5772/intechopen.82236

dT<sup>g</sup>

been ignored. In a real engine, the energy efficiency can be enhanced by transforming the radiation to the gas enthalpy. For the heat convection in the porous media, the local thermal equilibrium (LTE) model (T<sup>g</sup> = Tp) is frequently used. On the other hand, we shall use more general local thermal nonequilibrium (LTNE) model. Actually, the temperature difference between the gas temperature, Tg, and the bulk material temperature,Tp, drives the heat convection. Note that all

The calculation results for laser power at 100 kW and hydrogen propellant are shown in Figure 13. The total temperature and the energy efficiency defined as the fraction of the laser power that is converted to the gas enthalpy are plotted as functions of the incident laser power density, IL0. δT is defined as IL0/ρuCp,g, which is equal to the temperature of the gas without energy loss. On the curve of a

constant δT, the mass flux ρu increases as the laser power density, IL0, increases. As is clear from the figure, the energy conversion efficiency increases monotonously

with IL0. In order to attain 3000 K, IL0 should be larger than 10<sup>9</sup> W/m2

means that the 100 kW laser beam is focused on a spot of the order of 1 cm.

the variables are calculated in SI unit.

Figure 13.

85

Result of propulsion model.

dx � hav <sup>T</sup><sup>p</sup> � <sup>T</sup><sup>g</sup>

Here,T<sup>p</sup> and T<sup>g</sup> are the temperature of the bulk material of porous media and the gas, respectively. x is the coordinate inside of the porous media as illustrated in Figure 12. ρ is the density of gas. u is the velocity of the gas. Cp,g is the specific heat at constant pressure of the gas. h is the volumetric interfacial heat transfer coefficient, for which a number of empirical models have been presented, and is the function of Reynolds number whose standard size is element size of the bulk material of the porous media. For the carbon fiber-based porous media, the standard size should be the diameter of the carbon fiber, which is around 10 mm. When hydrogen is used as working gas, because hydrogen has the largest mean free path among the species at certain pressure, the porous flow features high Knudsen number, and then the analysis could become complicated. The fiber size significantly affects the energy transfer processes. a<sup>v</sup> is the specific surface area of porous medium (surface area per unit volume), and qrad is light power irradiated by the black-body radiation from the element surface of porous medium, which is determined on the basis of the Stefan-Boltzmann law. I<sup>L</sup> is the laser power density. The scattering of the laser light, the scattering and reabsorption of radiation from hightemperature part, the reflection of radiation on the interface between the porous media and the engine wall, and the heat conduction inside of the porous media have

<sup>¼</sup> <sup>0</sup> (4)

, which

Figure 11. Schematic of volume absorber-type laser rocket engine.

Figure 12. Surface absorber vs. volume absorber.

the graphite materials will be more than 3000 K. (Sublimation point of graphite under anaerobic condition at one atmosphere pressure lies between 3895 and 4020 K [19].) Another feature of the porous media heat exchanger is the moderate absorption length that can be varied by the bulk size and the porosity of the porous material. This may lead to the reabsorption of the radiation from the hightemperature part of the porous material and the suppression of the energy loss due to radiation. By heating up the hydrogen gas to around 3000 K, the specific impulse in vacuum condition will reach at 900 seconds. In Figure 12, the "volume absorber" in the present concept is compared with the conventional Kare's concept of the socalled surface absorber. We shall consider the theoretical modeling of the volume absorber.

For the energy balance of the bulk material of porous media and the gas past the porous media, we shall consider the one-dimensional model as

$$
tau\_v \left( T\_\text{p} - T\_\text{g} \right) + \frac{\text{d}I\_\text{L}}{\text{d}\infty} + q\_{\text{rad}} = \mathbf{0} \tag{3}
$$

Beamed Launch Propulsion DOI: http://dx.doi.org/10.5772/intechopen.82236

$$
\rho u C\_{p, \text{g}} \frac{\text{d}T\_{\text{g}}}{\text{d} \text{x}} - h a\_v \left( T\_{\text{p}} - T\_{\text{g}} \right) = \mathbf{0} \tag{4}
$$

Here,T<sup>p</sup> and T<sup>g</sup> are the temperature of the bulk material of porous media and the gas, respectively. x is the coordinate inside of the porous media as illustrated in Figure 12. ρ is the density of gas. u is the velocity of the gas. Cp,g is the specific heat at constant pressure of the gas. h is the volumetric interfacial heat transfer coefficient, for which a number of empirical models have been presented, and is the function of Reynolds number whose standard size is element size of the bulk material of the porous media. For the carbon fiber-based porous media, the standard size should be the diameter of the carbon fiber, which is around 10 mm. When hydrogen is used as working gas, because hydrogen has the largest mean free path among the species at certain pressure, the porous flow features high Knudsen number, and then the analysis could become complicated. The fiber size significantly affects the energy transfer processes. a<sup>v</sup> is the specific surface area of porous medium (surface area per unit volume), and qrad is light power irradiated by the black-body radiation from the element surface of porous medium, which is determined on the basis of the Stefan-Boltzmann law. I<sup>L</sup> is the laser power density. The scattering of the laser light, the scattering and reabsorption of radiation from hightemperature part, the reflection of radiation on the interface between the porous media and the engine wall, and the heat conduction inside of the porous media have been ignored. In a real engine, the energy efficiency can be enhanced by transforming the radiation to the gas enthalpy. For the heat convection in the porous media, the local thermal equilibrium (LTE) model (T<sup>g</sup> = Tp) is frequently used. On the other hand, we shall use more general local thermal nonequilibrium (LTNE) model. Actually, the temperature difference between the gas temperature, Tg, and the bulk material temperature,Tp, drives the heat convection. Note that all the variables are calculated in SI unit.

The calculation results for laser power at 100 kW and hydrogen propellant are shown in Figure 13. The total temperature and the energy efficiency defined as the fraction of the laser power that is converted to the gas enthalpy are plotted as functions of the incident laser power density, IL0. δT is defined as IL0/ρuCp,g, which is equal to the temperature of the gas without energy loss. On the curve of a constant δT, the mass flux ρu increases as the laser power density, IL0, increases. As is clear from the figure, the energy conversion efficiency increases monotonously with IL0. In order to attain 3000 K, IL0 should be larger than 10<sup>9</sup> W/m2 , which means that the 100 kW laser beam is focused on a spot of the order of 1 cm.

Figure 13. Result of propulsion model.

the graphite materials will be more than 3000 K. (Sublimation point of graphite under anaerobic condition at one atmosphere pressure lies between 3895 and 4020 K [19].) Another feature of the porous media heat exchanger is the moderate absorption length that can be varied by the bulk size and the porosity of the porous

temperature part of the porous material and the suppression of the energy loss due to radiation. By heating up the hydrogen gas to around 3000 K, the specific impulse in vacuum condition will reach at 900 seconds. In Figure 12, the "volume absorber" in the present concept is compared with the conventional Kare's concept of the socalled surface absorber. We shall consider the theoretical modeling of the volume

For the energy balance of the bulk material of porous media and the gas past the

dI<sup>L</sup> dx

þ qrad ¼ 0 (3)

material. This may lead to the reabsorption of the radiation from the high-

porous media, we shall consider the one-dimensional model as

hav T<sup>p</sup> � Tg <sup>þ</sup>

absorber.

84

Figure 12.

Surface absorber vs. volume absorber.

Figure 11.

Propulsion Systems

Schematic of volume absorber-type laser rocket engine.

Moreover, because the mass flux is quite large, the Reynolds number and the Mach number of the flow into the porous media become 102 and 0.3, respectively, which are extraordinarily large numbers for the porous flows. The heat transfer model should be further investigated experimentally.

deflection index C<sup>n</sup>

Beamed Launch Propulsion

integrating C<sup>n</sup>

2 . C<sup>n</sup>

DOI: http://dx.doi.org/10.5772/intechopen.82236

isoplanatic angle θ<sup>0</sup> is defined as

distribution of C<sup>n</sup>

C<sup>n</sup> 2

following. The unit of C<sup>n</sup>

typical value of C<sup>n</sup>

87

<sup>A</sup> <sup>¼</sup> <sup>1</sup>:<sup>29</sup> � <sup>10</sup>�<sup>12</sup>r<sup>0</sup>

passing of an aircraft, it is not easy to predict C<sup>n</sup>

<sup>2</sup> along the beam direction as

r0 <sup>¼</sup> <sup>0</sup>:423k<sup>2</sup> sec ð Þ<sup>β</sup>

<sup>2</sup> is the function of the altitude, depending on the local

<sup>2</sup> precisely in general cases. Here,

<sup>2</sup> [20].

(6)

(7)

(9)

<sup>3</sup> <sup>þ</sup> <sup>A</sup>e�10<sup>h</sup> (8)

<sup>λ</sup><sup>2</sup> � <sup>3</sup>:<sup>89</sup> � <sup>10</sup>�<sup>15</sup> (10)

<sup>2</sup> unpredictable. According to

δz (11)

weather condition. Since it is sensitive to the instantaneous perturbation as the

Fried parameter (or called coherence length or seeing size) r<sup>0</sup> is defined by

ðL 0 C<sup>n</sup> 2 ð Þz dz

� ��3=<sup>5</sup>

Here, k is the wave number of the laser beam, sec is the secant (trigonometric function), β is the zenith angle, and L is the propagation distance. When a laser beam is transmitted from the ground to space, r<sup>0</sup> is the maximum beam diameter on the ground for the diffraction limited focusing in space. Even when the beam diameter is larger than r0, the spot diameter on space object is larger than the diffraction limit of r0. Hence, it is useless to increase the beam diameter on the ground larger than r0. Fried parameter equals to the typical size of the turbulence. Its typical value is around 10 cm for the visible light around for λ = 500 nm. Because

we shall choose well-known HV 7/5 model for the altitude distribution of C<sup>n</sup>

r<sup>0</sup> ∝ λ1.2, it is around 20 cm for λ = 1 μm. For the flatness of the wave front,

ðL 0 C<sup>n</sup> 2 ð Þ<sup>z</sup> <sup>z</sup><sup>5</sup>=<sup>3</sup> dz

θ<sup>0</sup> ∝ λ1.2, typical value is 16 μrad for λ = 1 μm. In HV 7/5 model, the altitude

The typical value for the visible light (λ = 500 nm) is 7 μrad, and considering

Here, z is the altitude from sea level [km], and h is the altitude from the beam source [km]. When the laser beam is emitted from sea level, z = h as assuming in the

> 75θ<sup>0</sup> �5=3 <sup>λ</sup><sup>2</sup> � <sup>0</sup>:<sup>14</sup> <sup>q</sup>

Here, the units of r0, θ0, and λ are cm, μrad, and μm. As shown in Figure 14, the

2 ð Þ<sup>z</sup> � �<sup>6</sup>=<sup>5</sup>

k<sup>2</sup>=<sup>5</sup> z11<sup>=</sup><sup>5</sup>

<sup>λ</sup><sup>2</sup> � <sup>1</sup>:<sup>61</sup> � <sup>10</sup>�<sup>13</sup>θ<sup>0</sup>

10�<sup>17</sup> m�2/3 at altitude of 10 km. Then, it becomes almost constant. For the laser beam propagation from the ground to the sky, the atmospheric turbulence has significant impact at altitude lower than 10 km. In the actual atmosphere, the atmospheric boundary layer, typically lower than 2 km, is quite effective to the fluctuations in the laser beam. On the other hand, the conditions in the boundary layer depend on the local landform and are time-varying even in 1 day. This

ffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffiffi

<sup>2</sup> is 10�<sup>15</sup> m�2/3 near the ground and is reduced sharply to

�5=3

� ��3=<sup>5</sup>

<sup>2</sup> is formulated as follows taking r<sup>0</sup> and θ<sup>0</sup> as major parameters:

<sup>27</sup> � � <sup>þ</sup> <sup>2</sup>:<sup>7</sup> � <sup>10</sup>�16e�2<sup>z</sup>

(z) is m�2/3. W and A are the constants that represent the

<sup>θ</sup><sup>0</sup> <sup>¼</sup> <sup>2</sup>:91k<sup>2</sup>

ð Þ¼ <sup>z</sup> <sup>5</sup>:<sup>94</sup> � <sup>10</sup>�<sup>23</sup>z10e�<sup>z</sup> <sup>W</sup>

