**Suborbital Flight**

**Chapter 10**

**Provisional chapter**

**Suborbital Flight: An Affordable and Feasible Option**

**Suborbital Flight: An Affordable and Feasible Option** 

Suborbital flights are a low-cost option for universities. To perform suborbital missions, it is necessary to design, plan, test, verify, and validate each and every one of the subsystems that integrate the payload without leaving the Earth. In Mexico, some experiments have been carried out since the 1990s to test communication systems in case of disaster and emergency. The Mexican Service Gondola (CSM) from 2015 to date has made suborbital flights in conjunction with the National Polytechnic Institute and the group of Protective Coatings Resistant to Thermal Changes and Cosmic Radiation (CRTCR) to test communication systems and glass-ceramic coatings. Suborbital flights are a great opportunity to explore the national territory and test new communication systems, structures,

**Keywords:** suborbital flight, mexican service gondola, mission design, communication

Every day, there are new advances in the space area, specifically in satellites. Each time it becomes more complex to access space, due to requirements, restrictions (mass, volume,

> © 2016 The Author(s). Licensee InTech. 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.

© 2018 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.

DOI: 10.5772/intechopen.73859

**for Mexican Aerospace Development**

**for Mexican Aerospace Development**

Rafael Prieto-Melendez, Leopoldo Ruiz-Huerta and

Additional information is available at the end of the chapter

Additional information is available at the end of the chapter

Barbara Bermudez-Reyes, Frederic Trillaud,

Frederic Trillaud, Fernando Velazquez-Villegas, Jonathan Remba, Ana M. Arizmendi-Morquecho, Alberto Caballero-Ruíz, Mario A. Mendoza-Barcenas, Rafael Prieto-Melendez, Leopoldo Ruiz-Huerta and

Fernando Velazquez-Villegas,

Barbara Bermudez-Reyes,

Ana M. Arizmendi-Morquecho,

Mario A. Mendoza-Barcenas,

http://dx.doi.org/10.5772/intechopen.73859

Jonathan Remba-Uribe,

Alberto Caballero-Ruíz,

Lauro Santiago-Cruz

**Abstract**

and materials.

**1. Introduction**

systems, glass-ceramic coatings

Lauro Santiago-Cruz

#### **Suborbital Flight: An Affordable and Feasible Option for Mexican Aerospace Development Suborbital Flight: An Affordable and Feasible Option for Mexican Aerospace Development**

DOI: 10.5772/intechopen.73859

Barbara Bermudez-Reyes, Frederic Trillaud, Fernando Velazquez-Villegas, Jonathan Remba-Uribe, Ana M. Arizmendi-Morquecho, Alberto Caballero-Ruíz, Mario A. Mendoza-Barcenas, Rafael Prieto-Melendez, Leopoldo Ruiz-Huerta and Lauro Santiago-Cruz Barbara Bermudez-Reyes, Frederic Trillaud, Fernando Velazquez-Villegas, Jonathan Remba, Ana M. Arizmendi-Morquecho, Alberto Caballero-Ruíz, Mario A. Mendoza-Barcenas, Rafael Prieto-Melendez, Leopoldo Ruiz-Huerta and Lauro Santiago-Cruz Additional information is available at the end of the chapter

Additional information is available at the end of the chapter

http://dx.doi.org/10.5772/intechopen.73859

#### **Abstract**

Suborbital flights are a low-cost option for universities. To perform suborbital missions, it is necessary to design, plan, test, verify, and validate each and every one of the subsystems that integrate the payload without leaving the Earth. In Mexico, some experiments have been carried out since the 1990s to test communication systems in case of disaster and emergency. The Mexican Service Gondola (CSM) from 2015 to date has made suborbital flights in conjunction with the National Polytechnic Institute and the group of Protective Coatings Resistant to Thermal Changes and Cosmic Radiation (CRTCR) to test communication systems and glass-ceramic coatings. Suborbital flights are a great opportunity to explore the national territory and test new communication systems, structures, and materials.

**Keywords:** suborbital flight, mexican service gondola, mission design, communication systems, glass-ceramic coatings

#### **1. Introduction**

Every day, there are new advances in the space area, specifically in satellites. Each time it becomes more complex to access space, due to requirements, restrictions (mass, volume,

© 2016 The Author(s). Licensee InTech. 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. © 2018 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.

structural, etc.), cost and regulations, integration tests, and orbit assignment. For this, suborbital flights have become a viable option for probes of nanosatellite systems [1]. A suborbital flight can exceed 100 km in height and will not orbit the earth; that is, it will not leave the Earth's atmosphere [2]. Therefore, suborbital flights are viable for testing various subsystems and segments that are composing picosatellite and nanosatellite systems [3]. A nanosatellite system is characterized by its mass (1–10 kg), and its geometry can be cubic (CubeSat) or cylindrical (TubeSat). It should be noted that these nanosatellites can be composed by units of 1 kg (1 U) up to 10 units (10 U) [4]. These characteristics allow you to adjust and perform specific experiments or test subsystems in a timely manner and recover the nanosatellite, only if it does not go into space [3]. In addition, one of the advantages of performing suborbital flights is that you have a wide range of launch platforms as rockets, UAVs, and stratospheric balloons to climb into the high atmosphere [5]. The stratospheric balloons are an affordable platform for uploading nanosatellites and allow measurements during the ascent (infrared and ultraviolet radiation, X-rays, gamma rays, photographic recognition, and video capture) [6]. It also allows testing deployable systems (parachutes) for recovery of payloads without major damage [3]. For all the above, in Mexico, suborbital flights have become an alternative for sensors, communication, attitude, electrical subsystems, new materials resistant to thermal changes and cosmic radiation, etc., and the most important thing is that the universities allow the formation of human resources in the space area.

#### **2. Suborbital flight: a window into space**

In terms of altitude, a suborbital flight is limited by the Kármán line, which is a line 100 km above sea level. The Fédération Aéronautique Internationale defines this limit because it is roughly the point where a vehicle flying fast enough to support itself with aerodynamic lift from the Earth's atmosphere would be flying faster than orbital speed [7]. The Kármán line covers the troposphere (until 20 km), the stratosphere (until 50 km), the mesosphere (until 85 km), and a little section of the thermosphere, which extends until 690 km. **Figure 1** shows some physical properties of the atmosphere vs. altitude; this demonstrates that suborbital flight implies interesting conditions to implement scientific experiments [8].

and RCS-V, RCS-VI, RCS-VII, and RCS-VIII completing two test flights and eight full flights over the past two decades. In generally, flights only the RCS-V gondola was lost. The rate at recovery is over 90% owing to the use of onboard GPS that was not available at the time of the RCS-V project. The payloads were typically a 40 m band 4 W transmitter, 2 m band 0.7 W

**Figure 1.** Comparison of the 1962 US standard atmosphere graph of geometric altitude against density, pressure, the

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187

In 2009, the CRAEG (Club de Radio Amateur del Estado de Guanajuato A.C.) launched their first project and one of the most complexes to date, SARSEM-ICARUS 1 (Mexican Aerostatic Subspace Repeater System). The objective of this project was to provide wide area communications in case of disaster or emergency for handheld and mobile VHF and UHF radios. The system carried a flight computer, a camera, temperature sensors, an onboard radio repeater, a dual GPS system, and a data communication system to download data and upload commands for controlling remotely specific subsystems. The gondola reached 28.8 km, had a radio coverage of 800 km, and was successfully recovered. In the following years (2010, 2011, 2013), SARSEM II, SARSEM III, and SARSEM IV were improved and redesigned, reaching their

transmitter, altimeter, two temperature sensors, and a flight computer (**Figure 2**) [10].

nominal altitudes and coverage as expected (**Table 1** and **Figure 3**) [11].

speed of sound, and temperature with approximate altitudes of various objects [8].

Perhaps, the most attractive characteristic of suborbital flight is its cost, which is very low in comparison with space flight. According to NASA, today, it costs \$10,000 to put a pound of payload in Earth orbit [9]. However, it costs about \$1000 per pound to make a suborbital mission and that includes launching, tracking, and recovering the payload.

#### **3. Suborbital flight: historical panorama**

Mexico has been entering the area of suborbital flights in the late 1990s. Since then, a group of amateur radio operators from the "Radio Club Satélite" has sent several high-altitude balloons for experimental communication projects such as TSAT-1, TSAT-2, TSAT-3, and TSAT-4 Suborbital Flight: An Affordable and Feasible Option for Mexican Aerospace Development http://dx.doi.org/10.5772/intechopen.73859 187

structural, etc.), cost and regulations, integration tests, and orbit assignment. For this, suborbital flights have become a viable option for probes of nanosatellite systems [1]. A suborbital flight can exceed 100 km in height and will not orbit the earth; that is, it will not leave the Earth's atmosphere [2]. Therefore, suborbital flights are viable for testing various subsystems and segments that are composing picosatellite and nanosatellite systems [3]. A nanosatellite system is characterized by its mass (1–10 kg), and its geometry can be cubic (CubeSat) or cylindrical (TubeSat). It should be noted that these nanosatellites can be composed by units of 1 kg (1 U) up to 10 units (10 U) [4]. These characteristics allow you to adjust and perform specific experiments or test subsystems in a timely manner and recover the nanosatellite, only if it does not go into space [3]. In addition, one of the advantages of performing suborbital flights is that you have a wide range of launch platforms as rockets, UAVs, and stratospheric balloons to climb into the high atmosphere [5]. The stratospheric balloons are an affordable platform for uploading nanosatellites and allow measurements during the ascent (infrared and ultraviolet radiation, X-rays, gamma rays, photographic recognition, and video capture) [6]. It also allows testing deployable systems (parachutes) for recovery of payloads without major damage [3]. For all the above, in Mexico, suborbital flights have become an alternative for sensors, communication, attitude, electrical subsystems, new materials resistant to thermal changes and cosmic radiation, etc., and the most important thing is that the universities allow the formation of human resources in the space

In terms of altitude, a suborbital flight is limited by the Kármán line, which is a line 100 km above sea level. The Fédération Aéronautique Internationale defines this limit because it is roughly the point where a vehicle flying fast enough to support itself with aerodynamic lift from the Earth's atmosphere would be flying faster than orbital speed [7]. The Kármán line covers the troposphere (until 20 km), the stratosphere (until 50 km), the mesosphere (until 85 km), and a little section of the thermosphere, which extends until 690 km. **Figure 1** shows some physical properties of the atmosphere vs. altitude; this demonstrates that suborbital

Perhaps, the most attractive characteristic of suborbital flight is its cost, which is very low in comparison with space flight. According to NASA, today, it costs \$10,000 to put a pound of payload in Earth orbit [9]. However, it costs about \$1000 per pound to make a suborbital mis-

Mexico has been entering the area of suborbital flights in the late 1990s. Since then, a group of amateur radio operators from the "Radio Club Satélite" has sent several high-altitude balloons for experimental communication projects such as TSAT-1, TSAT-2, TSAT-3, and TSAT-4

flight implies interesting conditions to implement scientific experiments [8].

sion and that includes launching, tracking, and recovering the payload.

area.

186 Space Flight

**2. Suborbital flight: a window into space**

**3. Suborbital flight: historical panorama**

**Figure 1.** Comparison of the 1962 US standard atmosphere graph of geometric altitude against density, pressure, the speed of sound, and temperature with approximate altitudes of various objects [8].

and RCS-V, RCS-VI, RCS-VII, and RCS-VIII completing two test flights and eight full flights over the past two decades. In generally, flights only the RCS-V gondola was lost. The rate at recovery is over 90% owing to the use of onboard GPS that was not available at the time of the RCS-V project. The payloads were typically a 40 m band 4 W transmitter, 2 m band 0.7 W transmitter, altimeter, two temperature sensors, and a flight computer (**Figure 2**) [10].

In 2009, the CRAEG (Club de Radio Amateur del Estado de Guanajuato A.C.) launched their first project and one of the most complexes to date, SARSEM-ICARUS 1 (Mexican Aerostatic Subspace Repeater System). The objective of this project was to provide wide area communications in case of disaster or emergency for handheld and mobile VHF and UHF radios. The system carried a flight computer, a camera, temperature sensors, an onboard radio repeater, a dual GPS system, and a data communication system to download data and upload commands for controlling remotely specific subsystems. The gondola reached 28.8 km, had a radio coverage of 800 km, and was successfully recovered. In the following years (2010, 2011, 2013), SARSEM II, SARSEM III, and SARSEM IV were improved and redesigned, reaching their nominal altitudes and coverage as expected (**Table 1** and **Figure 3**) [11].

the leadership of the Institute of Nuclear Sciences of the National Autonomous University of Mexico (UNAM) [12]. The idea was to build a suborbital platform to promote aerospace technologies by means of stratospheric flights between the altitudes of 25 and 35 km to test electronic systems and detectors used in satellites. This project followed the need to increment the aerospace technologies following the creation of the Mexican Space Agency around 2010, merging the effort of the national academic and industrial sectors [13]. From the start, the participation of undergrad and graduate students was considered essential for the future development of the technology in Mexico. It should be noted that suborbital platforms provide a cheap and easy-to-handle test facility based on sounding stratospheric balloons, which have flight durations of the order of a few hours. They are useful tools to test in near-space conditions of various aerospace systems and subsystems. In 2015 and 2016, two suborbital flights were successfully carried out over the state of Guanajuato, Mexico. A first gondola of 2.5 kg was launched, the CSM-1 housed one of the subsystems of Ulises 2.0, a nanosatellite developed by the Unidad de Alta Tecnología (UAT) of the Faculty of Engineering of the UNAM. It reached an altitude of nearly 31 km monitoring the thermal behavior of the electronic payload in addition to the temperatures inside and outside the gondola. This first experience was achieved because of a successful collaboration between the Institute of Engineering, the Engineering Faculty of the UNAM, the company, Remtronic Telecomunicaciones, and the Amateur Radio Club of the state of Guanajuato (CRAEG). The second gondola, CSM-2, with a weight of 2.1 kg, was an improvement of the first version, CSM-1, which had a crude structural design. Indeed, during the first flight, the gondola reached a velocity of about 8 m/s at landing despite the use of a parachute. The force of the impact deformed the structure, but the payload did not suffer any damages thanks to a custom-designed floating structure holding the payload inside the gondola. However, it appeared necessary to improve the impact absorbers, and a new design was tested during the second flight. This second flight carried two payloads, a monitoring atmospheric electronic system (SADM-1) and an experiment to try out a ceramic coating for satellites [14, 15]. Indeed, the new design allowed lowering the overall mass of the gondola while ensuring a better absorption at impact. This flight reached an altitude of 34 km. The external temperature of the gondola reached −70°C over a flight duration of about 2 h. For this second flight, the original collaboration included a new member, the National Laboratory for Additive Manufacturing, 3D Digitization and Computed Tomography (MADiT) of the Center of Applied Sciences and Technological Development (CCADET) of the UNAM. MADiT has strong capabilities and experience in design and manufacturing by means of additive manufacturing technologies. This national laboratory collabo-

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rates in the development of the structure of the gondola.

with additional clients.

