**2.5 Temperature and water content effects**

The temperature of the high-energy gases in the TCA, using the basic ADNbased propellant, is higher than is normally tolerated by the standard hydrazine thrusters. In order to overcome this limitation, the ADN-based propellants can be adjusted so that the catalytic effect is maintained, whereas the reaction temperature is reduced in order to have the TCA materials of construction within their required temperature limits. For dual use thruster application, this issue can be tackled by (a) using suitable TCA materials with compatibility to higher temperatures and (b) lowering the gas temperature. A combination of both is also possible.

Even though using TCA materials suitable for higher temperatures is not dealt with here, it is notable that these would also be suitable for the lower temperature


**7**

*Green Comparable Alternatives of Hydrazines-Based Monopropellant and Bipropellant Rocket…*

hydrazine decomposition products, which are the chemically nonoxidizing N2,

As regards lowering the gas temperature, ECAPS presented their new LMP-103S #1127-3 propellant variant that also combusts at a lower temperature, giving a specific impulse comparable with the *Isp* for hydrazine. They stated that the lower combustion temperature may enable the usage of less expensive materials for the TCA [29]. For their low temperature derivative of LMP-103S, which ECAPS have recently developed, they have conducted hot-firing tests at their facility at FOI-Grindsjön [29, 39], as well as in cooperation with Airbus Safran Launchers (ASL) in the facility at DLR-Lampoldshausen [23, 40–43]. Although the declared intention of the ECAPS development was to handle significantly lower storage temperatures than specified for traditional storable monopropellants, for example, hydrazine (down to about −30°C), this propellant also exhibited a lower combustion temperature than LMP-103S, giving a specific impulse comparable with the *Isp* for hydrazine. The low-temperature derivative of the space-qualified LMP-103S was tested in a 22N development thruster, having 20% higher density than hydrazine, combusting at a lower temperature than LMP-103S and with *Isp* similar to

Lowering TCA gas temperatures by the effect of further dilution in water of established ADN-based propellants is expected to be in line with the reduction of the energetic content of the decomposition products. This leads directly to reduction in the temperature of the decomposition gases. The desired effect achieved by this is the possibility to use less demanding materials of construction, but with lower performance. Thruster performance is linked to the temperature directly, as follows. The specific impulse is proportional to the square root of the ratio between

expected, specific impulse is reduced too, per the square root of temperature.

to comparing hydrazine and the ADN-based monopropellant FLP-106 [8]. The specific impulse of FLP-106, as shown in **Table 2**, is higher than that of hydrazine,

A detailed investigation and analysis on the influence of the water content on the specific impulse and the thermochemical and density properties of the propellant has been presented in a recent conference by GRASP FP7 group participants. They presented the influence of water content on the ignition process and the spray behavior and the influence on the thermal field inside the combustion. The analysis of the spray behavior in vacuum near conditions was investigated by using different

\_\_\_\_\_\_

*Tc*/*M* , is similarly applicable

\_\_\_\_\_\_

*Tc*/*M*[44]. Therefore, as

temperature and molecular weight of the exhaust gas, or *Isp*~√

in accordance with its considerably higher chamber temperature.

The specific impulse relation to temperature, *Isp*~√

*DOI: http://dx.doi.org/10.5772/intechopen.82676*

*ADN-based liquid monopropellant catalytic decomposition [34].*

H2, and NH3.

**Figure 4.**

hydrazine [29].

#### **Table 1.**

*COTS parts and components flown in the Prisma satellite [33].*

*Green Comparable Alternatives of Hydrazines-Based Monopropellant and Bipropellant Rocket… DOI: http://dx.doi.org/10.5772/intechopen.82676*

#### **Figure 4.**

*Aerospace Engineering*

propulsion system components [31, 32].

presents these for the Prisma satellite.

filters, etc.), increases the flexibility, improves the reliability, and reduces the costs for introducing the ADN-based reduced hazards' propellants technology on future missions. FOI and ECAPS, who promote the ADN-based RHP FLP-106 and LMP-103S, respectively, have confirmed that most components can be COTS hydrazine

The materials, typically utilized for hydrazine propulsion systems, have been verified to be compatible with ADN-based liquid propellants and specifically with the LMP-103S space-proven aboard the Prisma satellite that was launched in June 2010 and operated successfully in space for 5 years. **Table 1** (based on Ref. [33])