W ¼ 27

�5=3

complexity in atmospheric boundary layer makes C<sup>n</sup>

<sup>δ</sup>w<sup>2</sup> <sup>¼</sup> <sup>128</sup>

Ref. [21], beam expansion of a Gaussian beam δw is formulated as

<sup>5</sup> <sup>0</sup>:545Cn

2

atmospheric condition, formulated using (r0, θ0) as

#### 5. Beam transmission through the atmosphere

The laser beam transmission through the atmosphere is a critical issue for the feasibility of LLS. For the LLS, the laser beam propagates from 0 to 100 km across the atmosphere to point continuously and precisely on a vehicle. The laser beam expands both due to diffraction and atmospheric turbulence. In the studies of the LLS, to the best of author's knowledge, as noted above, only Phipps et al. have considered the atmospheric beam transmission simply using the Fried parameter. Several studies considered the solar power satellite, SPS, using laser beam though there is no systematic study for the beam transmission through the atmospheric turbulence. This is partly because relevant theory has not been fully developed. The light wave propagation through the turbulent atmosphere has been studied in the field of astronomy in terms of the adaptive optics [20]. Only numerical simulations on the basis of the random phase screen method is useful for the exact analyses. To be exact, the atmospheric turbulence depends on local and instantaneous weather conditions. For a particular launch site, the numerical studies and the launch tests are necessary to verify the local effects of the atmospheric turbulence on the laser beam transmission. In some cases, it is necessary to apply the adaptive optics (AO) techniques like the large telescope like Subaru. On the other hand, qualitative discussion is also useful for typical cases, but it is not easy on the basis of numerical simulations. The analytical formula for the effects of the air turbulence on the Gaussian beam was found in 2009 [21]. What is necessary now is to know how and how much the air turbulence can affect the beam propagation. In this section, we shall discuss the impact of air turbulence on the laser beam propagation qualitatively on the basis of recent result in the research field of AO.

In order to attain high-transmission efficiency, the laser spot on the vehicle is adjusted to a proper size. The beam diffraction is regulated using the focusing optics. From the formula of diffraction limit, the minimum spot diameter of a Gaussian beam, ds, is formulated as a function of the propagated distance z and wavelength λ, the beam quality factor at M<sup>2</sup> , and the beam diameter on the source of beam, ϕ0:

$$d\_s = 1.2 \frac{z \lambda M^2}{\Phi\_0} \tag{5}$$

Substituting z = 100 km, λ = 1 μm, M<sup>2</sup> = 1.1 (typical value for fiber lasers), and ϕ<sup>0</sup> = 36 cm, we get d<sup>s</sup> � 37 cm. This means that a straight beam of 30 cm can be built.

The effect of the atmospheric turbulence on the laser beam propagation is categorized: (1) scintillation, (2) beam expansion, and (3) beam wonder. Scintillation is the so-called twinkle of stars. The intensity of the light varies unsteadily at high frequency. Beam expansion means the additional expansion of the beam diameter after propagating on a long path across the atmosphere. Beam wonder means the variation of the center of the laser beam axis on the cross-sectional surface. This is induced by the additional angular deflection of the laser beam due to the air turbulence. These effects are originated from the fluctuation in the deflection index distributed in the atmosphere; described using the structure constant of

#### Beamed Launch Propulsion DOI: http://dx.doi.org/10.5772/intechopen.82236

Moreover, because the mass flux is quite large, the Reynolds number and the Mach number of the flow into the porous media become 102 and 0.3, respectively, which are extraordinarily large numbers for the porous flows. The heat transfer model

The laser beam transmission through the atmosphere is a critical issue for the feasibility of LLS. For the LLS, the laser beam propagates from 0 to 100 km across the atmosphere to point continuously and precisely on a vehicle. The laser beam expands both due to diffraction and atmospheric turbulence. In the studies of the LLS, to the best of author's knowledge, as noted above, only Phipps et al. have considered the atmospheric beam transmission simply using the Fried parameter. Several studies considered the solar power satellite, SPS, using laser beam though there is no systematic study for the beam transmission through the atmospheric turbulence. This is partly because relevant theory has not been fully developed. The light wave propagation through the turbulent atmosphere has been studied in the field of astronomy in terms of the adaptive optics [20]. Only numerical simulations on the basis of the random phase screen method is useful for the exact analyses. To be exact, the atmospheric turbulence depends on local and instantaneous weather conditions. For a particular launch site, the numerical studies and the launch tests are necessary to verify the local effects of the atmospheric turbulence on the laser beam transmission. In some cases, it is necessary to apply the adaptive optics (AO) techniques like the large telescope like Subaru. On the other hand, qualitative discussion is also useful for typical cases, but it is not easy on the basis of numerical simulations. The analytical formula for the effects of the air turbulence on the Gaussian beam was found in 2009 [21]. What is necessary now is to know how and how much the air turbulence can affect the beam propagation. In this section, we shall discuss the impact of air turbulence on the laser beam propagation qualita-

should be further investigated experimentally.

Propulsion Systems

5. Beam transmission through the atmosphere

tively on the basis of recent result in the research field of AO.

wavelength λ, the beam quality factor at M<sup>2</sup>

of beam, ϕ0:

built.

86

In order to attain high-transmission efficiency, the laser spot on the vehicle is adjusted to a proper size. The beam diffraction is regulated using the focusing optics. From the formula of diffraction limit, the minimum spot diameter of a Gaussian beam, ds, is formulated as a function of the propagated distance z and

ds ¼ 1:2

zλM<sup>2</sup> ϕ0

Substituting z = 100 km, λ = 1 μm, M<sup>2</sup> = 1.1 (typical value for fiber lasers), and ϕ<sup>0</sup> = 36 cm, we get d<sup>s</sup> � 37 cm. This means that a straight beam of 30 cm can be

The effect of the atmospheric turbulence on the laser beam propagation is categorized: (1) scintillation, (2) beam expansion, and (3) beam wonder. Scintillation is the so-called twinkle of stars. The intensity of the light varies unsteadily at high frequency. Beam expansion means the additional expansion of the beam diameter after propagating on a long path across the atmosphere. Beam wonder means the variation of the center of the laser beam axis on the cross-sectional surface. This is induced by the additional angular deflection of the laser beam due to the air turbulence. These effects are originated from the fluctuation in the deflection index distributed in the atmosphere; described using the structure constant of

, and the beam diameter on the source

(5)

deflection index C<sup>n</sup> 2 . C<sup>n</sup> <sup>2</sup> is the function of the altitude, depending on the local weather condition. Since it is sensitive to the instantaneous perturbation as the passing of an aircraft, it is not easy to predict C<sup>n</sup> <sup>2</sup> precisely in general cases. Here, we shall choose well-known HV 7/5 model for the altitude distribution of C<sup>n</sup> <sup>2</sup> [20].

Fried parameter (or called coherence length or seeing size) r<sup>0</sup> is defined by integrating C<sup>n</sup> <sup>2</sup> along the beam direction as

$$\mathbf{r}\_0 = \left[\mathbf{0}.423k^2 \sec\left(\beta\right) \int\_0^L \mathbf{C}\_{\mathbf{n}}^2(z) d\mathbf{z}\right]^{-3/5} \tag{6}$$

Here, k is the wave number of the laser beam, sec is the secant (trigonometric function), β is the zenith angle, and L is the propagation distance. When a laser beam is transmitted from the ground to space, r<sup>0</sup> is the maximum beam diameter on the ground for the diffraction limited focusing in space. Even when the beam diameter is larger than r0, the spot diameter on space object is larger than the diffraction limit of r0. Hence, it is useless to increase the beam diameter on the ground larger than r0. Fried parameter equals to the typical size of the turbulence. Its typical value is around 10 cm for the visible light around for λ = 500 nm. Because r<sup>0</sup> ∝ λ1.2, it is around 20 cm for λ = 1 μm. For the flatness of the wave front, isoplanatic angle θ<sup>0</sup> is defined as

$$\Theta\_0 = \left[ 2.91k^2 \int\_0^L \mathbf{C}\_n^2(\mathbf{z}) \mathbf{z}^{5/3} \mathbf{d} \mathbf{z} \right]^{-3/5} \tag{7}$$

The typical value for the visible light (λ = 500 nm) is 7 μrad, and considering θ<sup>0</sup> ∝ λ1.2, typical value is 16 μrad for λ = 1 μm. In HV 7/5 model, the altitude distribution of C<sup>n</sup> <sup>2</sup> is formulated as follows taking r<sup>0</sup> and θ<sup>0</sup> as major parameters:

$$\left(\text{C}\_{\text{n}}\,^{2}\right) = 5.94 \cdot 10^{-23} z^{10} \text{e}^{-x} \left(\frac{\text{W}}{27}\right) + 2.7 \cdot 10^{-16} \text{e}^{-\frac{2\pi}{3}} + A \text{e}^{-10h} \tag{8}$$

Here, z is the altitude from sea level [km], and h is the altitude from the beam source [km]. When the laser beam is emitted from sea level, z = h as assuming in the following. The unit of C<sup>n</sup> 2 (z) is m�2/3. W and A are the constants that represent the atmospheric condition, formulated using (r0, θ0) as

$$\mathcal{W} = 2\Im\sqrt{75\theta\_0^{-5/3}\lambda^2 - 0.14} \tag{9}$$

$$A = \mathbf{1.29} \cdot \mathbf{10}^{-12} r\_0^{-5/3} \lambda^2 - \mathbf{1.61} \cdot \mathbf{10}^{-13} \theta\_0^{-5/3} \lambda^2 - \mathbf{3.89} \cdot \mathbf{10}^{-15} \tag{10}$$

Here, the units of r0, θ0, and λ are cm, μrad, and μm. As shown in Figure 14, the typical value of C<sup>n</sup> <sup>2</sup> is 10�<sup>15</sup> m�2/3 near the ground and is reduced sharply to 10�<sup>17</sup> m�2/3 at altitude of 10 km. Then, it becomes almost constant. For the laser beam propagation from the ground to the sky, the atmospheric turbulence has significant impact at altitude lower than 10 km. In the actual atmosphere, the atmospheric boundary layer, typically lower than 2 km, is quite effective to the fluctuations in the laser beam. On the other hand, the conditions in the boundary layer depend on the local landform and are time-varying even in 1 day. This complexity in atmospheric boundary layer makes C<sup>n</sup> <sup>2</sup> unpredictable. According to Ref. [21], beam expansion of a Gaussian beam δw is formulated as

$$
\delta w^2 = \frac{128}{5} \left( 0.545 \text{C}\_n^2(\text{z}) \right)^{6/5} k^{2/5} \text{z}^{11/5} \delta \text{z} \tag{11}
$$

Figure 14. Examples of C<sup>n</sup> <sup>2</sup> and δw for λ = 1 μm, θ<sup>0</sup> = 16.1 μrad, and r0 = 23.0 cm.


#### Table 1.

Beam expansion and wondering at altitude of 100 km.

Moreover, the angular fluctuation due to the beam wondering is formulated as

$$
\delta \alpha^2 = 0.364 \left( \frac{\phi\_0}{r\_0} \right)^{5/3} \left( \frac{\lambda}{\phi\_0} \right)^2 \tag{12}
$$

the laser beam transmission. The latter problem will become important in the near future when the laser launch vehicle can reach the altitude higher than 10 km. This problem is linked with the methodologies for the guidance and control of vehicle. Future studies will clarify the design features and technical problems of LLS in

more detail.

Beamed Launch Propulsion

DOI: http://dx.doi.org/10.5772/intechopen.82236

Author details

Nagoya University, Nagoya, Japan

provided the original work is properly cited.

\*Address all correspondence to: koichi.mori@mae.nagoya-u.ac.jp

© 2019 The Author(s). Licensee IntechOpen. This chapter is distributed under the terms of the Creative Commons Attribution License (http://creativecommons.org/licenses/ by/3.0), which permits unrestricted use, distribution, and reproduction in any medium,

Koichi Mori

89

As shown in Figure 14, δw saturates at the altitude around 10 km.