Part of the development of the gondola, CSM, a series of studies covering mechanical and thermal aspects have been conducted to improve the reliability of the structure and to diminish the risks, mainly associated with cold temperatures (<−60°C) and impact at landing [16, 17]. **Figure 4** shows a comparison between experimental and numerical values of the altitude of the balloon versus time of flight for both flights [17]. **Figure 5** is a photograph of the CSM-2 team just before the launch from the Explora Science Center, León, Guanajuato, Mexico, on November 2016. A new flight of CSM-2 is scheduled on April 2017

**Figure 2.** Projects (a) TSAT-1 (1992) and (b) RCS-V, RCS AC (1996).


**Table 1.** Mexican suborbital missions [11].

**Figure 3.** SARSEM-ICARUS III, CRAEG AC (2011).

#### **4. Mexican missions: CSM**

Following the achieved successes over the past two decades, the "Carga de Servicio Mexicana" (CSM) or Mexican Service Gondola was born in 2014. It should be mentioned that this project was initiated as a spin-off of the Pixqui payload which flew in a NASA gondola on August 2013 with the participation of the Engineering Faculty of the Institute of Engineering under the leadership of the Institute of Nuclear Sciences of the National Autonomous University of Mexico (UNAM) [12]. The idea was to build a suborbital platform to promote aerospace technologies by means of stratospheric flights between the altitudes of 25 and 35 km to test electronic systems and detectors used in satellites. This project followed the need to increment the aerospace technologies following the creation of the Mexican Space Agency around 2010, merging the effort of the national academic and industrial sectors [13]. From the start, the participation of undergrad and graduate students was considered essential for the future development of the technology in Mexico. It should be noted that suborbital platforms provide a cheap and easy-to-handle test facility based on sounding stratospheric balloons, which have flight durations of the order of a few hours. They are useful tools to test in near-space conditions of various aerospace systems and subsystems. In 2015 and 2016, two suborbital flights were successfully carried out over the state of Guanajuato, Mexico. A first gondola of 2.5 kg was launched, the CSM-1 housed one of the subsystems of Ulises 2.0, a nanosatellite developed by the Unidad de Alta Tecnología (UAT) of the Faculty of Engineering of the UNAM. It reached an altitude of nearly 31 km monitoring the thermal behavior of the electronic payload in addition to the temperatures inside and outside the gondola. This first experience was achieved because of a successful collaboration between the Institute of Engineering, the Engineering Faculty of the UNAM, the company, Remtronic Telecomunicaciones, and the Amateur Radio Club of the state of Guanajuato (CRAEG). The second gondola, CSM-2, with a weight of 2.1 kg, was an improvement of the first version, CSM-1, which had a crude structural design. Indeed, during the first flight, the gondola reached a velocity of about 8 m/s at landing despite the use of a parachute. The force of the impact deformed the structure, but the payload did not suffer any damages thanks to a custom-designed floating structure holding the payload inside the gondola. However, it appeared necessary to improve the impact absorbers, and a new design was tested during the second flight. This second flight carried two payloads, a monitoring atmospheric electronic system (SADM-1) and an experiment to try out a ceramic coating for satellites [14, 15]. Indeed, the new design allowed lowering the overall mass of the gondola while ensuring a better absorption at impact. This flight reached an altitude of 34 km. The external temperature of the gondola reached −70°C over a flight duration of about 2 h. For this second flight, the original collaboration included a new member, the National Laboratory for Additive Manufacturing, 3D Digitization and Computed Tomography (MADiT) of the Center of Applied Sciences and Technological Development (CCADET) of the UNAM. MADiT has strong capabilities and experience in design and manufacturing by means of additive manufacturing technologies. This national laboratory collaborates in the development of the structure of the gondola.

Part of the development of the gondola, CSM, a series of studies covering mechanical and thermal aspects have been conducted to improve the reliability of the structure and to diminish the risks, mainly associated with cold temperatures (<−60°C) and impact at landing [16, 17]. **Figure 4** shows a comparison between experimental and numerical values of the altitude of the balloon versus time of flight for both flights [17]. **Figure 5** is a photograph of the CSM-2 team just before the launch from the Explora Science Center, León, Guanajuato, Mexico, on November 2016. A new flight of CSM-2 is scheduled on April 2017 with additional clients.

**4. Mexican missions: CSM**

**Table 1.** Mexican suborbital missions [11].

188 Space Flight

**Figure 3.** SARSEM-ICARUS III, CRAEG AC (2011).

**Figure 2.** Projects (a) TSAT-1 (1992) and (b) RCS-V, RCS AC (1996).

**Mission Altitude (Km)** SARSEM II (2010) 28.7–800 km SARSEM III (2011) 33.6–900 km SARSEM IV (2013) 32.4–900 km

Following the achieved successes over the past two decades, the "Carga de Servicio Mexicana" (CSM) or Mexican Service Gondola was born in 2014. It should be mentioned that this project was initiated as a spin-off of the Pixqui payload which flew in a NASA gondola on August 2013 with the participation of the Engineering Faculty of the Institute of Engineering under

In the manufacture of satellites, aluminum alloys maintain the leadership among lightweight and relatively low-cost structural metal materials, followed by alloys based on titanium [20, 21]

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Glass-ceramic coatings have proven to be a key technology in thermal stability and protection coatings in extreme environments such as aerospace gas turbines, which together with the engine operate at high temperatures (1370–1425°C) [22, 23]. So, the selection of materials for thermal barrier coatings is restricted by some basic requirements: (1) high melting point of the ceramics, (2) they must not present phase transformation between the ambient temperature and the operating temperature, (3) low thermal conductivity, (4) chemically inert in space environment, (5) the thermal expansion must be related to that of the metallic substrate,

for their low density and high resistance to corrosion.

**Figure 6.** SEM images coatings: (a) SiO2

chemical mapping, (e) SiO2

TiO2


surface, and (f) SiO2


surface, (b) SiO2


chemical mapping.


chemical mapping, (c) SiO2


surface, (d) SiO2


**Figure 4.** Comparison between the recorded altitude and the modeling one as a function of time [17].

The development of space technologies requires the use of stratospheric balloons since they offer a cheap, near-space environmental platform to test electronic systems of any kind. For the Mexican team developing CSM, it is the belief that it can lead to a sustainable development of the aerospace sector in the country involving the academic, governmental, military, and industrial sector. The first step has been carried out in that direction, and the future goal is to lift greater mass of a few tens of kilograms to improve the service already provided by CSM.

On the other hand, in Mexico, they have been designing glass-ceramic coverings to protect satellite systems of cosmic radiation and thermal changes. Aerospace materials must be lightweight and resistant to structural stresses, as well as to conditions in space [18, 19].

In the manufacture of satellites, aluminum alloys maintain the leadership among lightweight and relatively low-cost structural metal materials, followed by alloys based on titanium [20, 21] for their low density and high resistance to corrosion.

Glass-ceramic coatings have proven to be a key technology in thermal stability and protection coatings in extreme environments such as aerospace gas turbines, which together with the engine operate at high temperatures (1370–1425°C) [22, 23]. So, the selection of materials for thermal barrier coatings is restricted by some basic requirements: (1) high melting point of the ceramics, (2) they must not present phase transformation between the ambient temperature and the operating temperature, (3) low thermal conductivity, (4) chemically inert in space environment, (5) the thermal expansion must be related to that of the metallic substrate,

The development of space technologies requires the use of stratospheric balloons since they offer a cheap, near-space environmental platform to test electronic systems of any kind. For the Mexican team developing CSM, it is the belief that it can lead to a sustainable development of the aerospace sector in the country involving the academic, governmental, military, and industrial sector. The first step has been carried out in that direction, and the future goal is to lift greater mass of a few tens of kilograms to improve the service already pro-

**Figure 5.** CSM-2 team just before the launch of CSM-2 in November 2016 at the Explora Science Center, León, Guanajuato,

**Figure 4.** Comparison between the recorded altitude and the modeling one as a function of time [17].

On the other hand, in Mexico, they have been designing glass-ceramic coverings to protect satellite systems of cosmic radiation and thermal changes. Aerospace materials must be light-

weight and resistant to structural stresses, as well as to conditions in space [18, 19].

vided by CSM.

Mexico.

190 Space Flight

**Figure 6.** SEM images coatings: (a) SiO2 -Al2 O3 surface, (b) SiO2 -Al2 O3 chemical mapping, (c) SiO2 -TiO2 surface, (d) SiO2 - TiO2 chemical mapping, (e) SiO2 -SiO2 surface, and (f) SiO2 -SiO2 chemical mapping.

(6) good adhesion to the metallic substrate, and (7) low heating rate during the sintering process [24, 25]. The materials that can be used as glass-ceramic coatings is very limited; so far, only some materials have been found that meet these requirements [26]: aluminum oxide (Al2 O3 ), mullite (3Al2 O3 -2SiO2 ), cordierite (2MgO-2Al2 O3 5SiO2 ), zirconium oxide (ZrO2 ), and zirconia stabilized with yttrium (ZrO2 -Y2 O3 ). These coatings have been initially proposed for use as thermal control surfaces in aircraft because lightweight coatings with good adhesive properties can be obtained under thermal shock conditions [27]. This is why it could be considered that glass-ceramic coatings applied in satellite systems could resist ultraviolet, cosmic, and high-energy particles over a wide temperature range [28].

balloons of larger flight durations should be investigated and tried out in conjunctions with

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Besides the practicability of the launch, tracking and recovery of the gondola, and the payload, it is important that the structure, which the gondola offers, houses onboard electronics and power services. The system should be able to monitor the environment and itself and to supply the payloads with the necessary power for their operation. Multidisciplinary teams of engineers in mechanics, aerospace technology, telecommunication, electrical and electronic systems, and aeronautics should be part of this effort, and likely a national center for aerospace technology can be the right place to gather those specialties. There is an important risk that allocating resources to small individual projects may not be fructuous on the long term and it may be sensible to lay out one or a couple of projects of national interest that can lead

Mexico has been involved into suborbital flights since the past two decades through amateur impulse. Lately, the academic sector has been starting to promote the use of stratospheric sounding balloons to provide the scientific and technological communities and inexpensive, easy-to-use facility to test their aerospace technologies. A few suborbital flights over the country have been carried out by different institutions with a large participation of students. Among the different projects, CSM is the only one involving mostly researchers and engineers from the academic and private sectors with the objective to professionalize this area of knowledge following the example of leading countries. This platform has provided two flights of 2-h duration for altitudes over 30 km to three clients. It is expected to provide yearly

Overall, suborbital flights are the first step for Mexico to get involved in the aerospace development. Increasing funding, engineering schools, and national laboratories dedicated to aerospace studies and technological and scientific projects with the participation of the government, the military and the private sector is considered mandatory to achieve a significant

The CSM team would like to thank the Institute of Engineering and the Faculty of Engineering

They also thank the DGAPA for the support through the project IN113315. Finally, the authors really appreciate the participation and enthusiasm of Eduardo Amaro Calderon,

flights to future clients accommodating payloads of larger masses.

large masses of a few hundreds of kilograms to a ton.

the effort with a sustainable funding.

contribution to the field in Latin America.

**Acknowledgements**

of the UNAM for sponsoring CSM.

Diego Dominguez Baez and Alfredo Sanchez Labra.

**6. Conclusions**

Therefore, the team of Coatings Resistant to Thermal Changes and Cosmic Radiation (CRTCR) from Space Science and Technology Network (REDCyTE) has designed glass-ceramic coatings reinforced with nanometric ceramic particles of Al2 O3 , SiO2 , and TiO2 , to obtain a multilayer system that is highly reflective and thin and homogeneous through the sol-gel. **Figure 6** shows images of scanning electron microscopy (SEM) in topography mode, in which the surfaces of the coatings are shown. In the same figure, SEM images are shown in chemical mapping mode in which the uniform dispersion of the ceramic nanoparticles immersed in the vitreous matrix is shown.

#### **5. Challenges in medium and long term**

In 2010, the Mexican Space Agency was born [29]. It followed a few decades of academic work and the need for an institution to formalize and federate projects toward the development of national space technologies. In 2015, a specific fund allocated to space technology was created by the government through the Federal Funding Agency (CONACYT) [30]. A few projects have been so far benefitting from this fund, and among them, one project dedicated to suborbital flights using stratospheric balloon was funded (ATON) [30]. Unfortunately, it was the only project funded to promote suborbital flights, and it is clear that more incentives are needed to allow the development of this sector in a near future and also to sustain the development of the space technologies in Mexico. In medium term, one can prospect for a need to increase the gondola size to provide services to larger payloads with greater mass. Typical light payloads range from less of a few kilograms to tens of kilograms. To lower the cost of flights and provide a service throughout the year, it is essential to be able to carry up to a ton if not a few hundreds of kilograms. At those masses, the risk associated with structural damages and health safety that such a gondola can create in semi-urban centers is too high. Therefore, flights over the oceans or the deserts, both available in Mexico, should be planned under the resources of the marine, air force, and the government for proper permits. It seems unlikely that flights over the jungle as one could plan in south of Mexico is realistic, due to the complexity to recover a gondola in a harsh, difficult-to-access environment with dense vegetation. States such as Sonora, Chihuahua, and Durango for desert lands or Baja California, Sinaloa, and Veracruz for oceans can be explored as possible sites for launches. Additionally, specific balloons of larger flight durations should be investigated and tried out in conjunctions with large masses of a few hundreds of kilograms to a ton.

Besides the practicability of the launch, tracking and recovery of the gondola, and the payload, it is important that the structure, which the gondola offers, houses onboard electronics and power services. The system should be able to monitor the environment and itself and to supply the payloads with the necessary power for their operation. Multidisciplinary teams of engineers in mechanics, aerospace technology, telecommunication, electrical and electronic systems, and aeronautics should be part of this effort, and likely a national center for aerospace technology can be the right place to gather those specialties. There is an important risk that allocating resources to small individual projects may not be fructuous on the long term and it may be sensible to lay out one or a couple of projects of national interest that can lead the effort with a sustainable funding.

#### **6. Conclusions**

(6) good adhesion to the metallic substrate, and (7) low heating rate during the sintering process [24, 25]. The materials that can be used as glass-ceramic coatings is very limited; so far, only some materials have been found that meet these requirements [26]: aluminum oxide

for use as thermal control surfaces in aircraft because lightweight coatings with good adhesive properties can be obtained under thermal shock conditions [27]. This is why it could be considered that glass-ceramic coatings applied in satellite systems could resist ultraviolet, cos-

Therefore, the team of Coatings Resistant to Thermal Changes and Cosmic Radiation (CRTCR) from Space Science and Technology Network (REDCyTE) has designed glass-ceramic coat-

layer system that is highly reflective and thin and homogeneous through the sol-gel. **Figure 6** shows images of scanning electron microscopy (SEM) in topography mode, in which the surfaces of the coatings are shown. In the same figure, SEM images are shown in chemical mapping mode in which the uniform dispersion of the ceramic nanoparticles immersed in the

In 2010, the Mexican Space Agency was born [29]. It followed a few decades of academic work and the need for an institution to formalize and federate projects toward the development of national space technologies. In 2015, a specific fund allocated to space technology was created by the government through the Federal Funding Agency (CONACYT) [30]. A few projects have been so far benefitting from this fund, and among them, one project dedicated to suborbital flights using stratospheric balloon was funded (ATON) [30]. Unfortunately, it was the only project funded to promote suborbital flights, and it is clear that more incentives are needed to allow the development of this sector in a near future and also to sustain the development of the space technologies in Mexico. In medium term, one can prospect for a need to increase the gondola size to provide services to larger payloads with greater mass. Typical light payloads range from less of a few kilograms to tens of kilograms. To lower the cost of flights and provide a service throughout the year, it is essential to be able to carry up to a ton if not a few hundreds of kilograms. At those masses, the risk associated with structural damages and health safety that such a gondola can create in semi-urban centers is too high. Therefore, flights over the oceans or the deserts, both available in Mexico, should be planned under the resources of the marine, air force, and the government for proper permits. It seems unlikely that flights over the jungle as one could plan in south of Mexico is realistic, due to the complexity to recover a gondola in a harsh, difficult-to-access environment with dense vegetation. States such as Sonora, Chihuahua, and Durango for desert lands or Baja California, Sinaloa, and Veracruz for oceans can be explored as possible sites for launches. Additionally, specific

O3 5SiO2

> O3 , SiO2

), zirconium oxide (ZrO2

). These coatings have been initially proposed

, and TiO2

), and

, to obtain a multi-

), cordierite (2MgO-2Al2


mic, and high-energy particles over a wide temperature range [28].

ings reinforced with nanometric ceramic particles of Al2

**5. Challenges in medium and long term**

(Al2 O3

192 Space Flight

), mullite (3Al2

vitreous matrix is shown.

O3 -2SiO2

zirconia stabilized with yttrium (ZrO2

Mexico has been involved into suborbital flights since the past two decades through amateur impulse. Lately, the academic sector has been starting to promote the use of stratospheric sounding balloons to provide the scientific and technological communities and inexpensive, easy-to-use facility to test their aerospace technologies. A few suborbital flights over the country have been carried out by different institutions with a large participation of students. Among the different projects, CSM is the only one involving mostly researchers and engineers from the academic and private sectors with the objective to professionalize this area of knowledge following the example of leading countries. This platform has provided two flights of 2-h duration for altitudes over 30 km to three clients. It is expected to provide yearly flights to future clients accommodating payloads of larger masses.

Overall, suborbital flights are the first step for Mexico to get involved in the aerospace development. Increasing funding, engineering schools, and national laboratories dedicated to aerospace studies and technological and scientific projects with the participation of the government, the military and the private sector is considered mandatory to achieve a significant contribution to the field in Latin America.

#### **Acknowledgements**

The CSM team would like to thank the Institute of Engineering and the Faculty of Engineering of the UNAM for sponsoring CSM.

They also thank the DGAPA for the support through the project IN113315. Finally, the authors really appreciate the participation and enthusiasm of Eduardo Amaro Calderon, Diego Dominguez Baez and Alfredo Sanchez Labra.