In the present work, this concept is extended beyond previous works [23], to

The catalytic effect on the ADN-based propellant has been proven with the same

catalyst as in the hydrazine thruster. In the TCA, in which the chemical reaction converts the liquid propellant to the necessary high energy gases, there is need for a catalyst to induce such an ignition reaction. In laboratory thermochemical tests, it has been proven that the same iridium catalyst, which decomposes hydrazine, has a

Differential scanning calorimetry (DSC) analysis has revealed an exothermal peak around 150°C in addition to the endothermic peak around 85°C of ADN melting and the thermal decomposition at 190°C. The peak around 150°C is attributable to the catalytic effect of heated iridium-based catalyst, which does not appear at low temperatures [34]. This has also been found in previously published works [35–38]. Ignition tests with DSC analysis, such as depicted in **Figure 4**, demonstrate this effect.

The temperature of the high-energy gases in the TCA, using the basic ADNbased propellant, is higher than is normally tolerated by the standard hydrazine thrusters. In order to overcome this limitation, the ADN-based propellants can be adjusted so that the catalytic effect is maintained, whereas the reaction temperature is reduced in order to have the TCA materials of construction within their required temperature limits. For dual use thruster application, this issue can be tackled by (a) using suitable TCA materials with compatibility to higher temperatures and (b) lowering the gas temperature. A combination of both is also possible.

Even though using TCA materials suitable for higher temperatures is not dealt with here, it is notable that these would also be suitable for the lower temperature

include also COTS monopropellant hydrazine thrusters, as described below.

**2.4 Dual capability catalytic effect with ADN-based propellants**

definite catalytic effect on the ADN-based propellant.

**Component Supplier Status**

Service valves Moog Qualified Pressure transducer Bradford Qualified

Latch valve Moog Qualified

*COTS parts and components flown in the Prisma satellite [33].*

Propellant tank Rafael Delta-qual by Rafael

System filter Sofrance Delta-qual by Sofrance

Thruster ECAPS Qualified by ECAPS Pipes and brackets ECAPS/SSC Qualified on STM

**2.5 Temperature and water content effects**

**6**

**Table 1.**

*ADN-based liquid monopropellant catalytic decomposition [34].*

hydrazine decomposition products, which are the chemically nonoxidizing N2, H2, and NH3.

As regards lowering the gas temperature, ECAPS presented their new LMP-103S #1127-3 propellant variant that also combusts at a lower temperature, giving a specific impulse comparable with the *Isp* for hydrazine. They stated that the lower combustion temperature may enable the usage of less expensive materials for the TCA [29]. For their low temperature derivative of LMP-103S, which ECAPS have recently developed, they have conducted hot-firing tests at their facility at FOI-Grindsjön [29, 39], as well as in cooperation with Airbus Safran Launchers (ASL) in the facility at DLR-Lampoldshausen [23, 40–43]. Although the declared intention of the ECAPS development was to handle significantly lower storage temperatures than specified for traditional storable monopropellants, for example, hydrazine (down to about −30°C), this propellant also exhibited a lower combustion temperature than LMP-103S, giving a specific impulse comparable with the *Isp* for hydrazine. The low-temperature derivative of the space-qualified LMP-103S was tested in a 22N development thruster, having 20% higher density than hydrazine, combusting at a lower temperature than LMP-103S and with *Isp* similar to hydrazine [29].

Lowering TCA gas temperatures by the effect of further dilution in water of established ADN-based propellants is expected to be in line with the reduction of the energetic content of the decomposition products. This leads directly to reduction in the temperature of the decomposition gases. The desired effect achieved by this is the possibility to use less demanding materials of construction, but with lower performance. Thruster performance is linked to the temperature directly, as follows. The specific impulse is proportional to the square root of the ratio between temperature and molecular weight of the exhaust gas, or *Isp*~√ \_\_\_\_\_\_ *Tc*/*M*[44]. Therefore, as expected, specific impulse is reduced too, per the square root of temperature. \_\_\_\_\_\_

The specific impulse relation to temperature, *Isp*~√ *Tc*/*M* , is similarly applicable to comparing hydrazine and the ADN-based monopropellant FLP-106 [8]. The specific impulse of FLP-106, as shown in **Table 2**, is higher than that of hydrazine, in accordance with its considerably higher chamber temperature.