The results are summarized in Table 1. The beam diameter on the ground is assumed 10 cm. The increment of the beam diameter δw is 1.7 cm (around 17%) at the altitude of 100 km. This should be regulated using the optics on the ground. The beam wondering is around 3 μrad. This means the beam position is deflected by 30 cm at the altitude of 100 km. Because the beam diameter is 10 cm and then the size of the beam-receiving surface on the vehicle is almost same, the fluctuation of the beam location at 30 cm is quite large. This fluctuation varies typically at a frequency of 1 kHz. Without any correction to the beam wondering, the thrust cannot be generated at altitude of the order of several tens kilo-meter. On the other hand, the beam wondering should be ignorable in the demonstration of launch up to the altitude of 1 km. Both r<sup>0</sup> and θ<sup>0</sup> increase with λ. Consequently, although δw decreases slightly with λ, δα is constant.

#### 6. Summary

It is clear that we need to develop a launch system of high specific impulse to expand our universe. Laser launch system (LLS) is a promising candidate that can generate the specific impulse higher than 900 seconds. As a 100-kW-class fiber laser has been developed today, actual launch to the orbit will happen in near future. In this chapter, we looked around the technical problems and tried some analyses for the propulsion performance and the atmospheric turbulence effect on Beamed Launch Propulsion DOI: http://dx.doi.org/10.5772/intechopen.82236

the laser beam transmission. The latter problem will become important in the near future when the laser launch vehicle can reach the altitude higher than 10 km. This problem is linked with the methodologies for the guidance and control of vehicle. Future studies will clarify the design features and technical problems of LLS in more detail.

### Author details

Moreover, the angular fluctuation due to the beam wondering is formulated as

λ (μm) ϕ<sup>0</sup> (cm) r<sup>0</sup> (cm) θ<sup>0</sup> (μrad) δα (μrad) δw (cm) 0.5 10 10 7 3.0 2.0 1 10 23 16 3.0 1.7 10 10 364 254 3.0 1.1

r0

The results are summarized in Table 1. The beam diameter on the ground is assumed 10 cm. The increment of the beam diameter δw is 1.7 cm (around 17%) at the altitude of 100 km. This should be regulated using the optics on the ground. The beam wondering is around 3 μrad. This means the beam position is deflected by 30 cm at the altitude of 100 km. Because the beam diameter is 10 cm and then the size of the beam-receiving surface on the vehicle is almost same, the fluctuation of the beam location at 30 cm is quite large. This fluctuation varies typically at a frequency of 1 kHz. Without any correction to the beam wondering, the thrust cannot be generated at altitude of the order of several tens kilo-meter. On the other hand, the beam wondering should be ignorable in the demonstration of launch up to the altitude of 1 km. Both r<sup>0</sup> and θ<sup>0</sup> increase with λ. Consequently, although δw

It is clear that we need to develop a launch system of high specific impulse to expand our universe. Laser launch system (LLS) is a promising candidate that can generate the specific impulse higher than 900 seconds. As a 100-kW-class fiber laser has been developed today, actual launch to the orbit will happen in near future. In this chapter, we looked around the technical problems and tried some analyses for the propulsion performance and the atmospheric turbulence effect on

<sup>5</sup>=<sup>3</sup> λ

ϕ0 <sup>2</sup>

(12)

δα<sup>2</sup> <sup>¼</sup> <sup>0</sup>:<sup>364</sup> <sup>ϕ</sup><sup>0</sup>

<sup>2</sup> and δw for λ = 1 μm, θ<sup>0</sup> = 16.1 μrad, and r0 = 23.0 cm.

As shown in Figure 14, δw saturates at the altitude around 10 km.

decreases slightly with λ, δα is constant.

Beam expansion and wondering at altitude of 100 km.

6. Summary

88

Figure 14. Examples of C<sup>n</sup>

Propulsion Systems

Table 1.

Koichi Mori Nagoya University, Nagoya, Japan

\*Address all correspondence to: koichi.mori@mae.nagoya-u.ac.jp

© 2019 The Author(s). Licensee IntechOpen. This chapter is distributed under the terms of the Creative Commons Attribution License (http://creativecommons.org/licenses/ by/3.0), which permits unrestricted use, distribution, and reproduction in any medium, provided the original work is properly cited.

#### References

[1] Pournelle J, Nieven L. Foot Fall. New York: Del Rey, Penguin Random House; 1985

[2] Wilson RC. New York: Burning Paradise, Tor Books; 2013

[3] https://en.wikipedia.org/wiki/Brea kthrough\_Starshot

[4] https://en.wikipedia.org/wiki/Spacebased\_solar\_power

[5] https://en.wikipedia.org/wiki/Lase r\_communication\_in\_space

[6] Ketsdever AD et al. Overview of advanced concepts for space access. Journal of Spacecraft and Rockets. 2010; 47:238-250

[7] Kantrowitz A. Laser population. Astronautics and Aeronautics. 1971; 10:74

[8] Myrabo LN. World record flights of beam-riding rocket lightcraft: Demonstration of disruptive propulsion technology. AIAA Paper 01-3798; 2001

[9] Komurasaki K, Wang B. Laser propulsion. In: Encyclopedia of Aerospace Engineering. New Jersey, USA: Wiley & Sons, Ltd, Hoboken; 2010

[10] Phipps C et al. Review: Laserablation propulsion. Journal of Propulsion and Power. 2010;26(4): 609-637

[11] Tran TD, Yogo A, Nishimura H, Mori K. Impulse and mass removal rate of aluminum target by nanosecond laser ablation in a wide range of ambient pressure. Journal of Applied Physics. 2017;122:233304

[12] Xie C, Tran DT, Mori K. Numerical estimation of laser-ablation propulsion performance in spherical capsule. In:

31st International Symposium on Shock Waves. Nagoya, Japan. No. SBM000190; 2017

Chapter 5

Abstract

1. Introduction

91

Hall Thruster Erosion

and results of tests with different ceramic are presented.

the requirements of the technical task [1].

Andrey Vitalievich Loyan and Alona Nikolaevna Khaustova

Hall thruster (HT) is one of the thrusters that are systematically applied in space.

If to compare HT with plasma ion thrusters, it has lower lifetime and specific impulse. HT has a set of advantages, and that is why interest to this plasma thruster is high. It has relatively simple design and technology of production. HT does not require a complex power supply unit, and it is very important for spacecraft. Propulsion system on the base of HT has lower mass, simpler technology, and less time of production. One of the main HT characteristics that require improvement is the lifetime of thruster. As it is known, one of the main factors that decrease thruster lifetime is the wear of discharge chamber (DCh). With the analysis of demands to HT, it is understandable that the required lifetime is more than 10 years. So the question about lifetime of the HT is still open. This chapter presents the overview of the thruster elements lifetimes and the overview of methods of thruster erosion investigation. It shows advantages and disadvantages of optical methods of DCh erosion rate investigation. Chapter presents modified method of optical investigation. The results of HT research under various modes of operation

Keywords: electric propulsion thruster, stationary plasma thruster, lifetime, discharge chamber, method diagnostics, optical emission spectroscopy, erosion rate

With the increasing of number and complexity of tasks performed by modern spacecrafts, more high requirements appear for the propulsion unit (PU). Today, units based on hall thruster (HT) are one of a promising type of PU. Modern HT meets customer requirements by the following parameters: efficiency, specific impulse, and thrust price. However, the claimed thruster lifetime cannot fully meet

The common symptom of the completion of the HT's lifetime is the moment, when the elements of the magnetic system, the cathode, are bombarded by the ion flow first time, and this prose is followed by degradation of integral characteristics. To estimate the technical state of the discharge chamber (DCh) wear, the HT tests

The problem of research of DCh wear is based on two factors. First, the maximum number of HT orders for a single customer does not exceed 10 units per year. To use the classical mathematical statistics, volumes of samples should count dozens of samples and tests should last hundreds thousands hours. To determine the HT lifetime, which is 4000 hours, it is necessary to provide tests of 130 thousand hours for 32 thrusters, which exceeds the cost of PU. The dissemination of the test results

and measures of the erosion of the material are carried out [1, 2].

[13] Legner HH, Douglas-Hamilton DH. CW laser propulsion. Journal of Energy. 1978;2(2):85-94

[14] Kare JT. Laser-powered heat exchanger rocket for ground-to-orbit launch. Journal of Propulsion and Power. 1995;11(3):535-543

[15] Katsurayama H, Komurasaki K, Arakawa Y. A preliminary study of pulse-laser powered orbital launcher. Acta Astronautica. 2009;65:1032-1041

[16] Phipps CR, Reilly JP, Campbell JW. Optimum parameters for laser-launching objects into low earth orbit. Laser and Particle Beams. 2000;18(4):661-695

[17] Robbins WH, Finger HB. An Historical Perspective of the NERVA Nuclear Rocket Engine Technology Program. NASA Contractor Report 187154, AIAA-91-3451, Prepared fro Lewis Research Center Under Contract NAS3-25266, 1991

[18] Gabrielli RA, Herdrich G. Review of nuclear thermal propulsion systems. Progress in Aerospace Sciences. 2015;79: 92-113

[19] Abrahamson J. Graphite sublimation temperatures, carbon arcs and crystallite erosion. Carbon. 1974;12(2):111-118

[20] Tyson RK. Principles of Adaptive Optics. 4th ed. Florida, USA: CRC Press, Boca Raton; 2015

[21] Ji X, Li X. Directionality of Gaussian array beam propagating in atmosphere turbulence. Journal of the Optical Society of America A. 2009;26(2): 236-243

## Chapter 5 Hall Thruster Erosion

Andrey Vitalievich Loyan and Alona Nikolaevna Khaustova

### Abstract

References

Propulsion Systems

1985

[1] Pournelle J, Nieven L. Foot Fall. New York: Del Rey, Penguin Random House;

31st International Symposium on Shock

[13] Legner HH, Douglas-Hamilton DH. CW laser propulsion. Journal of Energy.

[14] Kare JT. Laser-powered heat exchanger rocket for ground-to-orbit launch. Journal of Propulsion and Power. 1995;11(3):535-543

[15] Katsurayama H, Komurasaki K, Arakawa Y. A preliminary study of pulse-laser powered orbital launcher. Acta Astronautica. 2009;65:1032-1041

[16] Phipps CR, Reilly JP, Campbell JW. Optimum parameters for laser-launching objects into low earth orbit. Laser and Particle Beams. 2000;18(4):661-695

[18] Gabrielli RA, Herdrich G. Review of nuclear thermal propulsion systems. Progress in Aerospace Sciences. 2015;79:

[19] Abrahamson J. Graphite sublimation temperatures, carbon arcs and crystallite erosion. Carbon. 1974;12(2):111-118

[20] Tyson RK. Principles of Adaptive Optics. 4th ed. Florida, USA: CRC Press,

[21] Ji X, Li X. Directionality of Gaussian array beam propagating in atmosphere turbulence. Journal of the Optical Society of America A. 2009;26(2):

[17] Robbins WH, Finger HB. An Historical Perspective of the NERVA Nuclear Rocket Engine Technology Program. NASA Contractor Report 187154, AIAA-91-3451, Prepared fro Lewis Research Center Under Contract

NAS3-25266, 1991

Boca Raton; 2015

236-243

92-113

Waves. Nagoya, Japan. No.

SBM000190; 2017

1978;2(2):85-94

[2] Wilson RC. New York: Burning

[3] https://en.wikipedia.org/wiki/Brea

[4] https://en.wikipedia.org/wiki/Space-

[5] https://en.wikipedia.org/wiki/Lase

[6] Ketsdever AD et al. Overview of advanced concepts for space access. Journal of Spacecraft and Rockets. 2010;

[7] Kantrowitz A. Laser population. Astronautics and Aeronautics. 1971;

[8] Myrabo LN. World record flights of

beam-riding rocket lightcraft: Demonstration of disruptive propulsion technology. AIAA Paper

[9] Komurasaki K, Wang B. Laser propulsion. In: Encyclopedia of Aerospace Engineering. New Jersey, USA: Wiley & Sons, Ltd, Hoboken;

[10] Phipps C et al. Review: Laserablation propulsion. Journal of Propulsion and Power. 2010;26(4):

[11] Tran TD, Yogo A, Nishimura H, Mori K. Impulse and mass removal rate of aluminum target by nanosecond laser ablation in a wide range of ambient pressure. Journal of Applied Physics.