### **Author details**

Barbara Bermudez-Reyes1 \*, Frederic Trillaud2 , Fernando Velazquez-Villegas<sup>3</sup> , Jonathan Remba-Uribe<sup>4</sup> , Ana M. Arizmendi-Morquecho5 , Alberto Caballero-Ruíz<sup>6</sup> , Mario A. Mendoza-Barcenas<sup>7</sup> , Rafael Prieto-Melendez<sup>7</sup> , Leopoldo Ruiz-Huerta<sup>6</sup> and Lauro Santiago-Cruz2

\*Address all correspondence to: barbara.bermudezry@uanl.edu.mx

1 Faculty of Mechanical and Electrical Engineering, Autonomous University of Nuevo Leon, Nuevo León, México

www.webcitation.org/618QHms8h?url=http://www.fai.org/astronautics/100km.asp

Suborbital Flight: An Affordable and Feasible Option for Mexican Aerospace Development

http://dx.doi.org/10.5772/intechopen.73859

195

[8] Advanced Space Transportation Program: Paving the Highway to Space, Marshall Space Flight Center, NASA. Available from: https://www.nasa.gov/centers/marshall/news/

[9] U. S. Centennial Commission. Geometric altitude vs. temperature, pressure, density, and the speed of sound derived from the 1962 U.S. Standard Atmosphere. http://www.centennialofflight.net/essay/Theories\_of\_Flight/atmosphere/TH1G1.htm [Accessed: 12/07/2017]

[10] Solana J. Radio Club satellite A.C. BOLETIN XE1RCS. Historical Achieves (2008-2015).

[11] Club de Radio Amateur del Estado de Guanajuato A.C.-XE1CRG. Historical Archives

[12] CONACYT Press. Pixqui: plataforma mexicana de pruebas satelitales, Aug 18 2015.

[13] Agencia Espacial Mexicana (AEM) Antecedentes de la AEM. Available from: www.gob. mx/aem/acciones-y-programas/antecedentes-de-la-aem [Accessed: Apr 12, 2017]

[14] Mendoza Bárcenas MA, Prieto Meléndez R, Santiago Cruz L, Trillaud F, Espinosa Calderón A, Herraiz Sarachaga M, Velázquez Villegas F. Módulo experimental de carga útil "SADM-1" para fines de exploración atmosférica. Sociedad Mexicana de Instrumentacion. In: XXXII Congreso de Instrumentación. Guerrero, Mexico; October

[15] Herrera-Arroyave JE, Bermúdez-Reyes B, Ferrer-Pérez JA, Colín A. CubeSat system structural design. In: 67th International Astronautical Congress. Guadalajara, Mexico;

[16] Sebastian Rosas Contreras. Diseño de la Estructura Mecánica de una Carga de Servicio Ligera de Globos Estratosféricos Nacionales [undergraduate thesis]. UNAM; Oct 1, 2016

[17] Singer Genovese R, Trillaud F, Velazquez Villegas F, Santiago Cruz L, Remba J. Model and simulations of high altitude sounding balloons: Dynamics, stress-strain and thermal analysis. In: 67th International Astronautical Congress. Guadalajara, Mexico; Sep 2016.

[18] DiCarlo JA, Yun H-M. Non-oxide (Silicon Carbide) Fibers. In: Bansal NP, editor. Handbook of Ceramics Composites, NASA Glenn Research Center. United States of

[19] ECSS Secretariat ESA-ESTEC. Space Engineering. Structural Materials Handbook. Part 4: Integrity control, verification guidelines and manufacturing. European Cooperation for Space Standardization. Noordwijk, The Netherlands: Requirements & Standards

[20] ASM International, editor. Metals Handbook. Vol. 2. Properties and Selection: Nonferrous Alloys and Special Purpose Materials. 10 ed. Metals Park, OH: ASM; 1999

America: Kluwer Academic Publisher; 2005. pp. 33-38

Available from: www.conacytprensa.mx [Accessed: Aug 11, 2017]

background/facts/astp.html [Accessed: Dec 11, 2017]

(2009-2015). Guanajuato, Mexico; 2017

[accessed: 12/10/2017]

Mexico; 2017

2017

pp. 1-11

September 2016. pp. 1-5

Division; 2011. pp. 219-230

2 Institute of Engineering, National Autonomous University of Mexico, México City, México

3 Center of Advanced Engineering, Faculty of Engineering, National Autonomous University of Mexico, México City, México

4 Remtronic Telecommunications, Guanajuato, México

5 Center of Research in Advanced Materials-Monterrey, Nuevo León, México

6 National Laboratory of Additive Manufacturing, 3D digitalization and Computed Tomography, Center of Applied Sciences and Technological Development, National Autonomous University of Mexico, México City, México

7 Aerospace Development Center, National Polytechnic Institute, México City, México

#### **References**


www.webcitation.org/618QHms8h?url=http://www.fai.org/astronautics/100km.asp [accessed: 12/10/2017]

[8] Advanced Space Transportation Program: Paving the Highway to Space, Marshall Space Flight Center, NASA. Available from: https://www.nasa.gov/centers/marshall/news/ background/facts/astp.html [Accessed: Dec 11, 2017]

**Author details**

194 Space Flight

Barbara Bermudez-Reyes1

Mario A. Mendoza-Barcenas<sup>7</sup>

University of Mexico, México City, México

4 Remtronic Telecommunications, Guanajuato, México

Autonomous University of Mexico, México City, México

Conference 2015. Montana, USA: IEEE

Jonathan Remba-Uribe<sup>4</sup>

Lauro Santiago-Cruz2

Nuevo León, México

**References**

Inc.; 2013

\*, Frederic Trillaud2

\*Address all correspondence to: barbara.bermudezry@uanl.edu.mx

, Ana M. Arizmendi-Morquecho5

, Rafael Prieto-Melendez<sup>7</sup>

3 Center of Advanced Engineering, Faculty of Engineering, National Autonomous

6 National Laboratory of Additive Manufacturing, 3D digitalization and Computed Tomography, Center of Applied Sciences and Technological Development, National

7 Aerospace Development Center, National Polytechnic Institute, México City, México

[1] Welti CR. Satellite Basics for Everyone. Blomington: iUniverse, Inc.; 2012. pp. 10-16

[2] Thom Stone, Marcus Murbach, Richard Alema, Ray Gilstrap. SOAREX- Suborbital Experiments 2015 – A New Paradigm for Small Spacecraft Communication. In: Aerospace

[3] Sako N, Tsuda Y, Ota S, Eishima T, Yamamoto T, Ikeda I, Li H, Yamamoto H, Tanaka H, Tanaka A, Nakasuka S. Cansat suborbital launch experiment-university educational space program using can sized Pico-satellite. Acta Astronautica. 2001;**48**(5-12):767-766

[4] Briess K. Space design process. In: Ley W, Wittmann K, Hallmann W, editors. Handbook of Space Technology. 1st ed. United Kingdom: Jonh Wiley and Sons, Ltd; 2009

[5] Selva D, Dingwall B, Altunc S. A concept for an Agile Mission Development Facility for Cubesat and suborbital Missions. In: Aerospace Conference. Montana, USA: IEEE; 2016

[6] Antunes S. DIY Instruments for Amateur Space. 1st ed. California, USA: O'Reilly Media

[7] Sanz Fernández de Córdoba S. Karman separation line, used as the boundary separating Aeronautics and Astronautics. Fédération Aéronautique Internationale; 2004. https://

5 Center of Research in Advanced Materials-Monterrey, Nuevo León, México

1 Faculty of Mechanical and Electrical Engineering, Autonomous University of Nuevo Leon,

2 Institute of Engineering, National Autonomous University of Mexico, México City, México

, Fernando Velazquez-Villegas<sup>3</sup>

, Alberto Caballero-Ruíz<sup>6</sup>

, Leopoldo Ruiz-Huerta<sup>6</sup>

,

,

and


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**Chapter 11**

Provisional chapter

**Cost-Effective Platforms for Near-Space Research and**

DOI: 10.5772/intechopen.72168

Cost-Effective Platforms for Near-Space Research and

High-altitude balloons (HABs) are commonly used for atmospheric research. In recent years, newly developed platforms and instruments allow to measure position, temperature, radiation, humidity and gas profile in the troposphere and stratosphere. However, current platforms, such as radiosonde, have limited bandwidth and relatively small number of possible sensors on board. Furthermore, all the measuring instruments carried on board the balloon cannot be reused since most of the times the radiosonde cannot be retrieved. In this chapter, we present a generic near-space research platform based on an improved radio frequency (RF) communication, an advanced set of sensors that might also include a return-to-home (RTH) micro-UAV. We present the overall structure of an advanced HAB payload, which is equipped with a low-cost sophisticated set of sensors along with HD camera system, which weight less than 300 g. The payload is tied to a weather balloon with a smart autonomous release mechanism and two-way RF telemetry channel (LoRa or Iridium communication). The payload can be released from the balloon at any given time or position, allowing it to fall at a predicted area. In case the payload is attached to a micro UAV, it can return autonomously by multioptional smart decline to a predefined location using a built-in autopilot. The suggested new strategy is presented

Keywords: atmospheric and climate research, testing space components, near-space experiments, autonomous near-space flights, high altitude weather balloons, long-range

> © The Author(s). Licensee InTech. 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 eproduction in any medium, provided the original work is properly cited.

© 2018 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.

**Experiments**

Experiments

Abstract

RF communication

Kobi Gozlan, Yuval Reuveni, Kfir Cohen,

Kobi Gozlan, Yuval Reuveni, Kfir Cohen,

Additional information is available at the end of the chapter

Additional information is available at the end of the chapter

using several case studies and field experiments.

Boaz Ben-Moshe and Eyal Berliner

Boaz Ben-Moshe and Eyal Berliner

http://dx.doi.org/10.5772/intechopen.72168


## **Cost-Effective Platforms for Near-Space Research and Experiments** Provisional chapter Cost-Effective Platforms for Near-Space Research and

DOI: 10.5772/intechopen.72168

Kobi Gozlan, Yuval Reuveni, Kfir Cohen,

Boaz Ben-Moshe and Eyal Berliner Kobi Gozlan, Yuval Reuveni, Kfir Cohen,

Additional information is available at the end of the chapter Boaz Ben-Moshe and Eyal Berliner

http://dx.doi.org/10.5772/intechopen.72168 Additional information is available at the end of the chapter

#### Abstract

Experiments

[21] Hussey R, Wilson J, editors. Light Alloys Directory and Handbook. London: Chapman

[22] Berndt CC, Brindley W, Goland AN, Herman H, Houck DL, Jones K, Miller RA, Neiser R, Riggs W, Sampath S, Smith M, Spanne P. Journal of Thermal Science and Technology.

[23] Cortese B, Caschera D, de Caro T, Ingo GM. Micro-chemical and -morphological features of heat treated plasma sprayed zirconia-based thermal barrier coatings. Thin Solid

[24] Cernuschi F, Bianchi P, Leoni M, Scardi P. Thermal diffusivity/microstructure relationship in Y-PSZ thermal barrier coatings. Journal of Thermal Spray Technology.

[25] Vassen R, Tietz F, Kerkhoff G, Stoever D. New materials for advanced thermal barrier coatings. In: Lecomte-Beckers J, Schuber F, Ennis PJ, editors. Proceedings of the 6th Liége Conference on Materials for Advanced Power Engineering. Belgium: Universite

[26] Cao XQ, Vassen R, Stoever D. Ceramic materials for termal barrier coatings. Sciencie

[27] Singletary BJ. Ceramics, Refractories and Glasses. In: Rittenhouse JB, Singletary JB. editors. Space Materials Handbook. 3rd ed. United States of America: NASA SP-30 51;

[28] Nguyen CH, Chandrashekhara K, Birman V. Multifunctional thermal barrier coating in aerospace sándwich panels. Mechanics Research Communications. 2012;**39**:35-43 [29] Consejo Nacional de Ciencia y Tecnología. Comunicado 64/14, Invertirán CONACYT y Agencia Espacial Mexicana más de 30 MDP en investigación e innovación aeroespacial. Ciudad de México, Oct 8, 2014. Available from: https://www.conacyt.gob.mx/index.php/ comunicacion/comunicados-prensa/381-invertiran-conacyt-y-agencia-espacial-mexicanamas-de-30-mdp-en-investigacion-e-innovacion-aeroespacial [Accessed: Dec 16, 2017] [30] Consejo Nacional de Ciencia y Tecnología. Agencia Informativa, Montserrat Muñoz. Misión suborbital ATON – Ulises I: Ciencia y arte en la FIL. Dec 7, 2015. Available from: http://www.conacytprensa.mx/index.php/ciencia/universo/4619-prioritaria-mision-

Direct, Journal of European Ceramic Society. 2004;**24**:1-10

suborbital-aton-ulises-i-ciencia-y-arte [Accessed: Apr 12, 2017]

and Hall; 1998

Films. 2013;**549**:321-329

1999;**8**(1):102-109

de Liege; Nov 1998

1969. pp. 120-130

1992;**1**:1

196 Space Flight

High-altitude balloons (HABs) are commonly used for atmospheric research. In recent years, newly developed platforms and instruments allow to measure position, temperature, radiation, humidity and gas profile in the troposphere and stratosphere. However, current platforms, such as radiosonde, have limited bandwidth and relatively small number of possible sensors on board. Furthermore, all the measuring instruments carried on board the balloon cannot be reused since most of the times the radiosonde cannot be retrieved. In this chapter, we present a generic near-space research platform based on an improved radio frequency (RF) communication, an advanced set of sensors that might also include a return-to-home (RTH) micro-UAV. We present the overall structure of an advanced HAB payload, which is equipped with a low-cost sophisticated set of sensors along with HD camera system, which weight less than 300 g. The payload is tied to a weather balloon with a smart autonomous release mechanism and two-way RF telemetry channel (LoRa or Iridium communication). The payload can be released from the balloon at any given time or position, allowing it to fall at a predicted area. In case the payload is attached to a micro UAV, it can return autonomously by multioptional smart decline to a predefined location using a built-in autopilot. The suggested new strategy is presented using several case studies and field experiments.

Keywords: atmospheric and climate research, testing space components, near-space experiments, autonomous near-space flights, high altitude weather balloons, long-range RF communication

© 2018 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.

© The Author(s). Licensee InTech. 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 eproduction in any medium, provided the original work is properly cited.

### 1. Introduction

#### 1.1. Motivation

Traditionally, the space industry was mainly founded by governmental or military organizations. Yet, in recent years, the "new-space" environment attracts several major private companies such as Google, Facebook and OneWeb, each having a large-scale communication project involving global coverage using low earth orbit (LEO) nano-satellite swarm.

absorption and emission of radiation with respect to the recently revealed interesting insight regarding the radiation obsorbtion/emission dependency with altitude [2]. In addition, due to the fact that our plant's atmosphere is continuously bombarded by energetic particles, mainly galactic cosmic rays (GCR), along with sporadic space weather events, additional particles are introduced into the stratosphere and troposphere [7]. Regardless of this recurring impact, the effects of energetic particles in the troposphere and lower stratosphere are still inadequately understood. There are numerous mechanisms for explaining how weather and climate could potentially be modulated [8], but the majority of energetic particle effects in the lower atmosphere are linked to their potential for ionizing the surrounding air. The formed ions can accumulate on cloud tops, contributing to the microphysics [9], may play a key role in the formation of aerosol (e.g., [10]). In addition, atmospheric ions can absorb directly infrared radiation (IR) [11, 12], and high-energy particles are also presumed to impact lightning rates [13]. Above land and within the boundary layer (~few hundred meters), the atmosphere is mainly ionized by the radiation emitted from radioactive isotopes decay in the Earth's crust [3]. Hess [14] postulated that the ionization profile in the atmosphere should decrease with altitude due to the fact that the radioactive element source is located near the surface. However, after conducting balloon measurements, Hess discovered that the ionization increased at altitudes above 10 km and claimed that it is caused by GCR source. He also determined that penetration depth of these particles depends on the energy spectrum of the incoming radiation [14]. Two decades later, Regener extended Hess' experiments using HAB, measuring ionization rates up to altitudes of 20 km [15]. They discovered that cosmic ray ionization reaches its maximum value between altitudes of 17–24 km and is known as the Regener-Pfotzer maximum (RP max). The Pfotzer Maximum, which is also geomagneticlatitude dependent [16], formed within the tropopause layer below the stratosphere where primary particles (pions and hadrons) decrease and secondary particles (muons) increase [17]. This is a major source of ionization in the Earth's atmosphere. The establishment of an electromagnetic-muon stream results in ambient air ionization during the release of primary energies by the excitation of air molecules deeper in the atmosphere [17]. During this stream, a portion of the primary particles reach the ground as high-energy secondary particles [18]. The electromagnetic field also interacts with incoming particles, as the sun's solar radiation penetrates the atmosphere. This mixing is directly associated with the pressure decrease as

Cost-Effective Platforms for Near-Space Research and Experiments

http://dx.doi.org/10.5772/intechopen.72168

199

the differential absorption rate within tropopause heights varies [17].

can "land" in the middle of the ocean or in a high peak of a mountain.