A detailed investigation and analysis on the influence of the water content on the specific impulse and the thermochemical and density properties of the propellant has been presented in a recent conference by GRASP FP7 group participants. They presented the influence of water content on the ignition process and the spray behavior and the influence on the thermal field inside the combustion. The analysis of the spray behavior in vacuum near conditions was investigated by using different


#### **Table 2.**

*Comparison of the properties of the monopropellants hydrazine and the ADN-based FLP-106 [8].*

blends. For the analysis of the combustion chamber temperatures, the temperatures and the heat flux inside the combustion chamber in relation to the water content were estimated [45], as well as the impact of the water content and the results for a 500N class thruster.

The following results were obtained in Ref. [45]. Expansion ratio ε = 50 and chamber pressure Pc = 20 bars were assumed, and the reaction was feasible in the investigated concentrations in water, according to the calculations made. The density reduction still leaves the blend with a considerably higher density than hydrazine, with the corresponding gain in density-specific impulse *ρIsp*, decreasing to the lowest value of 280 kg s/L. This is nevertheless approximately 22% higher than the value calculated for hydrazine. Blends bringing *Isp* down to values similar to those of hydrazine are considered. Density and *ρIsp* as a function of temperature for water and FLP-106 at various degrees of water content are presented in **Figure 5** (from Ref. [45]).

#### **2.6 Preheating temperature capability**

The dedicated ADN-based thrusters are ignited with a preheated catalyst. The ECAPS 1N thrusters, specifically developed for ADN-based monopropellant, use a 10 W heater. The preheating time is 1800 s. In the case of the Prisma thruster, the maximum load during preheating was 9.25 and 8.3 W during firing [46].

The necessary preheat temperature for inducing the reaction of ADN-based propellant is considerably higher than the 120–180°C of hydrazine decomposition. For nominal performance, the required preheat temperatures were in the order of 200–300°C [33]. FLP-106 was experimentally ignited thermally and by resistive heating, within less than 2 ms. An optimal preheating temperature of about 300°C was found where the ignition delay was minimized [47].

Preheating tests have been carried out with a conventional 1N monopropellant hydrazine thruster, with nominal electrical supply voltage in a vacuum chamber that simulates space conditions. These have shown the capability to achieve with a conventional hydrazine thruster the necessary higher preheating of ADN-based propellant, as depicted in **Figure 6** [34]. This heating period compares well with the abovementioned 1800 s of the Prisma satellite in-orbit performance.

#### **2.7 Proof-of-concept firing program**

After removing, within the described initial risk reduction program, the major uncertainties regarding the proposed dual capability of monopropellant hydrazine propulsion systems to operate as equivalent reduced hazards propellant (RHP) systems, a proof-of-concept firing program has been proposed. This program entails end-to-end proof by firing testing in simulated space environment of a representative engineering model (EM) propulsion system.

**9**

**Figure 5.**

**Figure 6.**

*[45].*

firing and depicted in **Figure 7** [1].

lants' traces being in contact with each other.

*Green Comparable Alternatives of Hydrazines-Based Monopropellant and Bipropellant Rocket…*

*Density and ρIsp as a function of temperature for water and FLP-106 with various values of water content* 

The proposed test setup is based on existing hydrazine propulsion systems vacuum chamber infrastructure adaptation to ADN-based monopropellant firing, without long-term interference with the capability to continue with hydrazine system tests. The test chamber is the one previously used for the Offek satellite EM

*1N thruster catalyst bed temperature preheat in simulated space conditions [34].*

The difference between the propellants requires attention to aspects of quality and safety to personnel, as well as to those of the testing infrastructure. Primarily, the ADN-based propellant, which is an oxidizer by nature, needs to be very strictly separated from hydrazine, which is fuel by nature. This can be achieved by temporarily disconnecting the existing hydrazine feed lines from the vacuum test chamber setup and maintaining the necessary separation distances according to the materials involved and their quantities. Moreover, the vacuum pump lines should be equally separated, in order to prevent any concern regarding carried over propel-

A location independent feed system has been designed for the ADN-based "green" monopropellant, which would be entirely enclosed within the testing

*DOI: http://dx.doi.org/10.5772/intechopen.82676*

*Green Comparable Alternatives of Hydrazines-Based Monopropellant and Bipropellant Rocket… DOI: http://dx.doi.org/10.5772/intechopen.82676*

#### **Figure 5.**

*Aerospace Engineering*

500N class thruster.