[12] Xie C, Tran DT, Mori K. Numerical estimation of laser-ablation propulsion performance in spherical capsule. In:

r\_communication\_in\_space

Paradise, Tor Books; 2013

kthrough\_Starshot

based\_solar\_power

47:238-250

01-3798; 2001

10:74

2010

609-637

90

2017;122:233304

Hall thruster (HT) is one of the thrusters that are systematically applied in space. If to compare HT with plasma ion thrusters, it has lower lifetime and specific impulse. HT has a set of advantages, and that is why interest to this plasma thruster is high. It has relatively simple design and technology of production. HT does not require a complex power supply unit, and it is very important for spacecraft. Propulsion system on the base of HT has lower mass, simpler technology, and less time of production. One of the main HT characteristics that require improvement is the lifetime of thruster. As it is known, one of the main factors that decrease thruster lifetime is the wear of discharge chamber (DCh). With the analysis of demands to HT, it is understandable that the required lifetime is more than 10 years. So the question about lifetime of the HT is still open. This chapter presents the overview of the thruster elements lifetimes and the overview of methods of thruster erosion investigation. It shows advantages and disadvantages of optical methods of DCh erosion rate investigation. Chapter presents modified method of optical investigation. The results of HT research under various modes of operation and results of tests with different ceramic are presented.

Keywords: electric propulsion thruster, stationary plasma thruster, lifetime, discharge chamber, method diagnostics, optical emission spectroscopy, erosion rate

#### 1. Introduction

With the increasing of number and complexity of tasks performed by modern spacecrafts, more high requirements appear for the propulsion unit (PU). Today, units based on hall thruster (HT) are one of a promising type of PU. Modern HT meets customer requirements by the following parameters: efficiency, specific impulse, and thrust price. However, the claimed thruster lifetime cannot fully meet the requirements of the technical task [1].

The common symptom of the completion of the HT's lifetime is the moment, when the elements of the magnetic system, the cathode, are bombarded by the ion flow first time, and this prose is followed by degradation of integral characteristics. To estimate the technical state of the discharge chamber (DCh) wear, the HT tests and measures of the erosion of the material are carried out [1, 2].

The problem of research of DCh wear is based on two factors. First, the maximum number of HT orders for a single customer does not exceed 10 units per year. To use the classical mathematical statistics, volumes of samples should count dozens of samples and tests should last hundreds thousands hours. To determine the HT lifetime, which is 4000 hours, it is necessary to provide tests of 130 thousand hours for 32 thrusters, which exceeds the cost of PU. The dissemination of the test results from individual samples to a whole range of thrusters, even with one design, is unreasonable because of the lack of statistical data and high requirements for space technology.

Second, while using the reduced lifetime tests, the duration of the experiment is about 10% of the physical thruster lifetime. For example, for HT with 4000 hours of operation time, a test with a minimum base of 400 hours is required. The cost of such tests can be compared to the cost of one HT.

These factors determine the necessity of searching for new diagnostic methods of DCh wear—to increase the HT lifetime and to reduce the test duration [1–3].

Performance of the mentioned works on increase in a resource of the HT can be done in several directions:


The search for new materials is a direction of work that requires not only significant material but also durable time expenses; positive result will undoubtedly lead to DCh recourse increase, but will not be able to influence on characteristics of finished thrusters.

and also neutralizes the charge of the plume flowing out of the thruster. The anode block consists of a magnetic system and accelerating channel (2) with an anode (3). The magnetic system (MS) includes: magnetic part (4), inner core (5) and an outer core (6), an outer pole (7) and an internal pole (8), as well as a coils (9 and 10). The MS is constructed in such a way that a radial magnetic field is formed in the discharge channel (DCh) located in the gap between the magnetic poles. DCh is limited by coaxial cylindrical walls of dielectric material, in the base of which anode is located, which also performs the function of supplying gas to the thruster channel [4].

Principle scheme of the HT: 1—cathode-neutralizer; 2—discharge chamber; 3—anode; 4—magnetic part; 5 and 6—inner and outer cores; 7 and 8—outer and inner poles; 9 and 10—inner and outer coils.

The thruster works as follows: the propellant through the anode comes via the DСh. Discharge voltage is applied between the cathode and the anode. Under the action of electric field, electrons begin to move from the cathode to the anode and enter the crossed electric and magnetic fields. The magnitude of the magnetic field induction in the DCh is chosen such that the Larmor radii of the electrons re and the

ions ri satisfy the condition re ≪ L ≪ ri, where L is the length of the DCh.

the plasma oscillations.

93

Figure 1.

Hall Thruster Erosion

DOI: http://dx.doi.org/10.5772/intechopen.82654

ized by electron flow created by the cathode.

2.2 Discharge chamber lifetime

Under these conditions, the electron motion is performed in the azimuthal direction, and in the direction of the anode, displacement occurs due to collisions with neutral and charged particles, with the walls of the DCh, and also because of

During the operation of the thruster, the electrons moving to the anode ionize the atoms of the propellant, and the ions that are formed are accelerated along the electric field, creating a reactive thrust. At the thruster exit, the ion flow is neutral-

In the region of the DCh exit, an ion stream is produced, and some of which is directed not along the axis of the thruster, but on the wall of the DCh. In the plasma

The existing mathematical models of HT allow us to simulate the process of DCh insulator wear and based on the results of calculations, with reasonable accuracy, to predict the behavior of HT erosion. Despite this, leading HT developers are inclined to believe that trusted quantitative data can only be obtained experimentally on the basis of long-term resource tests.

The results of the analysis of the state of wear research of the DCh HT show that today there is no possibility to make a list of the main guaranteed recommendations according to which it would be possible to develop a design of the thruster with a lifetime up to two times higher than existing analogues; therefore, HT improvement works are carried out in all of the previously mentioned areas.

This chapter presents the results of research of the HT DCh edge wear at various operating modes (with different ratios of coil currents at stable values of the discharge voltage and mass flow rate).

The research results of low-power HT with different materials of ceramic insulators are presented.

#### 2. Possible Hall thruster lifetime limiting elements

#### 2.1 Main components and operating principle of the HT

HT is a plasma thruster in which thrust is created by propellant ions, formed and accelerated in a discharge in crossed electric and magnetic fields.

In the traditional version, HT consists of two main nodes—the cathode and anode blocks (Figure 1). The cathode (1) is performed according to the scheme of a hollow cathode, located on the side of the anode block and is a source of electrons. The purpose of the cathode is to ensure the thruster ignition, maintains its normal operation,

Hall Thruster Erosion DOI: http://dx.doi.org/10.5772/intechopen.82654

#### Figure 1.

from individual samples to a whole range of thrusters, even with one design, is unreasonable because of the lack of statistical data and high requirements for space

such tests can be compared to the cost of one HT.

• search for new ceramic materials;

• improvement of thruster design;

selection of its operation mode.

mentally on the basis of long-term resource tests.

2. Possible Hall thruster lifetime limiting elements

2.1 Main components and operating principle of the HT

accelerated in a discharge in crossed electric and magnetic fields.

charge voltage and mass flow rate).

Second, while using the reduced lifetime tests, the duration of the experiment is about 10% of the physical thruster lifetime. For example, for HT with 4000 hours of operation time, a test with a minimum base of 400 hours is required. The cost of

These factors determine the necessity of searching for new diagnostic methods of DCh wear—to increase the HT lifetime and to reduce the test duration [1–3]. Performance of the mentioned works on increase in a resource of the HT can be

• improvement of mathematical model of thruster in order to find new ways of

• search for new ways of HT resource increasing, for instance, by means of

The search for new materials is a direction of work that requires not only significant material but also durable time expenses; positive result will undoubtedly lead to DCh recourse increase, but will not be able to influence on characteristics

The existing mathematical models of HT allow us to simulate the process of DCh insulator wear and based on the results of calculations, with reasonable accuracy, to predict the behavior of HT erosion. Despite this, leading HT developers are inclined to believe that trusted quantitative data can only be obtained experi-

The results of the analysis of the state of wear research of the DCh HT show that today there is no possibility to make a list of the main guaranteed recommendations according to which it would be possible to develop a design of the thruster with a lifetime up to two times higher than existing analogues; therefore, HT improvement works are carried out in all of the previously mentioned areas.

This chapter presents the results of research of the HT DCh edge wear at various operating modes (with different ratios of coil currents at stable values of the dis-

The research results of low-power HT with different materials of ceramic insu-

HT is a plasma thruster in which thrust is created by propellant ions, formed and

In the traditional version, HT consists of two main nodes—the cathode and anode blocks (Figure 1). The cathode (1) is performed according to the scheme of a hollow cathode, located on the side of the anode block and is a source of electrons. The purpose of the cathode is to ensure the thruster ignition, maintains its normal operation,

enhancing sustainability of the DCh to high-energy ion flow; and

technology.

Propulsion Systems

done in several directions:

of finished thrusters.

lators are presented.

92

Principle scheme of the HT: 1—cathode-neutralizer; 2—discharge chamber; 3—anode; 4—magnetic part; 5 and 6—inner and outer cores; 7 and 8—outer and inner poles; 9 and 10—inner and outer coils.

and also neutralizes the charge of the plume flowing out of the thruster. The anode block consists of a magnetic system and accelerating channel (2) with an anode (3). The magnetic system (MS) includes: magnetic part (4), inner core (5) and an outer core (6), an outer pole (7) and an internal pole (8), as well as a coils (9 and 10). The MS is constructed in such a way that a radial magnetic field is formed in the discharge channel (DCh) located in the gap between the magnetic poles. DCh is limited by coaxial cylindrical walls of dielectric material, in the base of which anode is located, which also performs the function of supplying gas to the thruster channel [4].

The thruster works as follows: the propellant through the anode comes via the DСh. Discharge voltage is applied between the cathode and the anode. Under the action of electric field, electrons begin to move from the cathode to the anode and enter the crossed electric and magnetic fields. The magnitude of the magnetic field induction in the DCh is chosen such that the Larmor radii of the electrons re and the ions ri satisfy the condition re ≪ L ≪ ri, where L is the length of the DCh.

Under these conditions, the electron motion is performed in the azimuthal direction, and in the direction of the anode, displacement occurs due to collisions with neutral and charged particles, with the walls of the DCh, and also because of the plasma oscillations.

During the operation of the thruster, the electrons moving to the anode ionize the atoms of the propellant, and the ions that are formed are accelerated along the electric field, creating a reactive thrust. At the thruster exit, the ion flow is neutralized by electron flow created by the cathode.

#### 2.2 Discharge chamber lifetime

In the region of the DCh exit, an ion stream is produced, and some of which is directed not along the axis of the thruster, but on the wall of the DCh. In the plasma of the HT, the ion flux reaches an energy level of tens, hundreds, and thousands of electron volts. Upon impact against the surface, the ions begin to "dissipate" it, that is, knock out one or more atoms (physical sputtering—erosion), ions (emission), electrons (emission), or fragments of molecules (chemical sputtering). Bombarding ions can acquire electrons on the surface and be reflected from it in the form of neutral atoms, the neutralization process. Ions can be bound to the sample surface (adsorbed). The listed processes are accompanied by electromagnetic radiation from ultraviolet to X-ray ranges.

The ion energy greatly exceeds the covalent bond energy of the particles in the layer and the weak Van der Waals and electrostatic interaction between the layers, which leads to the appearance of various types of physical sputtering, which are divided into several types: primary direct knockout, sputtering by linear cascades, and thermal peaks [5].

In the first case, the ion transmits to atoms enough amount of energy, so that atoms are able to leave the material after a small number of collisions. In two other sputtering modes, the amount of transmitted energy is sufficient to excite the displacement from the equilibrium position of whole groups of atoms that overcome the surface energy barrier and leave the wall of the DCh. This can explain the occurrence of zones of normal and abnormal erosion on the surface of ceramic inserts.

The width of the erosion belt depends not only on the thruster design, the field distribution, but also on the DCh material and the choice of the thruster ope-

Erosion belt of the thruster of 0.1 kW after 92 hours of operation: 1—abnormal erosion; 2—normal erosion.