The necessity for developing new techniques and platforms for measuring and identifying energetic ionizing radiation in the atmosphere becomes vital. However, despite numerical model simulations for estimating flight trajectories, high-precision global positioning system (GPS) technology and the relatively slow balloon descent, recovering high-cost payload yet remains challenging, difficult and time-consuming, specifically around mountains or coastal areas [2].

Retrieving the payload enables us to acquire all the recorded data during the flight and that we were not been able to send using wireless communications. This is easier said than done and in practice, HAB payloads are not expected to be retrieved. For retrieving the payload, one should know the exact landing location of the payload, and more important, one must have access to that location. Thus, knowing the payload's exact landing location is not enough as it

The vision of having a reliable and affordable global network, which can be accessed from any location on Earth at any given time, is a challenging scientific task, which attracts both industrial and academic efforts during the last few decades. Currently, the majority of all proposed solutions are based on a network of numerous LEO nano-satellites, which will establish a global network using radio frequency (RF) communication data received on Earth. Major companies such as Google, Qualcomm, Facebook and SpaceX have each invested in similar projects, commonly referred as new-space and near-space projects. OneWeb is one example for such initiative project involving a large constellation of LEO satellites. Other projects such as Google's Loon or Facebook's Aquila Drone are not directly focused on satellite constellations but on near-space massive constellation of drones or balloons. The new-space industry includes various small to medium size companies, which are currently developing products for the near-space environment, e.g., Planet Labs and Spire companies are two examples for such effort, which is focused on global imaging and IoT.

Constructing a cost-effective global network requires the use of low-cost electronics, unlike the traditional space industry, which uses dedicated expensive hardware. In order to perform a "space-qualified" testing platform on such components, a flexible modulated testing platform is needed. In this work, we present a new generic methodology for performing near-space experiments based on advanced low-cost payload, which is tied to a weather balloon. The suggested strategy is based on more than dozen balloon-launch experiments encompassing a large number of components (electronics and mechanics), which were tested at 10–30 km heights.

#### 1.2. Related scientific work

High-altitude balloon (HAB) platforms have been used for direct atmospheric measurements for more than a century [1]. Measuring devices, which send data from HAB to a base-station located on the ground, using pocket-sized radio frequency (RF) transmitters and are widely known as radiosondes, were first invented by the French scientist Robert Bureau in 1929 [2]. Recently, HAB platforms have started to gain the ability of measuring, recording and transmitting other sources of data from a vast variety of instruments, substantially increasing HAB payload capabilities [3]. Furthermore, the increasing supporting evidence for climate change along with the understanding of real-time atmospheric composition measurements, both in the upper troposphere and lower stratosphere, is a key feature for studying radiative effects in our plant's climate system [2, 4], emphasizing the need for developing upper-air climate observation platforms [5, 6].

Although the main objective for HAB measurements is to monitor changes in temperature and water vapor vertical profiles in the troposphere and stratosphere, several new upper-air radiation profile measurements indicate supplemental valuable information regarding atmospheric absorption and emission of radiation with respect to the recently revealed interesting insight regarding the radiation obsorbtion/emission dependency with altitude [2]. In addition, due to the fact that our plant's atmosphere is continuously bombarded by energetic particles, mainly galactic cosmic rays (GCR), along with sporadic space weather events, additional particles are introduced into the stratosphere and troposphere [7]. Regardless of this recurring impact, the effects of energetic particles in the troposphere and lower stratosphere are still inadequately understood. There are numerous mechanisms for explaining how weather and climate could potentially be modulated [8], but the majority of energetic particle effects in the lower atmosphere are linked to their potential for ionizing the surrounding air. The formed ions can accumulate on cloud tops, contributing to the microphysics [9], may play a key role in the formation of aerosol (e.g., [10]). In addition, atmospheric ions can absorb directly infrared radiation (IR) [11, 12], and high-energy particles are also presumed to impact lightning rates [13].

1. Introduction

1.2. Related scientific work

Traditionally, the space industry was mainly founded by governmental or military organizations. Yet, in recent years, the "new-space" environment attracts several major private companies such as Google, Facebook and OneWeb, each having a large-scale communication project

The vision of having a reliable and affordable global network, which can be accessed from any location on Earth at any given time, is a challenging scientific task, which attracts both industrial and academic efforts during the last few decades. Currently, the majority of all proposed solutions are based on a network of numerous LEO nano-satellites, which will establish a global network using radio frequency (RF) communication data received on Earth. Major companies such as Google, Qualcomm, Facebook and SpaceX have each invested in similar projects, commonly referred as new-space and near-space projects. OneWeb is one example for such initiative project involving a large constellation of LEO satellites. Other projects such as Google's Loon or Facebook's Aquila Drone are not directly focused on satellite constellations but on near-space massive constellation of drones or balloons. The new-space industry includes various small to medium size companies, which are currently developing products for the near-space environment, e.g., Planet Labs and Spire companies are two

Constructing a cost-effective global network requires the use of low-cost electronics, unlike the traditional space industry, which uses dedicated expensive hardware. In order to perform a "space-qualified" testing platform on such components, a flexible modulated testing platform is needed. In this work, we present a new generic methodology for performing near-space experiments based on advanced low-cost payload, which is tied to a weather balloon. The suggested strategy is based on more than dozen balloon-launch experiments encompassing a large number

High-altitude balloon (HAB) platforms have been used for direct atmospheric measurements for more than a century [1]. Measuring devices, which send data from HAB to a base-station located on the ground, using pocket-sized radio frequency (RF) transmitters and are widely known as radiosondes, were first invented by the French scientist Robert Bureau in 1929 [2]. Recently, HAB platforms have started to gain the ability of measuring, recording and transmitting other sources of data from a vast variety of instruments, substantially increasing HAB payload capabilities [3]. Furthermore, the increasing supporting evidence for climate change along with the understanding of real-time atmospheric composition measurements, both in the upper troposphere and lower stratosphere, is a key feature for studying radiative effects in our plant's climate system [2, 4], emphasizing the need for developing upper-air climate observation platforms [5, 6].

Although the main objective for HAB measurements is to monitor changes in temperature and water vapor vertical profiles in the troposphere and stratosphere, several new upper-air radiation profile measurements indicate supplemental valuable information regarding atmospheric

of components (electronics and mechanics), which were tested at 10–30 km heights.

involving global coverage using low earth orbit (LEO) nano-satellite swarm.

examples for such effort, which is focused on global imaging and IoT.

1.1. Motivation

198 Space Flight

Above land and within the boundary layer (~few hundred meters), the atmosphere is mainly ionized by the radiation emitted from radioactive isotopes decay in the Earth's crust [3]. Hess [14] postulated that the ionization profile in the atmosphere should decrease with altitude due to the fact that the radioactive element source is located near the surface. However, after conducting balloon measurements, Hess discovered that the ionization increased at altitudes above 10 km and claimed that it is caused by GCR source. He also determined that penetration depth of these particles depends on the energy spectrum of the incoming radiation [14]. Two decades later, Regener extended Hess' experiments using HAB, measuring ionization rates up to altitudes of 20 km [15]. They discovered that cosmic ray ionization reaches its maximum value between altitudes of 17–24 km and is known as the Regener-Pfotzer maximum (RP max). The Pfotzer Maximum, which is also geomagneticlatitude dependent [16], formed within the tropopause layer below the stratosphere where primary particles (pions and hadrons) decrease and secondary particles (muons) increase [17]. This is a major source of ionization in the Earth's atmosphere. The establishment of an electromagnetic-muon stream results in ambient air ionization during the release of primary energies by the excitation of air molecules deeper in the atmosphere [17]. During this stream, a portion of the primary particles reach the ground as high-energy secondary particles [18]. The electromagnetic field also interacts with incoming particles, as the sun's solar radiation penetrates the atmosphere. This mixing is directly associated with the pressure decrease as the differential absorption rate within tropopause heights varies [17].

The necessity for developing new techniques and platforms for measuring and identifying energetic ionizing radiation in the atmosphere becomes vital. However, despite numerical model simulations for estimating flight trajectories, high-precision global positioning system (GPS) technology and the relatively slow balloon descent, recovering high-cost payload yet remains challenging, difficult and time-consuming, specifically around mountains or coastal areas [2].

Retrieving the payload enables us to acquire all the recorded data during the flight and that we were not been able to send using wireless communications. This is easier said than done and in practice, HAB payloads are not expected to be retrieved. For retrieving the payload, one should know the exact landing location of the payload, and more important, one must have access to that location. Thus, knowing the payload's exact landing location is not enough as it can "land" in the middle of the ocean or in a high peak of a mountain.

The rest of the paper is structured as follows: in Section 2, we survey the basic principle of flying a high-altitude balloon (HAB). In Section 3, we present design for a disposable costeffective payload for low-bandwidth applications, which provide the base platform for our experiments. In Section 4, we cover the HAB payload components, power supply behavior, thermal design and pre-flight tests. In Section 5, we present our investigations of long-range communications for low- and high-bandwidth applications. In Section 5, we present our own setup of a near-space return-to-home (RTH) micro-UAV for retrieving the payload with its recorded data. Finally, we discuss our efforts and future work.

Assuming 2000 l of helium were used, the 1000-g balloon should gain about 1000 g lift at 1 atm. Assuming a 500 g payload is attached to the balloon, one can expect an overall lift of

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Given the desired requirements for the experiment, e.g., max altitude, payload weight and required floating duration, one can adjust the amount of helium in the balloon accordingly.

As the balloon inclines, its surrounding air pressure decreases. Table 2 presents the expected

Balloon type Volume (L) Payload (g) Neck (g) Burst altitude (m) Ascent rate (m/s) Duration (m)

Table 1. HAB parameters: Few examples of the lift, duration and burst altitude with respect to the balloon type, payload

Altitude (m, 15 cel) Air pressure (atm) Balloon volume (L) Balloon diameter (m)

0 1.0 1000 1.24 2361 0.75 1333 1.36 5477 0.5 2000 1.56 10,278 0.25 4000 1.97 16,096 0.1 10,000 2.67 32,230 0.01 100,000 5.76 48,330 0.001 1,000,000 12.41

300 Kaymont 600 200 316 27,890 2.96 157 600 Kaymont 1500 200 1453 29,280 5.95 82 1000 Kaymont 2000 500 1053 35,070 3.95 148 1000 Kaymont 3000 1000 2080 32,135 4.84 111 1000 Kaymont 3000 1500 2080 32,135 3.54 151 1000 Kaymont 4000 2000 3106 30,053 4.44 113

500 g. In Figure 2, a basic calculation of the expected balloon parameters is presented.

Table 1 presents few examples for such adjustment.

Figure 2. An HAB's lift and burst calculator, from: http://habhub.org/calc/.

mass and amount of helium.

Table 2. Expected air pressure at a given altitude.

air pressure with respect to the balloon height.

#### 2. Preliminaries: basic principle of high-altitude balloon

In general, high-altitude balloon (HAB) is composed of the following components:


Consider a balloon with a self-weight of 1000 g, about 1000 l of helium is needed in order to allow the balloon to start floating (for each m3 of helium—one can expect a lift of 1000 g—1 kg).

Figure 1. Launching a HAB—yet another day at the office.

Assuming 2000 l of helium were used, the 1000-g balloon should gain about 1000 g lift at 1 atm. Assuming a 500 g payload is attached to the balloon, one can expect an overall lift of 500 g. In Figure 2, a basic calculation of the expected balloon parameters is presented.

Given the desired requirements for the experiment, e.g., max altitude, payload weight and required floating duration, one can adjust the amount of helium in the balloon accordingly. Table 1 presents few examples for such adjustment.

As the balloon inclines, its surrounding air pressure decreases. Table 2 presents the expected air pressure with respect to the balloon height.

Figure 2. An HAB's lift and burst calculator, from: http://habhub.org/calc/.

The rest of the paper is structured as follows: in Section 2, we survey the basic principle of flying a high-altitude balloon (HAB). In Section 3, we present design for a disposable costeffective payload for low-bandwidth applications, which provide the base platform for our experiments. In Section 4, we cover the HAB payload components, power supply behavior, thermal design and pre-flight tests. In Section 5, we present our investigations of long-range communications for low- and high-bandwidth applications. In Section 5, we present our own setup of a near-space return-to-home (RTH) micro-UAV for retrieving the payload with its

recorded data. Finally, we discuss our efforts and future work.

cream boxes).

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Figure 1. Launching a HAB—yet another day at the office.

2. Preliminaries: basic principle of high-altitude balloon

In general, high-altitude balloon (HAB) is composed of the following components:

inflated with helium, common use for HAB may weight 100–1200 g.

• A latex balloon—comes in a wide range of weights, which basically reflects its ability to be

• A payload—which includes all the necessary components for conducting the experiment and retrieving the data. In Figure 1, two payloads are connected (black and white ice-

• A ground station (GS)—commonly includes an RF receiver. In Figure 1, the GS also includes

Consider a balloon with a self-weight of 1000 g, about 1000 l of helium is needed in order to allow the balloon to start floating (for each m3 of helium—one can expect a lift of 1000 g—1 kg).

a robotic telescope and transmitter to control the payload detaching process.


Table 1. HAB parameters: Few examples of the lift, duration and burst altitude with respect to the balloon type, payload mass and amount of helium.


Table 2. Expected air pressure at a given altitude.

#### 3. Long-lasting "floating" balloon

In a typical HAB configuration using latex balloon, the balloon will ascend and expand as the air pressure decreases with height due to the thin atmosphere. At a certain point, it will inflate up to its elastic point, explode and fall. This means that if we can make the balloon float in a relatively constant altitude, we can extend its lifespan and endurance. Moreover, fixing the balloon at high altitudes could also enable to test any desired hardware under near-space conditions.

Google's "loon project" is a good example for an HAB setting that is capable of floating up in the atmosphere for a long duration. However, such settings are expensive and complicated, thus they are not a practical solution for scientific researchers.

Our approach for "fixing" the balloon's altitude was directed toward a simple constriction, i.e., a main latex balloon and a cluster of foil balloons. Foil balloons are not elastic and cannot expand, thus their volume can be approximated as constant. This means that as the outer pressure drops due to the thin atmosphere, its upthrust force will weaken and might even change its direction as dictated by the buoyancy force equation:

$$F\_B = \left(\rho\_{air} - \rho\_{gas}\right)gV \tag{1}$$

• Main payload (above) 500-g: Iridium transceiver, GPS, Solar panel, battery, Geiger counter and an autonomous release mechanism for a secondary payload.

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Figure 4. The retrieved payloads from the above experiment were found on a distant field, about 100 km from the point of launch. The crashing location was transmitted by the iridium communication system after the crash. In the picture, the

• Secondary payload (lower) 400-g: A long-range HD video streaming system based on Wi-Fi and a directional antenna pointing down (14 dBi flat panel antenna). In this test setting, each foil balloon has a net weight of approximately 0 g on ground level, while the expected weight at an altitude of 10 km is about 70 g. The overall setting provides a lift force of about 400 g on ground level. When reaching to 9–10 km height the system's net lift

force should be about 0, making the system relatively altitude-stationary (Figure 4).

In this subsection, we cover the HAB payload components. First, we present the common needed and used sensors in "near-space" experiments. Then, a brief discussion on energy and thermal design is presented—followed by a discussion of how to test a potential payload (on

• GNSS (e.g., GPS): Global Navigation Satellite Systems refer to a positioning sensor commonly used for computing the 3D position in a typical horizontal accuracy of 2–3 m in the open sky (the vertical accuracy is often not as accurate as the horizontal – errors of 10–20 m

4. Sensors, energy and thermal design

upper box is the secondary payload and the main is the lower box.

4.1. Sensors

the ground) for its ability to operate under near-space conditions.

where rair is the surrounding air density, rgas is the helium density, V is the balloon volume and g is the gravitational force. We present a basic example of our current test setting design (Figure 3):


Figure 3. A long-lasting HAB experiment. This setting retained its floating state for about 2.5 h before the main balloon exploded.

Figure 4. The retrieved payloads from the above experiment were found on a distant field, about 100 km from the point of launch. The crashing location was transmitted by the iridium communication system after the crash. In the picture, the upper box is the secondary payload and the main is the lower box.


In this test setting, each foil balloon has a net weight of approximately 0 g on ground level, while the expected weight at an altitude of 10 km is about 70 g. The overall setting provides a lift force of about 400 g on ground level. When reaching to 9–10 km height the system's net lift force should be about 0, making the system relatively altitude-stationary (Figure 4).