*expansion ratio ε = 50.*

**Table 2.**

**2.6 Preheating temperature capability**

**2.7 Proof-of-concept firing program**

tive engineering model (EM) propulsion system.

blends. For the analysis of the combustion chamber temperatures, the temperatures and the heat flux inside the combustion chamber in relation to the water content were estimated [45], as well as the impact of the water content and the results for a

*Comparison of the properties of the monopropellants hydrazine and the ADN-based FLP-106 [8].*

**Monopropellant Hydrazine FLP-106** Density *ρ* g/cm3 1.0037 1.357 Specific impulse based on mass flow *Isp* s 230 259 Specific impulse based on volume flow *ρIsp* s g/cm3 231 351 Chamber temperature *Tc* °C 1120 1880 *All properties at 25°C, Isp calculated for reaction chamber pressure Pc = 2.0 MPa, ambient pressure Pa = 0.0 MPa,* 

The following results were obtained in Ref. [45]. Expansion ratio ε = 50 and chamber pressure Pc = 20 bars were assumed, and the reaction was feasible in the investigated concentrations in water, according to the calculations made. The density reduction still leaves the blend with a considerably higher density than hydrazine, with the corresponding gain in density-specific impulse *ρIsp*, decreasing to the lowest value of 280 kg s/L. This is nevertheless approximately 22% higher than the value calculated for hydrazine. Blends bringing *Isp* down to values similar to those of hydrazine are considered. Density and *ρIsp* as a function of temperature for water and FLP-106 at

various degrees of water content are presented in **Figure 5** (from Ref. [45]).

maximum load during preheating was 9.25 and 8.3 W during firing [46].

abovementioned 1800 s of the Prisma satellite in-orbit performance.

was found where the ignition delay was minimized [47].

The dedicated ADN-based thrusters are ignited with a preheated catalyst. The ECAPS 1N thrusters, specifically developed for ADN-based monopropellant, use a 10 W heater. The preheating time is 1800 s. In the case of the Prisma thruster, the

The necessary preheat temperature for inducing the reaction of ADN-based propellant is considerably higher than the 120–180°C of hydrazine decomposition. For nominal performance, the required preheat temperatures were in the order of 200–300°C [33]. FLP-106 was experimentally ignited thermally and by resistive heating, within less than 2 ms. An optimal preheating temperature of about 300°C

Preheating tests have been carried out with a conventional 1N monopropellant hydrazine thruster, with nominal electrical supply voltage in a vacuum chamber that simulates space conditions. These have shown the capability to achieve with a conventional hydrazine thruster the necessary higher preheating of ADN-based propellant, as depicted in **Figure 6** [34]. This heating period compares well with the

After removing, within the described initial risk reduction program, the major uncertainties regarding the proposed dual capability of monopropellant hydrazine propulsion systems to operate as equivalent reduced hazards propellant (RHP) systems, a proof-of-concept firing program has been proposed. This program entails end-to-end proof by firing testing in simulated space environment of a representa-

**8**

*Density and ρIsp as a function of temperature for water and FLP-106 with various values of water content [45].*

#### **Figure 6.**

*1N thruster catalyst bed temperature preheat in simulated space conditions [34].*

The proposed test setup is based on existing hydrazine propulsion systems vacuum chamber infrastructure adaptation to ADN-based monopropellant firing, without long-term interference with the capability to continue with hydrazine system tests. The test chamber is the one previously used for the Offek satellite EM firing and depicted in **Figure 7** [1].

The difference between the propellants requires attention to aspects of quality and safety to personnel, as well as to those of the testing infrastructure. Primarily, the ADN-based propellant, which is an oxidizer by nature, needs to be very strictly separated from hydrazine, which is fuel by nature. This can be achieved by temporarily disconnecting the existing hydrazine feed lines from the vacuum test chamber setup and maintaining the necessary separation distances according to the materials involved and their quantities. Moreover, the vacuum pump lines should be equally separated, in order to prevent any concern regarding carried over propellants' traces being in contact with each other.

A location independent feed system has been designed for the ADN-based "green" monopropellant, which would be entirely enclosed within the testing

#### *Aerospace Engineering*

vacuum chamber (**Figure 7**). This is consistent with the end-to-end testing of the entire system, as was done with the hydrazine system EM firing tests.