Resistance to the wear of the DCh wall material depends on more than 20 parameters, the most important of which are composition, density, porosity, and

In spite of the fact that the cathode neutralizer is located in the zone with the minimal influence of the ion flow (i.e., outside the thruster plume), the fraction of ions with high-energy values is present in the near-cathode region. This leads to a

The most closely located to the DCh section element of the cathode is the ignition electrode (IE) (Figure 6). The goal of IE is to create an initial discharge between the emitter and the IE. Next, the main discharge is ignited between the

Analysis of the results of long-term lifetime tests carried out in the EDB Fakel Kaliningrad [9] showed that the volume of IE material sputtering is linear until the

2.3 The cathode lifetime (ignition electrode and emitter)

gradual sputtering of the cathode block (Figure 4).

ration mode.

Hall Thruster Erosion

DOI: http://dx.doi.org/10.5772/intechopen.82654

Figure 3.

manufacturing technology [7].

anode and cathode blocks.

Figure 4.

95

Position of cathode according to the DCh exit.

The main mass of sputtering product of the DCh material is neutral atoms, and the fraction of ions does not exceed 1% [6]. Ionized erosion products are recorded only at high values of the discharge voltage; the number of emitted ions of the material is by 2–4 orders of magnitude less than the number of emitted atoms.

The result of the erosion process can be traced visually after the tests. Figure 2 is the photo of new 0.1 kW power hall thruster that is developed by the Scientific-Technological Center of Space Power and Engines (STC SPE) of National Aerospace University named after N. Ye. Zhukovsky "Kharkiv Aviation Institute," Ukraine. Figure 3 presents the ceramic wall of the HT 0.1 kW after 92 hours of operation. There are areas of normal and abnormal erosion.

#### Figure 2.

Photo of the new hall thruster of 0.1 kW power and cathode: 1—cathode-neutralizer; 2—magnetic pole; 3—outer ceramic; 4–inner poles; 5–inner ceramic.

of the HT, the ion flux reaches an energy level of tens, hundreds, and thousands of electron volts. Upon impact against the surface, the ions begin to "dissipate" it, that is, knock out one or more atoms (physical sputtering—erosion), ions (emission), electrons (emission), or fragments of molecules (chemical sputtering). Bombarding ions can acquire electrons on the surface and be reflected from it in the form of neutral atoms, the neutralization process. Ions can be bound to the sample surface (adsorbed). The listed processes are accompanied by electromagnetic radiation

The ion energy greatly exceeds the covalent bond energy of the particles in the layer and the weak Van der Waals and electrostatic interaction between the layers, which leads to the appearance of various types of physical sputtering, which are divided into several types: primary direct knockout, sputtering by linear cascades,

In the first case, the ion transmits to atoms enough amount of energy, so that atoms are able to leave the material after a small number of collisions. In two other sputtering modes, the amount of transmitted energy is sufficient to excite the displacement from the equilibrium position of whole groups of atoms that overcome the surface energy barrier and leave the wall of the DCh. This can explain the occurrence

The main mass of sputtering product of the DCh material is neutral atoms, and the fraction of ions does not exceed 1% [6]. Ionized erosion products are recorded only at high values of the discharge voltage; the number of emitted ions of the material is by 2–4 orders of magnitude less than the number of emitted atoms.

The result of the erosion process can be traced visually after the tests. Figure 2 is the photo of new 0.1 kW power hall thruster that is developed by the Scientific-Technological Center of Space Power and Engines (STC SPE) of National Aerospace University named after N. Ye. Zhukovsky "Kharkiv Aviation Institute," Ukraine. Figure 3 presents the ceramic wall of the HT 0.1 kW after 92 hours of operation.

Photo of the new hall thruster of 0.1 kW power and cathode: 1—cathode-neutralizer; 2—magnetic pole;

of zones of normal and abnormal erosion on the surface of ceramic inserts.

There are areas of normal and abnormal erosion.

from ultraviolet to X-ray ranges.

and thermal peaks [5].

Propulsion Systems

Figure 2.

94

3—outer ceramic; 4–inner poles; 5–inner ceramic.

Figure 3. Erosion belt of the thruster of 0.1 kW after 92 hours of operation: 1—abnormal erosion; 2—normal erosion.

The width of the erosion belt depends not only on the thruster design, the field distribution, but also on the DCh material and the choice of the thruster operation mode.

Resistance to the wear of the DCh wall material depends on more than 20 parameters, the most important of which are composition, density, porosity, and manufacturing technology [7].

#### 2.3 The cathode lifetime (ignition electrode and emitter)

In spite of the fact that the cathode neutralizer is located in the zone with the minimal influence of the ion flow (i.e., outside the thruster plume), the fraction of ions with high-energy values is present in the near-cathode region. This leads to a gradual sputtering of the cathode block (Figure 4).

The most closely located to the DCh section element of the cathode is the ignition electrode (IE) (Figure 6). The goal of IE is to create an initial discharge between the emitter and the IE. Next, the main discharge is ignited between the anode and cathode blocks.

Analysis of the results of long-term lifetime tests carried out in the EDB Fakel Kaliningrad [9] showed that the volume of IE material sputtering is linear until the

Figure 4. Position of cathode according to the DCh exit.

#### Propulsion Systems

length of the outer insulator exceeds the length of the outer pole tip. As the length of the outer ceramic insert decreases, the insulator ceases to be a barrier between the plasma plume and the cathode block, and thereby opens both the pole tip and the cathode for ion flow influence. The results of long tests of about 3000 hours or more showed that even with full wear of the end part of the IE, it fully performs the function of starting the thruster. According to the forecast, the IE lifetime is more than 10,000 h.

Using direct methods, data on the rate of erosion of ceramic units of measurement—mm/h or mg/h are received. The error in determining the wear depends on the stability of the thruster operation and the accuracy of the measuring equipment.

In order to be able to monitor the erosion rate directly during thruster operation,

The principle of the method is to trap the sputtering products of ceramic insu-

The design of the recording device is designed in such a way that the ionic component of the plume would not be able to overcome the potential barrier created by the electric field applied to the housing. Permanent magnets are used to trap electron component of the flow. Magnets characteristics are selected depending on the thruster plume. The quartz crystal is attached to the resonating circuit. During the experiment, the frequency of the loop oscillation is monitored, which changes when the film forms on the surface of the crystal from the deposited neutral atoms—the erosion products of the thruster

Analysis of the data obtained using this method showed good alignment of

Laser-induced fluorescence (LIF) has also been used for several decades to study

LIF is considered a highly sensitive method [8], since the excitation rate is independent of the plasma parameters, with stable operation of the thruster, and there is a possibility to adjust the equipment to obtain maximum sensitivity for each

results and the method of optical emission spectroscopy (OES).

The device for recording the mass loss of the DCh insulators by the quartz crystal method.

there are a number of indirect measurement methods. These include optical

There is no need for a sophisticated calculation method.

3.2 Indirect methods

Hall Thruster Erosion

methods and special methods [10].

DOI: http://dx.doi.org/10.5772/intechopen.82654

lators on a quartz crystal (Figure 5).

3.2.1 Method of quartz crystal

3.2.2 Laser-induced fluorescence

the fluorescence spectra of HT plasma [9].

material.

Figure 5.

97

The cathode emitter lifetime is determined by the rate of material loss of its emission part. The statistical data show that during the time of the cathode block lifetime tests with duration of more than 8000 hours, the emitter volume loss is less than 8%. The predicted emitter lifetime is not less than 15,000 hours [15].

#### 2.4 Hall thruster magnetic system lifetime

HT magnetic system materials possess low sputtering coefficients as compared to ceramic insulators of DCh. The design of discharge chamber exit of the thruster is designed in such a way that the edges of both the outer ceramic insert and the inner ceramic insert block the ion flow to the magnetic poles.

Obviously, despite the applied constructional solutions, there is a weak flux of high-energy ions in this region. However, the material mass leakage is negligible. The results of experiments on the determination of the initial stage of materials of the magnetic system wear at different thicknesses of the walls of the DCh show that with a proper selection of the geometry of the DCh walls, a lifetime of a magnetic system of more than 15,000 hours can be achieved for a 1.35 kW HT model [8].

The results of long-term HT lifetime tests show that at first, DCh edges are susceptible to ion bombardment discharged, and only after the total sputtering of the edges of the insulators, magnetic system elements and cathode block are beginning to significantly erode. Hence, one of the most critical elements of HT from the point of view of the lifetime is the edge of the DCh insulators.

#### 3. Overview of methods of Hall thruster erosion measurements

A number of direct and indirect methods have been developed to study the rate of erosion of the HT DCh, each of which has unique methodological differences.

#### 3.1 Direct methods

During the thruster operation, the profile of the edges gradually changes due to the loss of the weight of the DCh insulators. If the design of the thruster allows rapid extraction of ceramic inserts from the HT anodic block quickly and without causing defects, then the mass change control (weighting methods) is carried out, for example, using laboratory scales (WA-21).

For the initial masses of outer and inner ceramics mOC\_0 and mIC\_0, the mass of the insulator is taken just after its manufacturing. Further tests are carried out with the HT unchanged operating mode. The time base of the experiment is divided into several stages, for example, 10–50 hours each, depending on the thruster power and the rate of erosion. After the end of each stage of the tests, the insulators are removed from the anode block, and their mass, mn1 and mv1, are measured. Mass loss rate, mg/h, of the sample is determined by calculating the difference between the initial and measured masses and dividing it into a time test base [9].

In order to control the radial erosion, mm/h, measuring instrumental microscopes are used.

#### Hall Thruster Erosion DOI: http://dx.doi.org/10.5772/intechopen.82654

Using direct methods, data on the rate of erosion of ceramic units of measurement—mm/h or mg/h are received. The error in determining the wear depends on the stability of the thruster operation and the accuracy of the measuring equipment. There is no need for a sophisticated calculation method.

#### 3.2 Indirect methods

length of the outer insulator exceeds the length of the outer pole tip. As the length of the outer ceramic insert decreases, the insulator ceases to be a barrier between the plasma plume and the cathode block, and thereby opens both the pole tip and the cathode for ion flow influence. The results of long tests of about 3000 hours or more showed that even with full wear of the end part of the IE, it fully performs the function of starting the thruster. According to the forecast, the IE lifetime is more

The cathode emitter lifetime is determined by the rate of material loss of its emission part. The statistical data show that during the time of the cathode block lifetime tests with duration of more than 8000 hours, the emitter volume loss is less

HT magnetic system materials possess low sputtering coefficients as compared to ceramic insulators of DCh. The design of discharge chamber exit of the thruster is designed in such a way that the edges of both the outer ceramic insert and the inner

Obviously, despite the applied constructional solutions, there is a weak flux of high-energy ions in this region. However, the material mass leakage is negligible. The results of experiments on the determination of the initial stage of materials of the magnetic system wear at different thicknesses of the walls of the DCh show that with a proper selection of the geometry of the DCh walls, a lifetime of a magnetic system of more than 15,000 hours can be achieved for a 1.35 kW HT model [8]. The results of long-term HT lifetime tests show that at first, DCh edges are susceptible to ion bombardment discharged, and only after the total sputtering of the edges of the insulators, magnetic system elements and cathode block are beginning to significantly erode. Hence, one of the most critical elements of HT from the

than 8%. The predicted emitter lifetime is not less than 15,000 hours [15].

2.4 Hall thruster magnetic system lifetime

ceramic insert block the ion flow to the magnetic poles.

point of view of the lifetime is the edge of the DCh insulators.

for example, using laboratory scales (WA-21).

3. Overview of methods of Hall thruster erosion measurements

A number of direct and indirect methods have been developed to study the rate of erosion of the HT DCh, each of which has unique methodological differences.

During the thruster operation, the profile of the edges gradually changes due to the loss of the weight of the DCh insulators. If the design of the thruster allows rapid extraction of ceramic inserts from the HT anodic block quickly and without causing defects, then the mass change control (weighting methods) is carried out,

For the initial masses of outer and inner ceramics mOC\_0 and mIC\_0, the mass of the insulator is taken just after its manufacturing. Further tests are carried out with the HT unchanged operating mode. The time base of the experiment is divided into several stages, for example, 10–50 hours each, depending on the thruster power and the rate of erosion. After the end of each stage of the tests, the insulators are removed from the anode block, and their mass, mn1 and mv1, are measured. Mass loss rate, mg/h, of the sample is determined by calculating the difference between

In order to control the radial erosion, mm/h, measuring instrumental micro-

the initial and measured masses and dividing it into a time test base [9].

than 10,000 h.