#### 4. Sensors, energy and thermal design

In this subsection, we cover the HAB payload components. First, we present the common needed and used sensors in "near-space" experiments. Then, a brief discussion on energy and thermal design is presented—followed by a discussion of how to test a potential payload (on the ground) for its ability to operate under near-space conditions.

#### 4.1. Sensors

3. Long-lasting "floating" balloon

thus they are not a practical solution for scientific researchers.

change its direction as dictated by the buoyancy force equation:

• A single 1000 g main latex balloon with a capability of 1300 g neck-lift.

conditions.

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• A two-parted payload:

exploded.

In a typical HAB configuration using latex balloon, the balloon will ascend and expand as the air pressure decreases with height due to the thin atmosphere. At a certain point, it will inflate up to its elastic point, explode and fall. This means that if we can make the balloon float in a relatively constant altitude, we can extend its lifespan and endurance. Moreover, fixing the balloon at high altitudes could also enable to test any desired hardware under near-space

Google's "loon project" is a good example for an HAB setting that is capable of floating up in the atmosphere for a long duration. However, such settings are expensive and complicated,

Our approach for "fixing" the balloon's altitude was directed toward a simple constriction, i.e., a main latex balloon and a cluster of foil balloons. Foil balloons are not elastic and cannot expand, thus their volume can be approximated as constant. This means that as the outer pressure drops due to the thin atmosphere, its upthrust force will weaken and might even

> FB ¼ rair � rgas

where rair is the surrounding air density, rgas is the helium density, V is the balloon volume and g is the gravitational force. We present a basic example of our current test setting design (Figure 3):

• Five non-lasting foil balloons with a fixed volume of about 110 l each, with a self-weight of approximately 90 g. Combining these balloons implies a weight variance of about 500 g.

Figure 3. A long-lasting HAB experiment. This setting retained its floating state for about 2.5 h before the main balloon

gV (1)

• GNSS (e.g., GPS): Global Navigation Satellite Systems refer to a positioning sensor commonly used for computing the 3D position in a typical horizontal accuracy of 2–3 m in the open sky (the vertical accuracy is often not as accurate as the horizontal – errors of 10–20 m are common even in the open sky). We have mostly used U-blox GNSS relievers which are becoming the industry standard for most COTS (Commercial Off-The-Shelf) drones. Remark: one should configure the GNSS receiver to a "balloon-mode" (Airborne) else the positioning might be limited to a low altitude of 12 km or less. Due to the nature of the balloon "Airborne < 1 g" is the preferred model.

[60–45]C, respectively. This makes the task of keeping the payload at "room level" temperature (i.e., [0, 45]C) vital. Packaging the payload with COTS boxes made from materials that provide proper thermal insulation such as expanded polystyrene (EPS) is sufficient for such need. Recall that in height of 16 km, the expected air pressure is 0.1 atm, while at 31 km, it is about 0.01 atm, so air-based passive cooling is significantly less efficient than on the ground. In practice: taking into an account the above considerations and the fact that most IoT components are suited for operating in near-space conditions it is easy to construct a thermalbalanced payload. Most of the required tests for the payload performances under near-space conditions can be performed with a simple setup, which consists of a vacuum chamber, a home freezer and a simple thermal camera (Figure 6). Table 3 depicts a thermal analysis of a

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Samsung Galaxy S6 mainboard under low ventilation conditions.

Figure 6. Thermal analysis of an android phone using a thermal camera (Op-gal's Therm-app).

velocity [m/s]

Portable 12,000 310 50 Altitude and velocity Medium Stationary 9000 10 6 Altitude and velocity Small Pedestrian 9000 30 20 Altitude and velocity Small Automotive 6000 100 15 Altitude and velocity Medium At sea 500 25 5 Altitude and velocity Medium Airborne <1 g 50,000 100 100 Altitude Large Airborne <2 g 50,000 250 100 Altitude Large Airborne <4 g 50,000 500 100 Altitude Large Wrist 9000 30 20 Altitude and velocity Medium

Table 3. Form: U-blox M8 N manual—make sure you use airborne mod (the default is portable—so the GPS will not

Max vertical velocity [m/s] Sanity check type Max position

deviation

Platform Max altitude [m] Max horizontal

work above 12 km).

Modern GPS can support 10 Hz position sampling rate—yet for most coming measurements, such sampling rate is not needed—as the dynamics of the balloon is very low. Lowering the positioning rate may also help reducing the energy consumption of the GNSS receiver.


#### 4.2. Energy and thermal design

It should be noted that performance of all batteries drops drastically at low temperatures starting 10C. At high elevation such as 10–30 km the outer temperature is expected to be

Figure 5. The balloon altitude over time, as recorded by the barometer sensor and our true altitude estimation.

[60–45]C, respectively. This makes the task of keeping the payload at "room level" temperature (i.e., [0, 45]C) vital. Packaging the payload with COTS boxes made from materials that provide proper thermal insulation such as expanded polystyrene (EPS) is sufficient for such need. Recall that in height of 16 km, the expected air pressure is 0.1 atm, while at 31 km, it is about 0.01 atm, so air-based passive cooling is significantly less efficient than on the ground.

are common even in the open sky). We have mostly used U-blox GNSS relievers which are becoming the industry standard for most COTS (Commercial Off-The-Shelf) drones. Remark: one should configure the GNSS receiver to a "balloon-mode" (Airborne) else the positioning might be limited to a low altitude of 12 km or less. Due to the nature of the

Modern GPS can support 10 Hz position sampling rate—yet for most coming measurements, such sampling rate is not needed—as the dynamics of the balloon is very low. Lowering the

• 9DoF: is basically a set of MEMS sensors: three axis magnetic field, three axis accelerometers, three axis gyroscopes. Combined they can be used to compute orientation. We have used Bosch BNO055 sensor which also has a true orientation filter—and found it both

• Barometer: this sensor measures the atmospheric pressure and temperature. This combination enables us to compute a naïve estimation of elevation in submeter accuracy. However, in our experiments, we noticed that in altitudes higher than 10 km, the barometer's altitude estimation started to slow its elevation in a certain pattern, which repeated itself. Using the GPS sensor measurements, we were able to estimate its true altitude, which consists with the expected altitudes. It should be denoted that although barometers mostly have an elevation accuracy of submeter (in some models subfeet), in high elevation the accuracy gets

• Temperature: thermocouple sensors are simple and robust sensors and being used to measure the inner and outer temperature of the payload. These values are significant for

It should be noted that performance of all batteries drops drastically at low temperatures starting 10C. At high elevation such as 10–30 km the outer temperature is expected to be

Figure 5. The balloon altitude over time, as recorded by the barometer sensor and our true altitude estimation.

worse, below is an example of real data of "faulty" barometer (Figure 5).

the proper operation of the electronic components and the batteries.

positioning rate may also help reducing the energy consumption of the GNSS receiver.

balloon "Airborne < 1 g" is the preferred model.

affordable and robust.

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4.2. Energy and thermal design

In practice: taking into an account the above considerations and the fact that most IoT components are suited for operating in near-space conditions it is easy to construct a thermalbalanced payload. Most of the required tests for the payload performances under near-space conditions can be performed with a simple setup, which consists of a vacuum chamber, a home freezer and a simple thermal camera (Figure 6). Table 3 depicts a thermal analysis of a Samsung Galaxy S6 mainboard under low ventilation conditions.

Figure 6. Thermal analysis of an android phone using a thermal camera (Op-gal's Therm-app).


Table 3. Form: U-blox M8 N manual—make sure you use airborne mod (the default is portable—so the GPS will not work above 12 km).

Figure 7. A typical thermal-balanced-payload, notice the ventilation hole marked with a circle.

A typical payload will include a GPS, microcontroller, LoRa modem, Geiger counter, barometer and humidity sensor (Figure 5). The total energy consumption is about 250 mW (Figure 7).

#### 5. Disposable cost-effective payload for low-bandwidth sensor data applications

In most cases, we usually direct our efforts toward recording and transmitting low-bandwidth sensor data or testing electronics at near-space conditions. As retrieving the payload with its data is not always certain, we designed the payload to be cost-effective and disposable and yet capable of long-range low-bandwidth communications.

6. Long-range communications

Figure 8. Gamma count vs. altitude.

reusability and security in most Radiosonde payloads.

Radiosonde is the most common type of payload which is capable of long-range communications suitable for HAB. A radiosonde can be regarded as a black-box which includes a variety of sensors and a radio transmitter. Radiosondes may come in various shapes and technologies, but in general, they measure: position (GPS), barometric pressure, humidity and temperature. They also may include some other related sensors such as Ozone meter. These data are transmitted to the GS using RF communications, commonly—UHF 400–406 MHz, and 1675– 1700 MHz. This solution's range is typically between 50 and 200 km that depends on environ-

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As mentioned above, radiosonde payloads are closed systems that limit the user's ability to customize them. This means that in order to transmit additional sensors data, an additional communication device is required as well. Moreover, they provide low-bandwidth and halfduplex (download only) communications. It should be noted that the RF, which is used by a radiosonde, is not an ISM band, and therefore, it might require RF approval by local authorities. Other concerns about using radiosonde communication abilities include the lack of frequency

Providing high-bandwidth communications enable us to obtain real-time measurements such as multi-spectral images and conducting high-resolution gamma-ray spectrometer measurements. Full-duplex communications enable us to interact with the payload, so we can remotely control the payload or the balloon motion. The ability to adapt the modem communications setting, i.e., reprogramming it in real-time makes it a more flexible solution that provides bandwidth and range according to the user's needs or environmental conditions. In our

6.1. Current state

mental conditions.

Our basic HAB payload setup typically includes the following components:


Such payload's BOM (Bill of Material) will cost about 100–120\$. The weight of the payload can be reduced to a sum of 150 g, making it suitable for 300 g HAB. The total cost including the cost of the launch will cost less than 200\$. In case there is no need for a Geiger counter, the overall BOM of the payload and balloon can be below 100\$.

Using this affordable payload design, we were able to perform several experiments in which we measured Gamma counts with respect to altitude and location in relative high accuracy. Figure 8 shows the real-time Geiger count as received at the GS from the payload (over 120 km range).

Figure 8. Gamma count vs. altitude.

#### 6. Long-range communications

#### 6.1. Current state

A typical payload will include a GPS, microcontroller, LoRa modem, Geiger counter, barometer and humidity sensor (Figure 5). The total energy consumption is about 250 mW (Figure 7).

In most cases, we usually direct our efforts toward recording and transmitting low-bandwidth sensor data or testing electronics at near-space conditions. As retrieving the payload with its data is not always certain, we designed the payload to be cost-effective and disposable and yet

• Environment conditions sensors (barometric pressure/altitude/temperature/humidity/

• Geiger counter-based on the new solid-state technology (which reduces the weight and

Such payload's BOM (Bill of Material) will cost about 100–120\$. The weight of the payload can be reduced to a sum of 150 g, making it suitable for 300 g HAB. The total cost including the cost of the launch will cost less than 200\$. In case there is no need for a Geiger counter, the overall

Using this affordable payload design, we were able to perform several experiments in which we measured Gamma counts with respect to altitude and location in relative high accuracy. Figure 8 shows the real-time Geiger count as received at the GS from the payload (over 120 km range).

5. Disposable cost-effective payload for low-bandwidth sensor data

Figure 7. A typical thermal-balanced-payload, notice the ventilation hole marked with a circle.

Our basic HAB payload setup typically includes the following components:

capable of long-range low-bandwidth communications.

• Versatile GNSS module capable of GPS, GLONASS.

• An actuator for releasing the payload on command.

BOM of the payload and balloon can be below 100\$.

applications

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• Arduino MCU.

Dewpoint).

• 433 MHz LoRa radio transceiver.

price of Geiger counter).

Radiosonde is the most common type of payload which is capable of long-range communications suitable for HAB. A radiosonde can be regarded as a black-box which includes a variety of sensors and a radio transmitter. Radiosondes may come in various shapes and technologies, but in general, they measure: position (GPS), barometric pressure, humidity and temperature. They also may include some other related sensors such as Ozone meter. These data are transmitted to the GS using RF communications, commonly—UHF 400–406 MHz, and 1675– 1700 MHz. This solution's range is typically between 50 and 200 km that depends on environmental conditions.

As mentioned above, radiosonde payloads are closed systems that limit the user's ability to customize them. This means that in order to transmit additional sensors data, an additional communication device is required as well. Moreover, they provide low-bandwidth and halfduplex (download only) communications. It should be noted that the RF, which is used by a radiosonde, is not an ISM band, and therefore, it might require RF approval by local authorities. Other concerns about using radiosonde communication abilities include the lack of frequency reusability and security in most Radiosonde payloads.

Providing high-bandwidth communications enable us to obtain real-time measurements such as multi-spectral images and conducting high-resolution gamma-ray spectrometer measurements. Full-duplex communications enable us to interact with the payload, so we can remotely control the payload or the balloon motion. The ability to adapt the modem communications setting, i.e., reprogramming it in real-time makes it a more flexible solution that provides bandwidth and range according to the user's needs or environmental conditions. In our experiments, we consider the minimally required coverage range to be about 40–50 km that is required for conducting HAB missions.

(~10 cents per 50 bytes)—so it is applicable for low-bandwidth missions. Yet it allows full control, two-way communication. Another satellite-related solution named "SPOT" is commonly used to track HABs. This one-way (transmission only) solution uses the "Global-Star" satellite network for near global coverage. Interestingly, we have found that the use of short message service (SMS) in cellular communications was relatively efficient and we were able to send and receive text messages from about 5000 m height when the expected height is about 2500 m.

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In this section, we present methods for constructing simple (DIY) payloads based on COTS devices. We start by presenting a naive attempt to shoot high-resolution images from high altitude—as part of a class challenge in the undergraduate "Autonomous Robotics" course during the year 2017 (given in the Computer Science Department at Ariel University). All suggested solutions included an Android phone with an international sim card and an app which captures time-laps photos with position while attempting to upload them using existing

• Smartphone-based payload—100–200 g. Android phones with the needed apps for timelapse camera (such as OpenCamera) and a cloud-based uploader app (such as Dropbox or Google drive). The phone was equipped with a sim card which can be used for uploading

• Thermal Box: the most common is polystyrene (ice-cream box)—which is needed to

Five different solutions were implemented (see Figure 10) mainly using the OpenCamera android open source. None of the payloads could capture reasonable images from high altitude

Figure 9. Three payloads ready to be launched—each with a smartphone and software for uploading the gathered data. As part of the navigation graduated course in Ariel University (Israel). None of the payloads could actually transmit good

and clear images from high altitude. All payloads eventually fall in Suraya. Over 200 km from launch.

maintain a controlled temperature for the phone electronics and batteries.

cloud uploader tools (Figure 9). The balloon launches included the following setting:

6.2. Cellular 3G/LTE communications

• A regular latex 600, 1000 g balloon.

the data—using a prepaid data plan.

Cellular 3G/LTE communications are intuitively a natural solution for full-duplex and highbandwidth communications that is communally used by "makers". However, a cellular device that has been used at high altitudes can be easily detected by multiple base-stations simultaneously. Generally, such device will be blocked by the cellular providers thus making its 3G/ LTE communications inoperable till the device returns to ground level.

This makes smartphones not suitable as a real-time communications solution for high altitudes. On the other hand, for low altitude applications or when it is known that the payload will fall in a cellular covered area, smartphones might be considered as a suitable communication solution. Denote that in many countries (e.g., USA) mobile phones are required to operate in "flight mode" while "in-air".

UHF RF communication such as 433, 866, and 915 MHz which are ISM RF bands can provide low-bandwidth and long-range communications solution. We have investigated many drones remote control (RC) two-way communication solutions, and we found that while they can provide long-range communications their high-energy consumption and cost make them less appealing for day-to-day HAB missions. As an example, the DragonLink RC technology, which is the gold standard for flying long-range drone's communications, required in our experiments 1.5 W transmitter for achieving the range of 40 km and a data rate of 19.2 kbps.

In our experiments, we found that LoRa technology-based devices are the most suitable and preferable solution for HAB missions' requirements. Meaning, they are robust, programmable, with a very low-energy consumption and affordable. With the right setting, we were able to achieve full-duplex communications with a 25 mW transmitter more than 120 km range and a data rate of 0.4 kbps.

Wi-Fi technology can provide high-bandwidth communications, however it was designed for as Wireless Local Area Network (WLAN). This means that with COTS hardware in a direct line of sight communications, its expected range is limited to hundreds of meters. We have designed a long-range Wi-Fi setting based on EZ-WiFiBroadcast settings. EZ-WiFiBroadcast is a special DIY design of Wi-Fi communications which is commonly used as a poor man's longrange HD FPV solution. With our current long-range Wi-Fi setting, we have been able to capture 720P video from a HAB at 9.8 km height and located about 15 km from the GS.