Here, the handling procedures can be simplified, thanks to the reduced hazards involved with the handling of RHP. The Prisma satellite fueling campaign serves as an example for that, as illustrated in **Figure 8**. During the launch campaign of the Prisma satellite, the first in-space demonstration of an ADN-based propulsion system, ECAPS loaded the hydrazine and RHP propellants at the Yasny launch base. The handling of ADN-based RHP was evaluated and declared as a "nonhazardous operation" by the Range Safety, so SCAPE suits were not required during the Prisma ADN-based propellant loading operation [20, 48].

The firing program, like any new type of testing, needs to go through the common procedural requirements. These include safety reviews and safety approvals, test procedure preparation and approval, allocation of thruster and propellant, and allocation of the test facility. The procedural requirements have been fulfilled, with the exception of the facility allocation.

**Figure 7.** *Entire system EM testing vacuum chamber [1].*

**11**

*Green Comparable Alternatives of Hydrazines-Based Monopropellant and Bipropellant Rocket…*

An initial risk reduction program has been performed for the concept of dualcapability propulsion systems. This was done by analysis of data as well as by dedicated tests. The program included proof of concept of dual use of all propulsion system parts and components, such as thrusters, valves, diaphragm tanks, pressure transducers, and pipework. The dual use of the propulsion systems' key components, the thrusters, is beyond any previous work. Both material compatibility and actual operation have been justified for both hydrazine and RHP, in view of an eventual system end-to-end proof by firing testing in space simulation environment.

The concept of dual-capability systems may serve as a vehicle toward gradual migration from monopropellant hydrazine propulsion systems to equivalent RHP systems. Hydrazine systems are prevalent in several applications and are still often the systems of choice in space propulsion as well as in other applications. The slow introduction rate of RHP or "green propellants," into space systems due to the conservatism of the space propulsion industry may be expedited thanks to the possibility for gradual conversion by dual capability of conventional hydrazine systems and ADN-based RHP. The presented propulsion system concept may accept last moment decisions on fueling with either hydrazine or with an RHP. This flexibility enables project progress until a very late stage without necessary commitment to either of the propellants, thus allowing a smoother transfer

This section describes a hypergolic system based on kerosene and hydrogen peroxide, similar in performance to MMH/N2O4 that has been developed by NewRocket© [24]. The NewRocket Green Propellant (NRGP) hypergolic bipropellant is based on concentrated hydrogen peroxide (HTP—high test peroxide) as oxidizer and on a kerosene-based fuel. NRGP is used in a family of bipropellant rocket and gas-generator applications. Neat HTP and kerosene are not hypergolic, while NRGP has been made such by addition of a minute amount of a solid energetic activator to the fuel. The activator is maintained homogeneously distributed in the fuel by its suitable gelation to a shear-thinning yield-stress fluid. Shear-thinning fluids exhibit decreased viscosity with increasing applied shear stresses, such as by pressure gradients (ΔP). The shear-thinning feature of the fuel enables its full functionality in propulsion systems, including pressurized or pumped feed flow

Usually, decomposition of hydrogen peroxide is achieved using catalyst beds

Another method is based on the idea of using catalytic or reactive material (such

**3. Comparable bipropellant rocket propulsion system**

and injection to the reaction chamber, just like any liquid propellant.

fuel; however, the system complexity and weight are both increased.

based on silver, platinum, and other materials. Catalyst beds produce hightemperature-decomposed hydrogen peroxide that can burn with a hydrocarbon

as metal oxides—MnO2, PbO2, F2O3, etc.) that is dissolved in a liquid fuel. The reactive material decomposes hydrogen peroxide and ignites the fuel, so hypergolic ignition is achieved without the use of a catalyst bed. However, this method requires fuels such as ethanol or methanol that serve as solvents for the reactive material. All these solvents used either alone or with kerosene-based fuels and have relatively low heat of combustion; therefore, the energetic performance of the system is low.

*DOI: http://dx.doi.org/10.5772/intechopen.82676*

from hydrazine to RHP.

**3.1 Overview**

**2.8 Concluding remarks for monopropellant**

**Figure 8.** *Propellant loading of satellite Prisma [20].*

*Green Comparable Alternatives of Hydrazines-Based Monopropellant and Bipropellant Rocket… DOI: http://dx.doi.org/10.5772/intechopen.82676*