Propulsion Systems

3.1 Direct methods

scopes are used.

96

In order to be able to monitor the erosion rate directly during thruster operation, there are a number of indirect measurement methods. These include optical methods and special methods [10].

#### 3.2.1 Method of quartz crystal

The principle of the method is to trap the sputtering products of ceramic insulators on a quartz crystal (Figure 5).

The design of the recording device is designed in such a way that the ionic component of the plume would not be able to overcome the potential barrier created by the electric field applied to the housing. Permanent magnets are used to trap electron component of the flow. Magnets characteristics are selected depending on the thruster plume. The quartz crystal is attached to the resonating circuit. During the experiment, the frequency of the loop oscillation is monitored, which changes when the film forms on the surface of the crystal from the deposited neutral atoms—the erosion products of the thruster material.

Analysis of the data obtained using this method showed good alignment of results and the method of optical emission spectroscopy (OES).

#### 3.2.2 Laser-induced fluorescence

Laser-induced fluorescence (LIF) has also been used for several decades to study the fluorescence spectra of HT plasma [9].

LIF is considered a highly sensitive method [8], since the excitation rate is independent of the plasma parameters, with stable operation of the thruster, and there is a possibility to adjust the equipment to obtain maximum sensitivity for each

Figure 5. The device for recording the mass loss of the DCh insulators by the quartz crystal method.

component under study. However, the need to install and set a large number of measuring equipment, the need to reconfigure the laser to investigate various components of the spectrum, critically limits the scope of the method [9].

4. Detailed description of the optical emission spectroscopy

4.1 Basic scheme and principle of measurements with the OES method

collimator (OESSC)

DOI: http://dx.doi.org/10.5772/intechopen.82654

Hall Thruster Erosion

wavelengths.

ceramic.

was 4.5°.

Figure 6.

99

The scheme of measurements of the rate of erosion by the OES method.

(OES) method. OES development: method of the optical emission spectroscopy with the scanning of thruster plasma through

The basic scheme of the OES method is shown in Figure 6. Radiation of the plasma of the thruster through the inspection flange and the lens by means of an optical cable is transferred to a spectrometer where the spectrum is decomposed. Further, the signal is transmitted to the computer and output as a dependence of the intensity of the spectral line on the wavelength. As a rule, the inspection flange of a vacuum chamber is made of quartz glass of KU-1 type with a low absorption coefficient of radiation in a wide range of

It is obvious from Figure 8 that when using the basic scheme of the OES method, it is not possible to register radiation separately from each edge of the DCh

4.2 Basic scheme and principle of measurements with the OESSC method

For the spectral measurements from small regions of the plasma radiation of the thruster, the OES method experimental scheme was modified. From the OES measurement scheme, the collecting lens was removed, and the optical cable was installed into the vacuum chamber at the minimum possible distance from the DCh cut. A collimator was installed on the fiber optic cable (Figure 7). Collimator is a device that geometrically reduces the divergence angle of the optical fiber. For a standard optical cable, the divergence angle is equal to 25.4°. For the developed optical receiver, which consists of a collimator and a cable, the divergence angle

#### 3.2.3 Mass spectrometry

One of the methods for recording HT sputtering products is the mass spectrometry method [8]. Modern mass spectrometers have high resolution and can be used with all kinds of particles in different states of excitation and different concentrations of the sample under study.

Despite the significant advantages of modern mass spectrometers, high sensitivity, in comparison with other optical methods, unambiguous identification of particles, and a large operating range of masses, they have a number of disadvantages. First, mass spectrometers are complex instruments both in operation and in maintenance. Second, modern mass spectrometers are an extremely expensive type of equipment, the cost of which is comparable to the cost of manufacturing several HTs.

#### 3.2.4 Optical emission spectroscopy

The principle of the method of optical emission spectroscopy consists in recording the emission spectrum of the thruster sputtering products and recalculating the intensities of the spectral lines in the erosion rate [9]. Advantages of the method, namely contactless, relatively simple procedure for processing experimental data, and relatively simple technical organization of measuring equipment, led to the extensive use of OES as one of the main diagnostic tools in the field of research of the technical state of HT.

The requirements to the method of HT insulators edges erosion rate are next:


Methodological requirements are as follows: relatively high measurement accuracy and unambiguous interpretation of the results.

The analysis of the methods for diagnosing the rate of erosion of the spraying products of the edge edges of HT insulators showed that in order to obtain qualitative measurement results, two diagnostic methods should be used throughout the thruster test cycle: OES and measurements by direct methods.

component under study. However, the need to install and set a large number of measuring equipment, the need to reconfigure the laser to investigate various com-

One of the methods for recording HT sputtering products is the mass spectrometry method [8]. Modern mass spectrometers have high resolution and can be used with all kinds of particles in different states of excitation and different concentra-

Despite the significant advantages of modern mass spectrometers, high sensitivity, in comparison with other optical methods, unambiguous identification of particles, and a large operating range of masses, they have a number of disadvantages. First, mass spectrometers are complex instruments both in operation and in maintenance. Second, modern mass spectrometers are an extremely expensive type of equipment, the cost of which is comparable to the cost of manufacturing

The principle of the method of optical emission spectroscopy consists in recording the emission spectrum of the thruster sputtering products and recalculating the intensities of the spectral lines in the erosion rate [9]. Advantages of the method, namely contactless, relatively simple procedure for processing experimental data, and relatively simple technical organization of measuring equipment, led to the extensive use of OES as one of the main diagnostic tools in the field of research of

The requirements to the method of HT insulators edges erosion rate are next:

• no need for multiple calibration and adjustment of measuring equipment;

• measurement of the external and internal insulators erosion rate separately.

Methodological requirements are as follows: relatively high measurement accu-

The analysis of the methods for diagnosing the rate of erosion of the spraying products of the edge edges of HT insulators showed that in order to obtain qualitative measurement results, two diagnostic methods should be used throughout the

• short-term preparation time of measuring equipment;

• the ability to do measurements during the thruster operation;

• short-term procedure of experimental data processing;

thruster test cycle: OES and measurements by direct methods.

• possibility of multiple use of recording sensors;

racy and unambiguous interpretation of the results.

• non-intrusive, estimation of erosion without any contact with thruster itself or

ponents of the spectrum, critically limits the scope of the method [9].

3.2.3 Mass spectrometry

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several HTs.

tions of the sample under study.

3.2.4 Optical emission spectroscopy

the technical state of HT.

plasma plume;

• technically simple;

98

#### 4. Detailed description of the optical emission spectroscopy (OES) method. OES development: method of the optical emission spectroscopy with the scanning of thruster plasma through collimator (OESSC)

#### 4.1 Basic scheme and principle of measurements with the OES method

The basic scheme of the OES method is shown in Figure 6. Radiation of the plasma of the thruster through the inspection flange and the lens by means of an optical cable is transferred to a spectrometer where the spectrum is decomposed. Further, the signal is transmitted to the computer and output as a dependence of the intensity of the spectral line on the wavelength. As a rule, the inspection flange of a vacuum chamber is made of quartz glass of KU-1 type with a low absorption coefficient of radiation in a wide range of wavelengths.

It is obvious from Figure 8 that when using the basic scheme of the OES method, it is not possible to register radiation separately from each edge of the DCh ceramic.

#### 4.2 Basic scheme and principle of measurements with the OESSC method

For the spectral measurements from small regions of the plasma radiation of the thruster, the OES method experimental scheme was modified. From the OES measurement scheme, the collecting lens was removed, and the optical cable was installed into the vacuum chamber at the minimum possible distance from the DCh cut. A collimator was installed on the fiber optic cable (Figure 7). Collimator is a device that geometrically reduces the divergence angle of the optical fiber. For a standard optical cable, the divergence angle is equal to 25.4°. For the developed optical receiver, which consists of a collimator and a cable, the divergence angle was 4.5°.

Figure 6. The scheme of measurements of the rate of erosion by the OES method.

thruster erosion. However, this problem was solved by installing of the quartz glass

The results of spectral measurements of two thrusters with different materials of ceramic insulators BN and BN + SiO2 are presented in Figure 8. As can be seen in the UV-range, boron lines with a wavelength of 249.66 and 249.77 nm are presented on the spectra of both thrusters. These boron lines that are suitable for qualitative analysis are well recorded. Silicon lines are presented on the thruster with BN + SiO2 ceramic and absent on the thruster with BN ceramic. The intensities of silicon line in UV-range are relatively high, so these lines can be used for the quantitative

Known method of erosion rate computation based on the thruster DCh radiation spectrum measurements has a comparable character. It means that, for example, computational data can give assessment of erosion rate differ at different thruster operational modes or how the erosion rate changes in time. OES equation does not describe how the optical fiber position and DCh erosion area influence the registered emission flow. These data are necessary for the computation of erosion rate

> R<sup>В</sup> 2 V<sup>В</sup> I Xe 828:01 I Xeþ<sup>m</sup> <sup>484</sup>:<sup>43</sup>

(1)

249.77 is the lighting of the optical

484.43 is the lighting

That is why for the OESSC method, next calculation expression is used

of the optical cable by the xenon metastable ions on the wavelength 484.43 nm;

828.01 is the lighting of the optical cable by the xenon atoms on the wavelength 828.01 nm; RB is the distance from the source to the radiation detector, m.; and VB is

This expression is used for defining the erosion rate from different parts of

Spectrum in the UV-range 240–270 nm: black—HT with BN insulator; blue—HT with BN + SiO2 insulator.

.

Er ¼ I B 249:77

cable by the boron atoms on the wavelength 249.77 nm; IXe + m

where Er is the erosion rate of the material; IB

the amount of radiation from DCh wear products, m<sup>3</sup>

protective screens to the sensitive element of the optical fiber [12].

4.3 Processing of the spectral measurement results

DOI: http://dx.doi.org/10.5772/intechopen.82654

analysis.

Hall Thruster Erosion

I Xe

the DCh.

Figure 8.

101

from different DCh walls.

#### Figure 7.

Scheme of measurements of the erosion rate by the OESSC method: 1—position of measurements of "inner ceramic"; and 2—position of measurements of "inner ceramic" and "outer ceramic."

It is also possible to scan radiation from different regions of the DCh by installing a registration element on a coordinate platform. According to the presented scheme, the improved method of OES was called the method of optical emission spectroscopy with scanning of the plasma radiation of the thruster through a collimator (OESSC).

The above scheme of measurements by the OESSC method has a number of advantages:


One of the method disadvantages is the installation of a sensitive optical cable in a vacuum chamber, which leads to the dusting of the fiber with the products of

thruster erosion. However, this problem was solved by installing of the quartz glass protective screens to the sensitive element of the optical fiber [12].

#### 4.3 Processing of the spectral measurement results

The results of spectral measurements of two thrusters with different materials of ceramic insulators BN and BN + SiO2 are presented in Figure 8. As can be seen in the UV-range, boron lines with a wavelength of 249.66 and 249.77 nm are presented on the spectra of both thrusters. These boron lines that are suitable for qualitative analysis are well recorded. Silicon lines are presented on the thruster with BN + SiO2 ceramic and absent on the thruster with BN ceramic. The intensities of silicon line in UV-range are relatively high, so these lines can be used for the quantitative analysis.

Known method of erosion rate computation based on the thruster DCh radiation spectrum measurements has a comparable character. It means that, for example, computational data can give assessment of erosion rate differ at different thruster operational modes or how the erosion rate changes in time. OES equation does not describe how the optical fiber position and DCh erosion area influence the registered emission flow. These data are necessary for the computation of erosion rate from different DCh walls.

That is why for the OESSC method, next calculation expression is used

$$\mathbf{Er} = \mathbf{I}\_{249.77}^{\mathbf{B}} \frac{\mathbf{R}\_{\mathbf{B}}^{\mathbf{2}}}{\mathbf{V}\_{\mathbf{B}}} \frac{\mathbf{I}\_{828.01}^{\mathbf{Xe}}}{\mathbf{I}\_{484.43}^{\mathbf{Xe}+\mathbf{m}}} \tag{1}$$

where Er is the erosion rate of the material; IB 249.77 is the lighting of the optical cable by the boron atoms on the wavelength 249.77 nm; IXe + m 484.43 is the lighting of the optical cable by the xenon metastable ions on the wavelength 484.43 nm; I Xe 828.01 is the lighting of the optical cable by the xenon atoms on the wavelength 828.01 nm; RB is the distance from the source to the radiation detector, m.; and VB is the amount of radiation from DCh wear products, m<sup>3</sup> .