Free-Space Optical (FSO) also known as laser communications are a less common highbandwidth communication solution which can be achieved by the use of a robotic telescope which tracks in real-time the HAB. In a clear day, such device can track an HAB for over 50 km. In our experiments, we successfully tracked HABs for more than 70 km using low-cost Celestron StarBright XLT telescope with 127 mm aperture Schmidt-Cassegrain lens. As shown recently by Google in their Loon project "Demonstration of free-space optical communication for long-range data links between balloons on Project Loon". This kind of high-bandwidth communications is still extremely complicated and requires technical skills and efforts which are not common in most research groups.

Another commercial solution is Global satellite communications (we have used Iridium's twoway Short Burst Data—SBD), this kind of solution requires a "pay per message" data plan (~10 cents per 50 bytes)—so it is applicable for low-bandwidth missions. Yet it allows full control, two-way communication. Another satellite-related solution named "SPOT" is commonly used to track HABs. This one-way (transmission only) solution uses the "Global-Star" satellite network for near global coverage. Interestingly, we have found that the use of short message service (SMS) in cellular communications was relatively efficient and we were able to send and receive text messages from about 5000 m height when the expected height is about 2500 m.

#### 6.2. Cellular 3G/LTE communications

experiments, we consider the minimally required coverage range to be about 40–50 km that is

Cellular 3G/LTE communications are intuitively a natural solution for full-duplex and highbandwidth communications that is communally used by "makers". However, a cellular device that has been used at high altitudes can be easily detected by multiple base-stations simultaneously. Generally, such device will be blocked by the cellular providers thus making its 3G/

This makes smartphones not suitable as a real-time communications solution for high altitudes. On the other hand, for low altitude applications or when it is known that the payload will fall in a cellular covered area, smartphones might be considered as a suitable communication solution. Denote that in many countries (e.g., USA) mobile phones are required to operate

UHF RF communication such as 433, 866, and 915 MHz which are ISM RF bands can provide low-bandwidth and long-range communications solution. We have investigated many drones remote control (RC) two-way communication solutions, and we found that while they can provide long-range communications their high-energy consumption and cost make them less appealing for day-to-day HAB missions. As an example, the DragonLink RC technology, which is the gold standard for flying long-range drone's communications, required in our experiments 1.5 W transmitter for achieving the range of 40 km and a data rate of 19.2 kbps. In our experiments, we found that LoRa technology-based devices are the most suitable and preferable solution for HAB missions' requirements. Meaning, they are robust, programmable, with a very low-energy consumption and affordable. With the right setting, we were able to achieve full-duplex communications with a 25 mW transmitter more than 120 km range and a

Wi-Fi technology can provide high-bandwidth communications, however it was designed for as Wireless Local Area Network (WLAN). This means that with COTS hardware in a direct line of sight communications, its expected range is limited to hundreds of meters. We have designed a long-range Wi-Fi setting based on EZ-WiFiBroadcast settings. EZ-WiFiBroadcast is a special DIY design of Wi-Fi communications which is commonly used as a poor man's longrange HD FPV solution. With our current long-range Wi-Fi setting, we have been able to capture 720P video from a HAB at 9.8 km height and located about 15 km from the GS.

Free-Space Optical (FSO) also known as laser communications are a less common highbandwidth communication solution which can be achieved by the use of a robotic telescope which tracks in real-time the HAB. In a clear day, such device can track an HAB for over 50 km. In our experiments, we successfully tracked HABs for more than 70 km using low-cost Celestron StarBright XLT telescope with 127 mm aperture Schmidt-Cassegrain lens. As shown recently by Google in their Loon project "Demonstration of free-space optical communication for long-range data links between balloons on Project Loon". This kind of high-bandwidth communications is still extremely complicated and requires technical skills and efforts which

Another commercial solution is Global satellite communications (we have used Iridium's twoway Short Burst Data—SBD), this kind of solution requires a "pay per message" data plan

LTE communications inoperable till the device returns to ground level.

required for conducting HAB missions.

208 Space Flight

in "flight mode" while "in-air".

data rate of 0.4 kbps.

are not common in most research groups.

In this section, we present methods for constructing simple (DIY) payloads based on COTS devices. We start by presenting a naive attempt to shoot high-resolution images from high altitude—as part of a class challenge in the undergraduate "Autonomous Robotics" course during the year 2017 (given in the Computer Science Department at Ariel University). All suggested solutions included an Android phone with an international sim card and an app which captures time-laps photos with position while attempting to upload them using existing cloud uploader tools (Figure 9). The balloon launches included the following setting:


Five different solutions were implemented (see Figure 10) mainly using the OpenCamera android open source. None of the payloads could capture reasonable images from high altitude

Figure 9. Three payloads ready to be launched—each with a smartphone and software for uploading the gathered data. As part of the navigation graduated course in Ariel University (Israel). None of the payloads could actually transmit good and clear images from high altitude. All payloads eventually fall in Suraya. Over 200 km from launch.

—although at least three (out of five) phones made it back safely to the ground and two of them even sent few images—until it was discovered by a "lucky founder" or simply run out of power.

Although the suggested concept failed the overall solution of using a software-only solution based on affordable smartphones seems to be a feasible cost-effective solution to many nearspace applications.

#### 6.3. Long-range Wi-Fi communications

Long-range and high-bandwidth communication solutions suitable for HAB missions are not common, especially not as COTS hardware. High-bandwidth data applications such as multispectral imagery or high-resolution measurements have a great value for exploring various electrical phenomena such as lightening discharges, sprites or blue-jets in the atmosphere and other aspects of this environment.

For providing high-bandwidth communication capabilities, we are directing our efforts on utilizing COTS communications hardware based on IEEE 802.11 standard WLAN which is also known as Wi-Fi. Wi-Fi networks can easily provide high-bandwidth communications but with COTS hardware they have a very limited range. Using a much more sophisticated hardware can extend its range dramatically to a few kilometers, but such systems are costly, with high-power demands and with a form factor that is not suited for a typical HAB's payload (Figure 11).

In theory, the use of a high gain directional antenna about 18–24 dBi for the receiver at the ground station and a directional antenna with a gain of about 10–14 dBi should provide us a link budget greater than 150 dB. Such link budget should enable communications for long ranges estimated at 10–30 km on regular conditions. In optimal conditions and a Forward Error Correction (FEC) mechanism, the range can be extended to about 50 km (Figure 12).

This lead us to investigate a different approach that uses COTS Wi-Fi hardware but in a nontraditional way. Some of the IEEE 802.11 network interfaces can operate in a special debug

mode that allows them to transmit and receive Wi-Fi communications with no regards to the IEEE 802.11 standard itself. As such, "makers" have used this feature for creating a "poor man's" long-range HD FPV solution. We based our system on "bortek"'s version of EZ-

Figure 12. Keep it simple: launching two simple payloads: Raspberry-Pi (upper) and an android smartphone (taking this

Figure 11. HAB ground station (GS): Left: The GS in general: two robotic telescopes with (auto-track) and a high gain 24dBi Wi-Fi antenna. Right: the robotic telescope: (a) A view-finder webcam. (b) A Wi-Fi + 3G router. So the telescope can be controlled globally. (c) A Pi-camera mounted to the telescope eye-view. (d) A Raspberry Pi which controls the telescope

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• Raspberry-Pi (RPi) device usually Raspberry Pi 0, with RPi-Cam camera and a Wi-Fi Network Interface Card (NIC) with an external directional 14 dBi flat panel antenna with

• Ground station based on another RPi device with an external 20 dBi directional antenna.

WiFiBroadcast (Figure 13).

image).

its face directed down.

Typically, the Wi-Fi long-range system includes:

using either visual tracking and GPS coordinates.

Figure 10. An image of the sky that was made by an android smartphone (Xiaomi Redmi 4A). In this experiment, the phone's camera was out of focus. It might be due to ice on the camera lens. The images were successfully uploaded after the payload has made it to the ground.

Figure 11. HAB ground station (GS): Left: The GS in general: two robotic telescopes with (auto-track) and a high gain 24dBi Wi-Fi antenna. Right: the robotic telescope: (a) A view-finder webcam. (b) A Wi-Fi + 3G router. So the telescope can be controlled globally. (c) A Pi-camera mounted to the telescope eye-view. (d) A Raspberry Pi which controls the telescope using either visual tracking and GPS coordinates.

Figure 12. Keep it simple: launching two simple payloads: Raspberry-Pi (upper) and an android smartphone (taking this image).

mode that allows them to transmit and receive Wi-Fi communications with no regards to the IEEE 802.11 standard itself. As such, "makers" have used this feature for creating a "poor man's" long-range HD FPV solution. We based our system on "bortek"'s version of EZ-WiFiBroadcast (Figure 13).

Typically, the Wi-Fi long-range system includes:

—although at least three (out of five) phones made it back safely to the ground and two of them even sent few images—until it was discovered by a "lucky founder" or simply run out of power. Although the suggested concept failed the overall solution of using a software-only solution based on affordable smartphones seems to be a feasible cost-effective solution to many near-

Long-range and high-bandwidth communication solutions suitable for HAB missions are not common, especially not as COTS hardware. High-bandwidth data applications such as multispectral imagery or high-resolution measurements have a great value for exploring various electrical phenomena such as lightening discharges, sprites or blue-jets in the atmosphere and

For providing high-bandwidth communication capabilities, we are directing our efforts on utilizing COTS communications hardware based on IEEE 802.11 standard WLAN which is also known as Wi-Fi. Wi-Fi networks can easily provide high-bandwidth communications but with COTS hardware they have a very limited range. Using a much more sophisticated hardware can extend its range dramatically to a few kilometers, but such systems are costly, with high-power demands and with a form factor that is not suited for a typical HAB's

In theory, the use of a high gain directional antenna about 18–24 dBi for the receiver at the ground station and a directional antenna with a gain of about 10–14 dBi should provide us a link budget greater than 150 dB. Such link budget should enable communications for long ranges estimated at 10–30 km on regular conditions. In optimal conditions and a Forward Error Correction (FEC) mechanism, the range can be extended to about 50 km (Figure 12).

This lead us to investigate a different approach that uses COTS Wi-Fi hardware but in a nontraditional way. Some of the IEEE 802.11 network interfaces can operate in a special debug

Figure 10. An image of the sky that was made by an android smartphone (Xiaomi Redmi 4A). In this experiment, the phone's camera was out of focus. It might be due to ice on the camera lens. The images were successfully uploaded after

space applications.

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payload (Figure 11).

the payload has made it to the ground.

6.3. Long-range Wi-Fi communications

other aspects of this environment.


Even though that our research on this approach is at early stages we have been able already to capture 720P video from a HAB at 9.8 km height and in an estimated distance of about 15 km with our long-range Wi-Fi communication system. Figure 13 is an image captured in this particular HAB mission. We found that flat directional antennas perform quite well as long as the angle between the balloon and the GS was not too wide.

As the GS design is compact it can be used as a mobile ground station located on top of a car which "chases" the balloon.

7. Near-space return to home micro drone

The proposed platform has the following properties:

tion system for RC & telemetry data.

3. Near-space flight mode for smart decline.

We present a near-space drone, which is affordable, robust and may weight below the FAA regulations (300 g). The micro-UAV has a unique RTH control algorithm adjust to near-space conditions and on board black box for storing a wide range of sensor measurements (Figure 15).

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1. Low-cost, lightweight electric UAV which was equipped with: multiple real-time sensors, HD cameras, a Pixhawk flight controller, GNSS receiver and long-range RF communica-

Figure 15. Four different models of RTH micro drones. Each of them was tested for autonomous flight launched from a

2. Smart release mechanism with several parameters for autonomous operation.

7.1. Drone structure

balloon.

Figure 14. LoRa module vs. iridium module.

#### 6.4. Long-range communication LoRa vs. iridium

In Europe, there is a well-established RF solution for tracking on HABs led by the UK High Altitude Society (https://ukhas.org.uk). This cooperative solution allows an online tracking mechanism based on COTS Software Defined Radio (SDR). This system is based on a fixed lowbandwidth protocol—mostly at the 434.075 MHz frequency and has been successfully in use for hundreds of launches annually. Yet, in many cases the UKHAS system is not suitable due to geolocation, bandwidth or even security reasons. In this work, we mainly focus on such cases in which a "real" Ground Station is needed. The iridium modem allows true global coverage and two-way communication, yet it is relatively expensive (300\$) and requires a data plan which cost about 2\$ for kB (a compressed single JPG image of 100 kB—will cost about 200\$). This kind of pricing makes it applicable mainly for strictly low bit rate application. The LoRa modem is an affordable (10–20\$) system with adaptive bit rate and works in unlessens band. The expected range for LoRa communications is over 120 km, while in a few places around the world, LoRa gateways are started to be deployed so that the expected route can be covered. But in general even with a single LoRa gateway it is expected to cover the balloon route (50–200 km)—using a standard UHF Yagi antenna in the expected range. We conclude that the LoRa solution can be an affordable complementary communication solution. It can be connected to a smartphone allowing long-range communications coverage and with actively connected to the Iridium system it can benefit the most to the satellite communication (Figure 14).

Figure 13. The smartphone payload shot from an upper payload based on a Raspberry-Pi camera equipped with a longrange Wi-Fi transmitter. The picture was taken at about 9.7 km above ground.

Figure 14. LoRa module vs. iridium module.

#### 7. Near-space return to home micro drone

#### 7.1. Drone structure

Even though that our research on this approach is at early stages we have been able already to capture 720P video from a HAB at 9.8 km height and in an estimated distance of about 15 km with our long-range Wi-Fi communication system. Figure 13 is an image captured in this particular HAB mission. We found that flat directional antennas perform quite well as long as

As the GS design is compact it can be used as a mobile ground station located on top of a car

In Europe, there is a well-established RF solution for tracking on HABs led by the UK High Altitude Society (https://ukhas.org.uk). This cooperative solution allows an online tracking mechanism based on COTS Software Defined Radio (SDR). This system is based on a fixed lowbandwidth protocol—mostly at the 434.075 MHz frequency and has been successfully in use for hundreds of launches annually. Yet, in many cases the UKHAS system is not suitable due to geolocation, bandwidth or even security reasons. In this work, we mainly focus on such cases in which a "real" Ground Station is needed. The iridium modem allows true global coverage and two-way communication, yet it is relatively expensive (300\$) and requires a data plan which cost about 2\$ for kB (a compressed single JPG image of 100 kB—will cost about 200\$). This kind of pricing makes it applicable mainly for strictly low bit rate application. The LoRa modem is an affordable (10–20\$) system with adaptive bit rate and works in unlessens band. The expected range for LoRa communications is over 120 km, while in a few places around the world, LoRa gateways are started to be deployed so that the expected route can be covered. But in general even with a single LoRa gateway it is expected to cover the balloon route (50–200 km)—using a standard UHF Yagi antenna in the expected range. We conclude that the LoRa solution can be an affordable complementary communication solution. It can be connected to a smartphone allowing long-range communications coverage and with actively connected to the Iridium

Figure 13. The smartphone payload shot from an upper payload based on a Raspberry-Pi camera equipped with a long-

the angle between the balloon and the GS was not too wide.

system it can benefit the most to the satellite communication (Figure 14).

range Wi-Fi transmitter. The picture was taken at about 9.7 km above ground.

6.4. Long-range communication LoRa vs. iridium

which "chases" the balloon.

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We present a near-space drone, which is affordable, robust and may weight below the FAA regulations (300 g). The micro-UAV has a unique RTH control algorithm adjust to near-space conditions and on board black box for storing a wide range of sensor measurements (Figure 15). The proposed platform has the following properties:


Figure 15. Four different models of RTH micro drones. Each of them was tested for autonomous flight launched from a balloon.

The basic requirement of the UAV is the ability of autonomous RTH or any other Geo location. The UAV needs to be lightweight, aerodynamic wing structure for fast and smooth flight and at list extended range of 50 km for RTH. Denote that in most cases flying back home will require flying against the wind (Figure 16).

#### 7.2. Smart release mechanism

The smart release mechanism is established from two main elements: mechanical mechanism and autonomous smart release software. The mechanical mechanism has two construction sets: Servo or Fuse wire. The servo is operated with PWM signal, and the fuse wire burns from relay. One of the most important things is the way the balloon attached the release mechanism to the UAV without affecting the UAV fly ability and minimal change of the aerodynamic, because of that the release mechanism mounted on the balloon payload. The autonomous smart release software is an algorithm that gets a several sensor parameters and decides if to release the UAV. The algorithm has the next prioritization: balloon burst, RC signal, altitude, battery, and geo fence. The RC signal is the only parameter that comes from the ground, the rest calculated on the MCU (Figure 17).

#### 7.3. Near-space flight mode

This mode has few parameters for controlling on smart decline. After the UAV release from the balloon, it will open parachute to altitude that set on the algorithm, the next step is to release

the parachute and glide with a constant decline rate to altitude that set on the algorithm and

Figure 18. Getting back home: A massive UHF transition caused the drone to get into the "fail-safe" state, releasing the

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Currently, we are constructing a micro wing-shape UAV with solar panels for energy harvesting; this will allow us to perform a much longer time and range experiment using super-pressure balloons. Release the drone on a "sunny morning"—allowing it to fly for up to 6 h during daytime covering 100–200 km. Such distance should be sufficient for finding a proper landing region

Launching a HAB requires authorization and following local regulations. We present here some of the US Federal Aviation Administration (FAA) regulations. Please note that even though many countries tend to adapt these regulations, local regulations might differ from the following

then open the motor and fly back home (Figure 18).

drone which in turn flyied back to "home" autonomously.

(Figures 19 and 20).

7.4. Regulation and safety

Figure 17. The RTH payload is going up.

(FAA Part 101 and 14 CFR Part 48):

Figure 16. Full flight path of an HAB and RTH payload by a micro drone.

Figure 17. The RTH payload is going up.

The basic requirement of the UAV is the ability of autonomous RTH or any other Geo location. The UAV needs to be lightweight, aerodynamic wing structure for fast and smooth flight and at list extended range of 50 km for RTH. Denote that in most cases flying back home will

The smart release mechanism is established from two main elements: mechanical mechanism and autonomous smart release software. The mechanical mechanism has two construction sets: Servo or Fuse wire. The servo is operated with PWM signal, and the fuse wire burns from relay. One of the most important things is the way the balloon attached the release mechanism to the UAV without affecting the UAV fly ability and minimal change of the aerodynamic, because of that the release mechanism mounted on the balloon payload. The autonomous smart release software is an algorithm that gets a several sensor parameters and decides if to release the UAV. The algorithm has the next prioritization: balloon burst, RC signal, altitude, battery, and geo fence. The RC signal is the only parameter that comes from the ground, the

This mode has few parameters for controlling on smart decline. After the UAV release from the balloon, it will open parachute to altitude that set on the algorithm, the next step is to release

require flying against the wind (Figure 16).

rest calculated on the MCU (Figure 17).

Figure 16. Full flight path of an HAB and RTH payload by a micro drone.

7.3. Near-space flight mode

7.2. Smart release mechanism

214 Space Flight

Figure 18. Getting back home: A massive UHF transition caused the drone to get into the "fail-safe" state, releasing the drone which in turn flyied back to "home" autonomously.

the parachute and glide with a constant decline rate to altitude that set on the algorithm and then open the motor and fly back home (Figure 18).

Currently, we are constructing a micro wing-shape UAV with solar panels for energy harvesting; this will allow us to perform a much longer time and range experiment using super-pressure balloons. Release the drone on a "sunny morning"—allowing it to fly for up to 6 h during daytime covering 100–200 km. Such distance should be sufficient for finding a proper landing region (Figures 19 and 20).

#### 7.4. Regulation and safety

Launching a HAB requires authorization and following local regulations. We present here some of the US Federal Aviation Administration (FAA) regulations. Please note that even though many countries tend to adapt these regulations, local regulations might differ from the following (FAA Part 101 and 14 CFR Part 48):

4. The balloon cannot use a rope or other device for suspension of the payload that requires an impact force of more than 50 pounds to separate the suspended payload from the balloon.

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5. No person may operate any balloon in a manner that creates a hazard to other persons, or

6. No person operating any balloon may allow an object to be dropped therefrom, if such

7. The owner must register their HAB as part of the FAA's new Unmanned Aircraft System

Here are the main rules of thumb we have used in our HAB launches (on top of the local

1. It is highly recommended to update the related FAA authorities and get a permission in

2. Validate in real-time the conformation for the launch, a few minutes prior to the lunch.

4. The overall weight of all payloads should not be more than 1 kg, "Return to Launch" UAVs should weigh less than 500 g—preferable below 300 g (FAA regulations).

5. The maximal declining speed of the falling payload (below 5000 m) should not exceed

6. The usage of a parachute cannot guarantee declining speed or velocity. As in this method the overall max weight per square cm should be below some value, we strongly recommend a weight-to-size ratio of no more than 2.5 g per cm square, e.g., a cube payload of 1 l

8. If there are still some safety issues with the HAB, make sure its planned route is not above populated areas—preferably above the sea. Aborting a HAB-UAV mission into the sea is a

In the last decade, HAB experiments, which were considered esoteric and rare, have become more applicable for scientific researchers and near-space experiments. Today, the overall cost of an HAB experiment can reach up to \$500. Radiosondes are commonly used for transmitting the sensory data in real-time. However, using this technology has a limited communication capability and is very hard to customize. New long-range wireless communication technologies such as LoRa allow us to transmit a wide range of sensory data with both substantial low-cost and light

3. Make sure you are not launching the HAB nearby airports or other no-flight-zones.

(UAS) laws. The registration number must be marked on each HAB flight.

action creates a hazard to other persons or their property.

their property.

aviation regulations):

some velocity (say e.g., 12 m/s).

should not weight more than 250 g.

8. Discussion and conclusion

7. Secure each payload's component to prevent its fall.

safe backup plan—and in HAB lots can go wrong.

9. Only launch at a safe zone—where there are no power-lines or buildings.

advance.

Figure 19. 290 g RTH micro-UAV, with a release carbon strip on its backend.

Figure 20. RTH micro drone lunching.


Here are the main rules of thumb we have used in our HAB launches (on top of the local aviation regulations):


#### 8. Discussion and conclusion

1. Any cellular phones must be turned off (airplane mode enabled) for any aircraft and/or

2. Any individual payload must weight less than 4 pounds and have a weight-to-size ratio of less than 3.0 ounces/square inch (total weight of the payload only divided by its smallest face).

3. Total payload of two or more packages carried by one balloon must be less than 12 pounds

balloon as soon as it leaves the ground.

Figure 20. RTH micro drone lunching.

Figure 19. 290 g RTH micro-UAV, with a release carbon strip on its backend.

total.

216 Space Flight

In the last decade, HAB experiments, which were considered esoteric and rare, have become more applicable for scientific researchers and near-space experiments. Today, the overall cost of an HAB experiment can reach up to \$500. Radiosondes are commonly used for transmitting the sensory data in real-time. However, using this technology has a limited communication capability and is very hard to customize. New long-range wireless communication technologies such as LoRa allow us to transmit a wide range of sensory data with both substantial low-cost and light

weight setup. The maximum data rate provided by LoRa technology is 37.5 kbps, which is sufficient for two-way telemetry along with a wide range of sensory data but is not suitable for high-data-rate applications such as real-time video data. For that we found long-range Wi-Fi techniques to be a prominent strategy: allowing us transmission of live video data up to ranges of about 15–30 km. For long duration application in which the balloon may circle the world, we also present a global two-way communication solution based on Iridium modem.

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[2] Kräuchi A, Philipona R. Return glider radiosonde for in situ upper-air research measure-

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[4] Solomon S, Rosenlof KH, Portmann RW, Daniel JS, Davis SM, Sanford TJ, et al. Contributions of stratospheric water vapor to decadal changes in the rate of global warming.

[5] Seidel DJ, Berger FH, Immler F, Sommer M, Vömel H, Diamond HJ, et al. Reference upper-air observations for climate: Rationale, progress, and plans. Bulletin of the Amer-

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As the state-of-the-art of communications is still limited, we presented a whole different approach which focused on retrieving the payload in a safe and secure way. Such solution overcomes the need for transmitting the measured data wirelessly—as all the needed information are stored on board of the UAV.

Moreover, this approach highly reduces the risk of losing precious equipment and enables reusing the experiment platform over and over again. In the past, developing and operating an autonomous UAV system was a complicated and costly project. However, in recent years the successful efforts of the toy and hobbies industries to make UAVs accessible and simple to operate provided the opportunity for using UAVs as a common research tool. As such it can be used as a practical and cost-effective solution for returning the payload home with a relatively simple release mechanism and auto-pilot controller.

Based on six different experiments performed during 2016–2017, we conclude that the suggested strategy of using an autonomous UAVas a generic multi-parametric near-space platform is suitable for tropospheric remote sensing and for testing electronic components in near-space conditions.

Current research focuses on exceeding the operational capabilities of long-range Wi-Fi to a fullduplex communication channel and extending its range even further with the development of a high-gain antenna tracker. The deployment of LoRa WAN infrastructure can extend the HAB's communication service over huge areas.

Finally, the current range of RTF autonomous micro UAV is about 30 km. We expect that after optimizing the algorithm for the decline mode (from near space to ground), such range may be extended to 50–100 km with a relatively high probability of success.

#### Author details

Kobi Gozlan<sup>1</sup> , Yuval Reuveni1,2,3,4, Kfir Cohen1 , Boaz Ben-Moshe<sup>1</sup> \* and Eyal Berliner1,2,5

\*Address all correspondence to: benmo@g.ariel.ac.il


#### References

weight setup. The maximum data rate provided by LoRa technology is 37.5 kbps, which is sufficient for two-way telemetry along with a wide range of sensory data but is not suitable for high-data-rate applications such as real-time video data. For that we found long-range Wi-Fi techniques to be a prominent strategy: allowing us transmission of live video data up to ranges of about 15–30 km. For long duration application in which the balloon may circle the world, we

As the state-of-the-art of communications is still limited, we presented a whole different approach which focused on retrieving the payload in a safe and secure way. Such solution overcomes the need for transmitting the measured data wirelessly—as all the needed informa-

Moreover, this approach highly reduces the risk of losing precious equipment and enables reusing the experiment platform over and over again. In the past, developing and operating an autonomous UAV system was a complicated and costly project. However, in recent years the successful efforts of the toy and hobbies industries to make UAVs accessible and simple to operate provided the opportunity for using UAVs as a common research tool. As such it can be used as a practical and cost-effective solution for returning the payload home with a relatively

Based on six different experiments performed during 2016–2017, we conclude that the suggested strategy of using an autonomous UAVas a generic multi-parametric near-space platform is suitable for tropospheric remote sensing and for testing electronic components in near-space conditions.

Current research focuses on exceeding the operational capabilities of long-range Wi-Fi to a fullduplex communication channel and extending its range even further with the development of a high-gain antenna tracker. The deployment of LoRa WAN infrastructure can extend the

Finally, the current range of RTF autonomous micro UAV is about 30 km. We expect that after optimizing the algorithm for the decline mode (from near space to ground), such range may be

, Boaz Ben-Moshe<sup>1</sup>

\* and Eyal Berliner1,2,5

also present a global two-way communication solution based on Iridium modem.

tion are stored on board of the UAV.

218 Space Flight

simple release mechanism and auto-pilot controller.

HAB's communication service over huge areas.

Author details

Kobi Gozlan<sup>1</sup>

extended to 50–100 km with a relatively high probability of success.

, Yuval Reuveni1,2,3,4, Kfir Cohen1

2 Department of Management, Bar-Ilan University, Israel

5 School of Sustainability, Interdisciplinary Center (IDC) Herzliya, Israel

\*Address all correspondence to: benmo@g.ariel.ac.il

3 Department of Physics, Ariel University, Israel

1 K&CG lab, Ariel University, Israel

4 Eastern R&D Center, Ariel, Israel


**Section 6**

**Deep-Space Flight**


**Section 6**

**Deep-Space Flight**

[15] Regener E. New results in cosmic ray measurements. Nature. 1933;132:696-698

History of geo - and space. Sciences. 2014;5(2):175

southern Appalachians. 2014 NCUR; 2015

220 Space Flight

[16] Carlson P, Watson AA. Erich Regener and the ionisation maximum of the atmosphere.

[17] Carmichael-Coker MK. Increase of ionizing radiation at the Pfotzer maximum over the

[18] Mishev A. Short- and medium-term induced ionization in the earth atmosphere by galactic and solar cosmic rays. International Journal of Atmospheric Sciences. 2013;2013:9

**Chapter 12**

**Provisional chapter**

**Cassini Spacecraft-DSN Communications, Handling**

**Anomalous Link Conditions, and Complete Loss-of-**

**Cassini Spacecraft-DSN Communications, Handling** 

DOI: 10.5772/intechopen.72075

**Anomalous Link Conditions, and Complete Loss-of-**

Once spacecraft are launched, it is impossible for engineers to physically repair anything that breaks onboard the vehicle. Instead, remote solutions must be employed to address spacecraft anomalies and fault conditions. To achieve this goal, telemetered data from the spacecraft are collected and assess by ground personnel to resolve problems. However, if the ground-to-spacecraft communication system breaks down, or the vehicle delivers an anomalous signal, a rigorous protocol must be employed in order to re-establish or fix the telecommunications link. There are several factors that can contribute to link problems, such as malfunctions or mishandling of the ground station equipment, onboard failures of the spacecraft's flight software coding, or even mishaps caused by the space environment itself. This chapter details the anomaly recovery protocols developed for the Cassini Mission-to-Saturn project, to resolve anomalous link problems as well re-acquisition of the spacecraft should a complete Loss of Signal (LOS)

**Keywords:** Cassini, spacecraft, Saturn, deep space network communications, fault

protection, loss-of-spacecraft-signal, anomalous downlink

© 2016 The Author(s). Licensee InTech. 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,

© 2018 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.

and reproduction in any medium, provided the original work is properly cited.

Despite the vast distance between remote-controlled interplanetary spacecraft launched from earth and the Deep Space Network (DSN) ground stations that operate them, the communications link to the spacecraft is very reliable, thanks to the extraordinary telecommunication capabilities built into NASA's DSN antennas around the world and the spacecraft's own system design. For the Cassini Mission-to-Saturn spacecraft (**Figure 1**), it takes nearly an hour

**Spacecraft Signal**

**Spacecraft Signal**

Additional information is available at the end of the chapter

Additional information is available at the end of the chapter

http://dx.doi.org/10.5772/intechopen.72075

Paula S. Morgan

**Abstract**

condition occur.

**1. Introduction**

Paula S. Morgan

**Provisional chapter**

#### **Cassini Spacecraft-DSN Communications, Handling Anomalous Link Conditions, and Complete Loss-of-Spacecraft Signal Anomalous Link Conditions, and Complete Loss-of-Spacecraft Signal**

**Cassini Spacecraft-DSN Communications, Handling** 

DOI: 10.5772/intechopen.72075

Paula S. Morgan Paula S. Morgan Additional information is available at the end of the chapter

Additional information is available at the end of the chapter

http://dx.doi.org/10.5772/intechopen.72075

#### **Abstract**

Once spacecraft are launched, it is impossible for engineers to physically repair anything that breaks onboard the vehicle. Instead, remote solutions must be employed to address spacecraft anomalies and fault conditions. To achieve this goal, telemetered data from the spacecraft are collected and assess by ground personnel to resolve problems. However, if the ground-to-spacecraft communication system breaks down, or the vehicle delivers an anomalous signal, a rigorous protocol must be employed in order to re-establish or fix the telecommunications link. There are several factors that can contribute to link problems, such as malfunctions or mishandling of the ground station equipment, onboard failures of the spacecraft's flight software coding, or even mishaps caused by the space environment itself. This chapter details the anomaly recovery protocols developed for the Cassini Mission-to-Saturn project, to resolve anomalous link problems as well re-acquisition of the spacecraft should a complete Loss of Signal (LOS) condition occur.

**Keywords:** Cassini, spacecraft, Saturn, deep space network communications, fault protection, loss-of-spacecraft-signal, anomalous downlink

#### **1. Introduction**

Despite the vast distance between remote-controlled interplanetary spacecraft launched from earth and the Deep Space Network (DSN) ground stations that operate them, the communications link to the spacecraft is very reliable, thanks to the extraordinary telecommunication capabilities built into NASA's DSN antennas around the world and the spacecraft's own system design. For the Cassini Mission-to-Saturn spacecraft (**Figure 1**), it takes nearly an hour

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. © 2018 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.