This expression is used for defining the erosion rate from different parts of the DCh.

Figure 8. Spectrum in the UV-range 240–270 nm: black—HT with BN insulator; blue—HT with BN + SiO2 insulator.

It is also possible to scan radiation from different regions of the DCh by installing a registration element on a coordinate platform. According to the presented scheme, the improved method of OES was called the method of optical emission spectroscopy with scanning of the plasma radiation of the thruster through

Scheme of measurements of the erosion rate by the OESSC method: 1—position of measurements of "inner

The above scheme of measurements by the OESSC method has a number of

• the possibility to obtain data on the wear rate of insulators from each of the

• the ability to carry out measurements of the entrainment of the material separately of different parts of the thruster magnetic system during the tests;

• possibility to register the degree of wear of ceramics in the azimuth

• nonintrusive—no contact of measuring equipment with thruster plasma, which may lead to change thruster operation mode and the change of the

• absence of complicated procedures for calibration and adjustment of

at different voltages, mass flow rates, and coil currents [11].

• the possibility to get erosion data at different thruster operating modes, namely

One of the method disadvantages is the installation of a sensitive optical cable in a vacuum chamber, which leads to the dusting of the fiber with the products of

ceramic inserts separately during the tests of the thruster;

ceramic"; and 2—position of measurements of "inner ceramic" and "outer ceramic."

• a simple procedure of experimental data processing;

a collimator (OESSC).

direction;

100

plasma parameters;

measuring equipment; and

advantages:

Figure 7.

Propulsion Systems

#### 5. OES studies of plasma propulsion systems of 100 W power

#### 5.1 OES test equipment and its adjustment

For the purpose of preliminary studies on the influence of the modes of operation of the HT on its erosion characteristics for measuring the spectra of optical radiation of the plasma plume of the 0.1 kW HT, an experimental installation was established, and the scheme of which is shown in Figure 9. The radiation emitted by the plasma (1) 0.1 kW HT, through a quartz window (2) (diameter—40 mm and thickness—5 mm), was observed with the help of a device for measuring the spectrum located outside the camera. The angle of observation of the plasma plume at the cut of the thruster was 60° relative to the chamber axis. With 10 cm of focusing lens (3), the radiation is collected and focused on a light conductive fiber (4) with a diameter of 1 mm. The distance between the plasma plume and the lens was about 80 cm. The end of the fiber is connected by a special adapter with the inlet of the spectrometer (5). A lens with a focal length of 10 cm gave an approximately 9 magnification. With an optical fiber with a diameter of 1 mm, the spatial resolution was about 9 mm in diameter. Thus, the installation captured most of the plasma jet in the thruster channel and its exit.

through the computer, and the graphical interface allows you to get a spectral image

Such a measuring spectrometric system is quite user-friendly, promptly provides information about the controlled spectrum, and thanks to a numerical representation, which allows us to conveniently store information and compare it with

Together with the Laboratoire de Physique des Gaz et des Plasmas, Université

Paris, Laboratoire d'Analyse Spectroscopique et d'Energétique des Plasmas, Université d'Orléans, and Laboratoire d'Aérothermique, the 0.1 W HT optical (spectral) tests were conducted to determine the effect of the regime of the thruster on its erosion characteristics [13]. Measurements were provided on the test facility as shown in Figure 10. Experimental parameters, discharge voltage (Ud) from 200 to 300 V in a step of 20 V, and also the current on the solenoid of the magnetic system of the thruster varied in the range of 1.5–3.0 A. All measurements were

In calculating of this parameter, the power of the solenoid was taken into account, but was neglected by the mass flow of xenon through the cathode. The maximum value of the efficiency (0.353) of HT during this experiment was obtained at a discharge voltage of 300 V and a mass flow of xenon 0.25 mg/s. Xe I (828.01 nm), Xe II (484.43 nm), and Al I (396.152 nm) lines were distinguished for the analysis of the level of erosion of the stationary plasma thruster with ceramics from the ABN. These three lines made it possible to analyze the rate of erosion. This analysis was conducted with Laboratoire d'Analyze Spectroscopique et d'Energétique des Plasmas, Université d'Orléans, and its results are published in the

The Xe II xenon ion emission occurs due to the transition from the wavelength

7=2

) to the level 11.54 eV (68045.156 cm<sup>1</sup>

<sup>6</sup><sup>s</sup> <sup>5</sup>p5 2P<sup>0</sup>

3=2

D7=<sup>2</sup> and 5d<sup>4</sup>

<sup>6</sup><sup>p</sup> of 14.10 eV

). The upper level

F7<sup>=</sup>2.

5.2 OES studies of the HT erosion at different operation modes

made at two values of xenon consumption: 0.20 and 0.25 mg/s.

materials of the international conference [14].

<sup>6</sup><sup>s</sup> is basically two metastable states Xe<sup>+</sup> <sup>5</sup>d<sup>4</sup>

of 484.43 nm from the initial level 5p<sup>4</sup> D<sup>0</sup>

Dependence of the erosion from the discharge voltage.

(80118.962 cm<sup>1</sup>

5p<sup>4</sup> D<sup>0</sup> 7=2

Figure 10.

103

that can be saved as a text file with data.

DOI: http://dx.doi.org/10.5772/intechopen.82654

Hall Thruster Erosion

another that was previously registered.

The end of the optical cable through a special adapter connects to the input port of the spectrometer USB2000. This mini spectrometer, manufactured by Ocean Optics (USA), was kindly provided by Laboratoire de Physique des Gaz et des Plasmas [13]. It was used to cover the considered range with a resolution that is sufficient to distinguish optical lines. Moreover, its weight (600 g), dimensions (15 11 5 cm), and power consumption (100 mA) make it an excellent candidate for the use in summer experiments to study the erosion of ceramics and the optical spectrum of the plasma plume of the thruster (Figure 9).

The light entering the spectrometer is refracted by a fixed dispersing prism and is directed to the charge coupled device (CCD) line with 2048 elements. So, in one shot, this spectrometer gives a spectrum of wavelengths of 380–830 nm with a resolution of 0.2–0.22 nm. The spectrometer is connected to the laptop through the USB interface. The computer controls the parameters of the spectrometer USB2000

#### Figure 9.

Experimental equipment: HT—Hall thruster; 1—plasma plume; 2—quartz view port; 3—quartz lens; 4—optical fiber; 5—spectrometer USB-2000; 6—USB connection cable; and 7—computer.

#### Hall Thruster Erosion DOI: http://dx.doi.org/10.5772/intechopen.82654

5. OES studies of plasma propulsion systems of 100 W power

For the purpose of preliminary studies on the influence of the modes of operation of the HT on its erosion characteristics for measuring the spectra of optical radiation of the plasma plume of the 0.1 kW HT, an experimental installation was established, and the scheme of which is shown in Figure 9. The radiation emitted by the plasma (1) 0.1 kW HT, through a quartz window (2) (diameter—40 mm and thickness—5 mm), was observed with the help of a device for measuring the spectrum located outside the camera. The angle of observation of the plasma plume at the cut of the thruster was 60° relative to the chamber axis. With 10 cm of focusing lens (3), the radiation is collected and focused on a light conductive fiber (4) with a diameter of 1 mm. The distance between the plasma plume and the lens was about 80 cm. The end of the fiber is connected by a special adapter with the inlet of the spectrometer (5). A lens with a focal length of 10 cm gave an approximately 9 magnification. With an optical fiber with a diameter of 1 mm, the spatial resolution was about 9 mm in diameter. Thus, the installation captured most of the

The end of the optical cable through a special adapter connects to the input port of the spectrometer USB2000. This mini spectrometer, manufactured by Ocean Optics (USA), was kindly provided by Laboratoire de Physique des Gaz et des Plasmas [13]. It was used to cover the considered range with a resolution that is sufficient to distinguish optical lines. Moreover, its weight (600 g), dimensions (15 11 5 cm), and power consumption (100 mA) make it an excellent candidate for the use in summer experiments to study the erosion of ceramics and the

The light entering the spectrometer is refracted by a fixed dispersing prism and is directed to the charge coupled device (CCD) line with 2048 elements. So, in one shot, this spectrometer gives a spectrum of wavelengths of 380–830 nm with a resolution of 0.2–0.22 nm. The spectrometer is connected to the laptop through the USB interface. The computer controls the parameters of the spectrometer USB2000

Experimental equipment: HT—Hall thruster; 1—plasma plume; 2—quartz view port; 3—quartz lens;

4—optical fiber; 5—spectrometer USB-2000; 6—USB connection cable; and 7—computer.

5.1 OES test equipment and its adjustment

Propulsion Systems

plasma jet in the thruster channel and its exit.

Figure 9.

102

optical spectrum of the plasma plume of the thruster (Figure 9).

through the computer, and the graphical interface allows you to get a spectral image that can be saved as a text file with data.

Such a measuring spectrometric system is quite user-friendly, promptly provides information about the controlled spectrum, and thanks to a numerical representation, which allows us to conveniently store information and compare it with another that was previously registered.

#### 5.2 OES studies of the HT erosion at different operation modes

Together with the Laboratoire de Physique des Gaz et des Plasmas, Université Paris, Laboratoire d'Analyse Spectroscopique et d'Energétique des Plasmas, Université d'Orléans, and Laboratoire d'Aérothermique, the 0.1 W HT optical (spectral) tests were conducted to determine the effect of the regime of the thruster on its erosion characteristics [13]. Measurements were provided on the test facility as shown in Figure 10. Experimental parameters, discharge voltage (Ud) from 200 to 300 V in a step of 20 V, and also the current on the solenoid of the magnetic system of the thruster varied in the range of 1.5–3.0 A. All measurements were made at two values of xenon consumption: 0.20 and 0.25 mg/s.

In calculating of this parameter, the power of the solenoid was taken into account, but was neglected by the mass flow of xenon through the cathode. The maximum value of the efficiency (0.353) of HT during this experiment was obtained at a discharge voltage of 300 V and a mass flow of xenon 0.25 mg/s.

Xe I (828.01 nm), Xe II (484.43 nm), and Al I (396.152 nm) lines were distinguished for the analysis of the level of erosion of the stationary plasma thruster with ceramics from the ABN. These three lines made it possible to analyze the rate of erosion. This analysis was conducted with Laboratoire d'Analyze Spectroscopique et d'Energétique des Plasmas, Université d'Orléans, and its results are published in the materials of the international conference [14].

The Xe II xenon ion emission occurs due to the transition from the wavelength of 484.43 nm from the initial level 5p<sup>4</sup> D<sup>0</sup> 7=2 <sup>6</sup><sup>s</sup> <sup>5</sup>p5 2P<sup>0</sup> 3=2 <sup>6</sup><sup>p</sup> of 14.10 eV (80118.962 cm<sup>1</sup> ) to the level 11.54 eV (68045.156 cm<sup>1</sup> ). The upper level 5p<sup>4</sup> D<sup>0</sup> 7=2 <sup>6</sup><sup>s</sup> is basically two metastable states Xe<sup>+</sup> <sup>5</sup>d<sup>4</sup> D7=<sup>2</sup> and 5d<sup>4</sup> F7<sup>=</sup>2.

Figure 10. Dependence of the erosion from the discharge voltage.

The two lines for aluminum are close to the transition 3s 23<sup>p</sup> � <sup>3</sup><sup>s</sup> 24s to the emission level of 3.14 eV (25347.756 cm�<sup>1</sup> ): the first one is 394.40 nm and the other one is 396.15 nm. Second quasi-resonance line is used. This line Al I has an amplitude of light oscillations and is located near the lines Xe II. The spectrometer should be able to isolate this line from the set of other lines of Xenon plasma emission. To determine the exact value of the AlI line 396.15 nm, a numerical method of data approximation was used for Gauss's multi-profile profile. The radiation level is also occupied by electrons excited from their ground state 2P<sup>0</sup> 1=2 [15].

The erosion of the thruster ceramics was determined from the ratio of the intensity of the lines

$$E\_{roison} = \frac{I(AlI, \text{396 } nm) \cdot I(\text{XeI}, \text{828 } nm)}{I(\text{XeII}, \text{484 } nm)} . \tag{2}$$

To understand the behavior of the erosion in the regimes of thruster that were not measured, the experimental data were approximated by the polynomial depen-

Figure 11 shows the dependence of the erosion on the coil current for a constant

discharge voltage of 260 V and for two xenon mass flow (0.20 and 0.25 mg/s).

6. Experiment investigation of HT erosion with the OESSC and method

As it was told, before OESSC method was developed to provide measurements of the HT erosion of outer and inner ceramics separately during thruster test was not possible for OES method. In this section, the results of the experimental investigation of two erosion measurements methods are presented: direct measurements

In the Scientific-Technological Center of Space Power and Engines (STC SPE) of National Aerospace University named after N. Ye. Zhukovsky "Kharkiv Aviation Institute," Ukraine, 1.35 kW thruster was tested to determine the dependence of the DCh edges erosion parameters from the ratio of the magnetic system coil currents.

Algorithm of the diagnostics by OESSC method consists of several steps.

Measuring equipment of the OESSC method consists of a block of high-resolution spectrometers for recording the radiation of HT in the range with the wavelength of 240–830 nm and two-coordinate platform for moving the optical receiver and the

3.Measurement of the radiation spectra for each of the insulators and calculation

4.Repeat the procedure of spectrum measurements for each of the DCh ceramic

2. Launch of the thruster and stabilization of the erosion parameters.

1. Maintenance of the experimental equipment (Figure 12).

inserts at different ratios of the thruster coil currents.

dence (Figures 10 and 11).

Hall Thruster Erosion

of radial erosion measurements

DOI: http://dx.doi.org/10.5772/intechopen.82654

6.1 The algorithm of the OESSC method

The thruster was tested with a ceramic BN.

unit for protecting the optical fiber from dust.

of the erosion rate from the formula (6).

Figure 12.

105

Measuring equipment of the OESSC method.

radial erosion and OESSC method.

where intensities were determined as:

$$I\_{Al}(\text{396 nm}) \propto n\_{Al} \cdot n\_{\epsilon} \cdot k\_{\epsilon}^{Al}, I\_{\text{Xc}}(\text{828 nm}) \propto n\_{\text{Xc}} \cdot n\_{\epsilon} \cdot k\_{\epsilon}^{\text{Xe}}, I\_{\text{Xe}^{+m}}(\text{484 nm}) \propto n\_{\text{Xc}^{+m}} \cdot n\_{\epsilon} \cdot k\_{\epsilon}^{\text{Xe}^{+m}} \tag{3}$$

where nAl, nXe, nXeþ<sup>m</sup>—metastable aluminum densities, neutral and ionized xenon, respectively, and kAl <sup>e</sup> , kXe <sup>e</sup> , kXeþ<sup>m</sup> <sup>e</sup> —electronic excitation coefficient for the corresponding upper state. The ratio of the intensity of the lines is used to determine the concentration of aluminum.

$$E\_{ratio} = \frac{I(Al\,I,396\,nm) \cdot I(XeI, 828\,nm)}{I(Xe\,II, 484\,nm)} \propto \frac{n\_{Al} \cdot n\_{\varepsilon} \cdot k\_{\varepsilon}^{Al} \cdot n\_{Xe} \cdot n\_{\varepsilon} \cdot k\_{\varepsilon}^{Xe}}{n\_{Xe^{+}} \cdot n\_{\varepsilon} \cdot k\_{\varepsilon}^{Xe^{+}}}.\tag{4}$$

Applying an actinometric hypothesis, we obtain that the density of ionic metastable is proportional to the concentration of ions, reducing the value kAl <sup>e</sup> and <sup>k</sup>Xeþ<sup>m</sup> <sup>e</sup> because of the equivalence of energy change (several eV), and this equation takes the form:

Figure 11. Dependence of the erosion from the coil current.

Hall Thruster Erosion DOI: http://dx.doi.org/10.5772/intechopen.82654

The two lines for aluminum are close to the transition 3s

occupied by electrons excited from their ground state 2P<sup>0</sup>

<sup>e</sup> , kXe

is proportional to the concentration of ions, reducing the value kAl

Erosion <sup>∝</sup> nAl � ne � kAl

Erosion <sup>¼</sup> I Al I ð Þ� ; <sup>396</sup> nm I Xe I ð Þ ; <sup>828</sup> nm

<sup>e</sup> , kXeþ<sup>m</sup>

one is 396.15 nm. Second quasi-resonance line is used. This line Al I has an amplitude of light oscillations and is located near the lines Xe II. The spectrometer should be able to isolate this line from the set of other lines of Xenon plasma emission. To determine the exact value of the AlI line 396.15 nm, a numerical method of data approximation was used for Gauss's multi-profile profile. The radiation level is also

The erosion of the thruster ceramics was determined from the ratio of the

Erosion <sup>¼</sup> I Al I ð Þ� ; <sup>396</sup> nm I Xe I ð Þ ; <sup>828</sup> nm

<sup>e</sup> , IXeð Þ <sup>828</sup> nm <sup>∝</sup>nXe � ne � kXe

where nAl, nXe, nXeþ<sup>m</sup>—metastable aluminum densities, neutral and ionized

corresponding upper state. The ratio of the intensity of the lines is used to deter-

I Xe II ð Þ ; <sup>484</sup> nm <sup>∝</sup> nAl � ne � <sup>k</sup>Al

the equivalence of energy change (several eV), and this equation takes the form:

nXeþ<sup>m</sup> � n<sup>2</sup>

Applying an actinometric hypothesis, we obtain that the density of ionic metastable

<sup>e</sup> � nXe � ne � <sup>k</sup>Xe

<sup>e</sup> � kXe<sup>þ</sup> � <sup>k</sup>Xeþ<sup>m</sup> e

e

emission level of 3.14 eV (25347.756 cm�<sup>1</sup>

where intensities were determined as:

intensity of the lines

Propulsion Systems

IAlð Þ <sup>396</sup> nm <sup>∝</sup>nAl � ne � kAl

xenon, respectively, and kAl

Figure 11.

104

Dependence of the erosion from the coil current.

mine the concentration of aluminum.

23<sup>p</sup> � <sup>3</sup><sup>s</sup>

[15].

<sup>e</sup> , IXeþ<sup>m</sup> ð Þ <sup>484</sup> nm <sup>∝</sup>nXeþ<sup>m</sup> � ne � kXeþ<sup>m</sup>

<sup>e</sup> � nXe � ne � <sup>k</sup>Xe

e

<sup>e</sup> and <sup>k</sup>Xeþ<sup>m</sup>

∝nAl: (5)

e (3)

: (4)

<sup>e</sup> because of

): the first one is 394.40 nm and the other

1=2 

I Xe II ð Þ ; <sup>484</sup> nm : (2)

<sup>e</sup> —electronic excitation coefficient for the

nXeþ<sup>m</sup> � ne � <sup>k</sup>Xeþ<sup>m</sup> e

24s to the

To understand the behavior of the erosion in the regimes of thruster that were not measured, the experimental data were approximated by the polynomial dependence (Figures 10 and 11).

Figure 11 shows the dependence of the erosion on the coil current for a constant discharge voltage of 260 V and for two xenon mass flow (0.20 and 0.25 mg/s).

#### 6. Experiment investigation of HT erosion with the OESSC and method of radial erosion measurements

As it was told, before OESSC method was developed to provide measurements of the HT erosion of outer and inner ceramics separately during thruster test was not possible for OES method. In this section, the results of the experimental investigation of two erosion measurements methods are presented: direct measurements radial erosion and OESSC method.

#### 6.1 The algorithm of the OESSC method

In the Scientific-Technological Center of Space Power and Engines (STC SPE) of National Aerospace University named after N. Ye. Zhukovsky "Kharkiv Aviation Institute," Ukraine, 1.35 kW thruster was tested to determine the dependence of the DCh edges erosion parameters from the ratio of the magnetic system coil currents. The thruster was tested with a ceramic BN.

Algorithm of the diagnostics by OESSC method consists of several steps.

1. Maintenance of the experimental equipment (Figure 12).

Measuring equipment of the OESSC method consists of a block of high-resolution spectrometers for recording the radiation of HT in the range with the wavelength of 240–830 nm and two-coordinate platform for moving the optical receiver and the unit for protecting the optical fiber from dust.


Figure 12. Measuring equipment of the OESSC method.

5. Comparison of the erosion rate of insulators in different modes of HT.

The ratio of the erosion rates of edges obtained by the OESSC method (ΔEr) was determined like:

$$\Delta Er = \frac{Er\_{\text{inner\\_cremicic}} - Er\_{\text{outer\\_crcarrier}}}{Er\_{\text{inner\\_crcarrier}}} \cdot 100\% \quad (\text{if} \, Er\_{\text{inner\\_crcarrier}} \, Er\_{\text{outer\\_crcarrier}})or$$

$$\Delta Er = \frac{Er\_{\text{outer\\_crcarrier}} - Er\_{\text{inner\\_crcarrier}}}{Er\_{\text{outer\\_crcarrier}}} \cdot 100\% (\text{if} \, Er\_{\text{outer\\_crcarrier}} \, SEr\_{\text{inner\\_crcarrier}})$$

6.Development of recommendations regarding the optimal operating mode with the maximum lifetime of the thruster DCh edges.

#### 6.2 The method of the radial erosion measurements

After taking measurements by the OESSC method and selecting the operating mode, the thruster is tested on a time base of 50 hours, and measurements of radial erosion by a direct method are performed (Figure 13 and Table 1). As a basis for measuring the insulator edge thickness, the outer diameter of the outer insulator and the inner diameter of the inner insulator were chosen.

Measurements are provided with the help of an instrumental measuring microscope with an error of less than 0.0005 mm. To eliminate errors in the installation of the DCh relative to the microscope, each measured position is photographed.

Radial erosion is defined as the difference between the coordinates of the edges of the insulators before and after the thruster operation.

$$\begin{aligned} E\_{RAD\\_OUT} &= \mathbf{e}\_{OUT\\_INITIAL} - \mathbf{e}\_{OUT\\_FINAL} \\ E\_{RAD\\_IN} &= \mathbf{e}\_{IN\\_INITIAL} - \mathbf{e}\_{IN\\_FINAL} \end{aligned} \tag{7}$$

Comparison of erosion rates of external and internal insulators was carried out

eOUT\_INITIAL Thickness of outer ceramic before the experiment eOUT\_FINAL Thickness of outer ceramic after the experiment eIN\_INITIAL Thickness of inner ceramic before the experiment eIN\_FINAL Thickness of inner ceramic after the experiment

where VRAD\_max—erosion of the insulator for which the wear is greater, mm/h;

Thruster operation modes: I—mode with the lowest discharge current, current

Results of the erosion measurements with direct methods are presented in

As it is easy to see from the results on the mode with the lowest discharge

6.3 Comparison of results accepted by the OESSC method and direct radial

Figure 14 shows the results of the erosion rate measurements by the OESSC method. It was found that in the operating mode of the thruster with the minimum discharge current (mode I), the difference in the erosion rate of insulators was about 32%, which corresponded to the data obtained with direct measurements. After diagnostics and comparing the irregularity of erosion in 25 regimes, it was found that at the operating mode II, the difference in the erosion rate of the edges does not exceed 0.3%. The results were confirmed in subsequent tests and measurements by a direct method. The time spent on the study of the thruster DCh wear by the OESSC method on 25 operating modes, taking into account the expectation of the time for stabilizing the erosion rate, was 7 hours, which made it possible to reduce significantly the search for optimal current of the coils of the magnetic system from the point of view of uniform wear of the edges of the

� 100%, (9)

<sup>Δ</sup>VR <sup>¼</sup> VRAD\_max � VRAD\_min VRAD\_max

VRAD\_min—erosion of the insulator for which the wear is lower, mm/h.

DIN\_OUT Inner diameter of outer ceramic DOUT\_OUT Outer diameter of outer ceramic DOUT\_IN Outer diameter of inner ceramic DIN\_IN Inner diameter of inner ceramic

of inner coil—5 A, and current of outer coil—5 A.

current, erosion of ceramic walls is not uniform.