© 2016 The Author(s). Licensee InTech. This chapter is distributed under the terms of the Creative Commons

ground commands for a predetermined (programmable) period of time, using a "countdown timer" which decrements until it is reset by a ground command or reaches "0" (which triggers the response). An extended series of actions are then commanded by FP to re-establish ground commandability by configuring various telecom arrangements and spacecraft attitudes in an attempt to find a viable U/L path. Each attempt by the response to command a new path is separated by an appropriate ground response interval for the SOFS team to re-acquire the spacecraft

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225

In all anomalous spacecraft D/L cases, it is desirable to re-establish spacecraft communications before the Command Loss Response activates in order to avoid the autonomous commanded actions of the FP: termination of the onboard running sequence (lost science opportunities), device swaps, propellant consumption via commanded turns, etc. Therefore, an expedient method for identifying possible anomalous/LOS causes is highly desirable before the FP activates, if possible. To aid in this goal, an Excel tool was developed to supplement the LOS Recovery Protocol in "timeline" format. Described herein are the optimized solutions implemented on Cassini for re-acquisition of the spacecraft's signal during anomalous D/L and LOS events, as well as an expedient method for recovery from the actions of the Command Loss

NASA's Cassini Mission-to-Saturn spacecraft is the first robotic mission ever to orbit the planet Saturn. Managed by the Jet Propulsion Laboratory (JPL) in Pasadena, this flagship-class mission is composed of 11 operating scientific instruments which study many intriguing features of Saturn, its moons, and ring system. The Cassini Program is an international cooperative effort involving primarily NASA, the European Space Agency (ESA), and the Italian Space Agency (Agenzia Spaziale Italiana, ASI). Cassini is the fourth spacecraft to visit the Saturnian system (but is the first vehicle to enter its orbit), and is composed of the NASA/ASI Cassini orbiter and the ESA-developed Huygens probe. Cassini launched on October 15, 1997, arriving at Saturn in 2004, after performing scientific observation of Earth's moon, Venus, and Jupiter (as well as participating in several scientific experiments) during its 6.7 year cruise period. Cassini's suite of (currently operating) science instruments consists of the following (**Figure 2**):

via U/L command.

Response, if activated.

**2. The Cassini mission**

**1.** Composite Infrared Spectrometer (CIRS)

**2.** Ion & Neutral Mass Spectrometer (INMS)

**4.** Ultraviolet Imaging Spectrograph (UVIS)

**6.** Magnetospheric Imaging Instrument (MIMI)

**5.** Imaging Science Subsystem (ISS)

**3.** Visible & Infrared Mapping Spectrometer (VIMS)

**Figure 1.** The Cassini-Huygens spacecraft.

and a half for commands from the Spacecraft Operations Flight Team (SOFS) here on earth to reach Cassini, where the orbiter is touring the Saturnian system (~8.5 AU). Yet, an anomalous downlink (D/L) signal condition can occur (or complete LOS) from several sources: environmental effects such as bad weather conditions at the DSN station or station problems (broken equipment), erroneous ground commands uplinked (U/L) to the spacecraft by the SOFS team, errors in the onboard running sequence, spacecraft pointing errors, internal FSW errors, or computer platform failures can cause problems when attempting to acquire the spacecraft's D/L signal. The space environment itself can also contribute to an LOS condition, since cosmic ray bombardment on the spacecraft's systems can cause spurious Solid State Power Switch (SSPS) trip-off of the spacecraft's Radio Frequency System (RFS) units, as well activations of the onboard Fault Protection (FP) routines which will reconfigure to redundant backup RFS units, so that reconfiguration by the ground is required in order to lock-up on the spacecraft's D/L signal.

To safeguard against these DSN-spacecraft link problems, troubleshooting methods have been developed by the Cassini SOFS team to diagnose and resolve conditions that inhibit spacecraft signal acquisition. A "Loss of Downlink Signal Recovery" protocol was developed for the SOFS team to follow in the event of an anomalous D/L signal (or completed LOS), as well as special FP which is implemented into Cassini's onboard FSW. This algorithm will monitor for prolonged absence of ground commanding, eventually invoking a "Loss of Commandability" FP (FP which is typically implemented into most deep space missions to safeguard against these undetected, sometimes waived or ground-induced failure conditions). Called "Command Loss FP" (from the perspective of the spacecraft since it's no longer receiving ground commands), this "catch-all" type of autonomous monitor-response algorithm will observe the absence of ground commands for a predetermined (programmable) period of time, using a "countdown timer" which decrements until it is reset by a ground command or reaches "0" (which triggers the response). An extended series of actions are then commanded by FP to re-establish ground commandability by configuring various telecom arrangements and spacecraft attitudes in an attempt to find a viable U/L path. Each attempt by the response to command a new path is separated by an appropriate ground response interval for the SOFS team to re-acquire the spacecraft via U/L command.

In all anomalous spacecraft D/L cases, it is desirable to re-establish spacecraft communications before the Command Loss Response activates in order to avoid the autonomous commanded actions of the FP: termination of the onboard running sequence (lost science opportunities), device swaps, propellant consumption via commanded turns, etc. Therefore, an expedient method for identifying possible anomalous/LOS causes is highly desirable before the FP activates, if possible. To aid in this goal, an Excel tool was developed to supplement the LOS Recovery Protocol in "timeline" format. Described herein are the optimized solutions implemented on Cassini for re-acquisition of the spacecraft's signal during anomalous D/L and LOS events, as well as an expedient method for recovery from the actions of the Command Loss Response, if activated.

#### **2. The Cassini mission**

and a half for commands from the Spacecraft Operations Flight Team (SOFS) here on earth to reach Cassini, where the orbiter is touring the Saturnian system (~8.5 AU). Yet, an anomalous downlink (D/L) signal condition can occur (or complete LOS) from several sources: environmental effects such as bad weather conditions at the DSN station or station problems (broken equipment), erroneous ground commands uplinked (U/L) to the spacecraft by the SOFS team, errors in the onboard running sequence, spacecraft pointing errors, internal FSW errors, or computer platform failures can cause problems when attempting to acquire the spacecraft's D/L signal. The space environment itself can also contribute to an LOS condition, since cosmic ray bombardment on the spacecraft's systems can cause spurious Solid State Power Switch (SSPS) trip-off of the spacecraft's Radio Frequency System (RFS) units, as well activations of the onboard Fault Protection (FP) routines which will reconfigure to redundant backup RFS units, so that reconfiguration by the ground is required in order to lock-up on the spacecraft's

To safeguard against these DSN-spacecraft link problems, troubleshooting methods have been developed by the Cassini SOFS team to diagnose and resolve conditions that inhibit spacecraft signal acquisition. A "Loss of Downlink Signal Recovery" protocol was developed for the SOFS team to follow in the event of an anomalous D/L signal (or completed LOS), as well as special FP which is implemented into Cassini's onboard FSW. This algorithm will monitor for prolonged absence of ground commanding, eventually invoking a "Loss of Commandability" FP (FP which is typically implemented into most deep space missions to safeguard against these undetected, sometimes waived or ground-induced failure conditions). Called "Command Loss FP" (from the perspective of the spacecraft since it's no longer receiving ground commands), this "catch-all" type of autonomous monitor-response algorithm will observe the absence of

D/L signal.

224 Space Flight

**Figure 1.** The Cassini-Huygens spacecraft.

NASA's Cassini Mission-to-Saturn spacecraft is the first robotic mission ever to orbit the planet Saturn. Managed by the Jet Propulsion Laboratory (JPL) in Pasadena, this flagship-class mission is composed of 11 operating scientific instruments which study many intriguing features of Saturn, its moons, and ring system. The Cassini Program is an international cooperative effort involving primarily NASA, the European Space Agency (ESA), and the Italian Space Agency (Agenzia Spaziale Italiana, ASI). Cassini is the fourth spacecraft to visit the Saturnian system (but is the first vehicle to enter its orbit), and is composed of the NASA/ASI Cassini orbiter and the ESA-developed Huygens probe. Cassini launched on October 15, 1997, arriving at Saturn in 2004, after performing scientific observation of Earth's moon, Venus, and Jupiter (as well as participating in several scientific experiments) during its 6.7 year cruise period. Cassini's suite of (currently operating) science instruments consists of the following (**Figure 2**):


probe consisted of six scientific instruments which performed experiments in aerosol collection, descent imaging & spectral radiometry, gas chromatography & mass spectrometry, atmospheric

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During the cruise portion of the journey to the Saturnian system, two gravity assist maneuvers were required from Venus, one from Earth, and one from Jupiter. Until Cassini reached 2.7 AU from the sun (during the inner cruise phase), communications between earth and the spacecraft were accomplished via the Low Gain Antenna (LGA), since the 4-m diameter High Gain Antenna (HGA) must be used to shield the spacecraft from the sun's heating (i.e. used as a sunshade). After reaching this distance (begin Outer cruise phase), communications begin on

sampling, and surface science. The entire Cassini mission consists of seven phases:

• Launch and initial acquisition of the spacecraft (October 15, 1997)

• Inner cruise (beginning October 20, 1997) • Outer cruise (beginning February 2000)

• Science cruise (starting July 2002)

• Saturn Tour continues (2004–2017)

**Figure 3.** Cassini's prime, equinox XM, & solstice XXM tours.

the earth-pointed HGA.

• Saturn Orbit Insertion (SOI; July 2004) • Huygens Probe Release (January 2005)

**Figure 2.** Cassini's instrument suite.


Also included onboard Cassini is the Huygens Probe; an atmospheric laboratory designed to collect data in the Titan Moon atmosphere and its surface. Deployed in January 2005, the probe consisted of six scientific instruments which performed experiments in aerosol collection, descent imaging & spectral radiometry, gas chromatography & mass spectrometry, atmospheric sampling, and surface science. The entire Cassini mission consists of seven phases:


During the cruise portion of the journey to the Saturnian system, two gravity assist maneuvers were required from Venus, one from Earth, and one from Jupiter. Until Cassini reached 2.7 AU from the sun (during the inner cruise phase), communications between earth and the spacecraft were accomplished via the Low Gain Antenna (LGA), since the 4-m diameter High Gain Antenna (HGA) must be used to shield the spacecraft from the sun's heating (i.e. used as a sunshade). After reaching this distance (begin Outer cruise phase), communications begin on the earth-pointed HGA.

**Figure 3.** Cassini's prime, equinox XM, & solstice XXM tours.

**7.** Dual Technique Magnetometer (MAG)

**9.** Radio & Plasma Wave Science instrument (RPWS)

Also included onboard Cassini is the Huygens Probe; an atmospheric laboratory designed to collect data in the Titan Moon atmosphere and its surface. Deployed in January 2005, the

**8.** Cosmic Dust Analyzer (CDA)

**Figure 2.** Cassini's instrument suite.

**10.**Radio Science Subsystem (RSS)

**11.**Radar

226 Space Flight

Cassini's "Prime Tour Mission" began in 2004, where planet/moon science investigation activities continued until 2008. Two mission extensions were granted: the "Equinox Mission" from 2008 to 2010, and the "Solstice Mission" from 2010 to 2017 (**Figure 3**, [1]). The spacecraft's 20 year mission ends with 42 orbits around the main ring system (**Figure 4**, [2]). Beginning on November 30, 2016, Cassini's orbit reoriented the spacecraft to the outer edge of the main rings to perform a series of 20 F-Ring orbits; a region of Saturn's rings which look like an odd "interwoven" structure. The last time that Cassini observed these rings close-up was at Saturn arrival in 2004, which allowed observation of only the dim, backlit side. But in November of 2016, numerous opportunities became available to examine the F-Ring's structure, with high-resolution observation of both sides of the F-Ring. The final mission phase called "The Grand Finale" began in April 2017 with a close flyby of Saturn's giant moon Titan, which provided re-orientation of the spacecraft's trajectory, allowing it to pass through the gap between Saturn and the D-Ring; the closest ring to the planet. With only a 1500 mile-wide corridor to fly through, Cassini will investigate this unexplored region of the Saturnian system, making the closest observations of Saturn to date. During these last 22 (D-Ring) orbits of the Cassini mission, the planet's magnetic and gravity fields will be mapped with high precision, and extremely close views of the atmosphere will be observed. New insights into Saturn's interior structure, the precise length of a Saturnian day, and the age and total mass of the rings will also be evaluated. On September 15, 2017, Cassini will end its 20 year mission with a fiery plunge into Saturn, providing valuable data about the planet's chemical composition as the friction forces (from the atmospheric entry) cause the vehicle to burn up, thus satisfying Planetary Protection requirements [3].

**3. The Cassini radio communications system**

**Figure 5.** The Cassini spacecraft.

Cassini's onboard telecommunications system consists of three antennas: a High-Gain Antenna and two Low-Gain Antennas (LGA-1 & LGA-2); all which interface with the RFS system (which performs command, telemetry, and radio-metric communications) and Radio Frequency Instrument Subsystem (RFIS); **Figure 5**. Cassini's 4-m Cassegrain HGA

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**Figure 4.** Cassini's end-of-Mission F & D Ring Orbits.

#### **3. The Cassini radio communications system**

Cassini's "Prime Tour Mission" began in 2004, where planet/moon science investigation activities continued until 2008. Two mission extensions were granted: the "Equinox Mission" from 2008 to 2010, and the "Solstice Mission" from 2010 to 2017 (**Figure 3**, [1]). The spacecraft's 20 year mission ends with 42 orbits around the main ring system (**Figure 4**, [2]). Beginning on November 30, 2016, Cassini's orbit reoriented the spacecraft to the outer edge of the main rings to perform a series of 20 F-Ring orbits; a region of Saturn's rings which look like an odd "interwoven" structure. The last time that Cassini observed these rings close-up was at Saturn arrival in 2004, which allowed observation of only the dim, backlit side. But in November of 2016, numerous opportunities became available to examine the F-Ring's structure, with high-resolution observation of both sides of the F-Ring. The final mission phase called "The Grand Finale" began in April 2017 with a close flyby of Saturn's giant moon Titan, which provided re-orientation of the spacecraft's trajectory, allowing it to pass through the gap between Saturn and the D-Ring; the closest ring to the planet. With only a 1500 mile-wide corridor to fly through, Cassini will investigate this unexplored region of the Saturnian system, making the closest observations of Saturn to date. During these last 22 (D-Ring) orbits of the Cassini mission, the planet's magnetic and gravity fields will be mapped with high precision, and extremely close views of the atmosphere will be observed. New insights into Saturn's interior structure, the precise length of a Saturnian day, and the age and total mass of the rings will also be evaluated. On September 15, 2017, Cassini will end its 20 year mission with a fiery plunge into Saturn, providing valuable data about the planet's chemical composition as the friction forces (from the atmospheric entry) cause the vehicle to burn up, thus satisfying Planetary

Protection requirements [3].

228 Space Flight

**Figure 4.** Cassini's end-of-Mission F & D Ring Orbits.

Cassini's onboard telecommunications system consists of three antennas: a High-Gain Antenna and two Low-Gain Antennas (LGA-1 & LGA-2); all which interface with the RFS system (which performs command, telemetry, and radio-metric communications) and Radio Frequency Instrument Subsystem (RFIS); **Figure 5**. Cassini's 4-m Cassegrain HGA

**Figure 5.** The Cassini spacecraft.

communicates with earth on X-band, and on S-Band with the Huygens probe (and radioscience). It also communicates on Ka-band to support radio science activities, and Ku-band for the imaging RADAR subsystem. The two LGA antennas operate on X-band only, with LGA-1 mounted on the top of the HGA (giving it an unobstructed field of view of 112°), and LGA-2 which is mounted on a boom below the Huygens probe near the bottom of the vehicle, yielding a 120° field of view. The LGA antennas were used for communication with the ground when the HGA could not be configured on earth-point due to thermal constraints (when in close proximity to the sun). In this case, the spacecraft had to be shielded by the HGA, leaving the LGA antennas to transmit and receive data at very low delivery rates. The LGA antennas are configured when FP executes.

being received from the ground). Precise tracking of the spacecraft is accomplished through this method, as well as the ability to carry out high precision science experiments onboard

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Cassini carries its own Ultra-Stable Oscillator (USO). During the one-way phase when the spacecraft transmits its signal to the ground (before two-way communication is established), the spacecraft must generate its own D/L signal using the on-board USO. Once the ground's U/L signal is acquired by the vehicle, it will abandon its own D/L signal to regenerate the D/L, thus changing the frequency. During this time, the ground station will "lose lock" on the spacecraft and must tune in the new frequency. This "out-of-lock" condition is predetermined by the ground (on the order of a minute or two), so that data delivery to the ground is temporarily halted during this transition period, in order to preserve the precious science data. The USO device is quite reliable in generating a stable D/L signal, more so than the 2-way method with the ground, because the ground U/L signal phase is subject to corruption by atmospheric effects, solar wind, etc. Therefore, the USO is more desirable than the hydrogen maser. However, the USO frequency cannot be precisely known if the D/L frequency changes due to relative motion of the spacecraft (as well as vehicle drifting). Since ranging is fundamentally a phase measurement, the ground must use the hydrogen maser referenced U/L along with phase coherent receivers on the spacecraft and on the ground to determine

**4. Cassini mission telecommunications operations in flight**

NASA's DSN is a part of JPL, consisting of a worldwide network of US spacecraft communication facilities. Placed approximately 120° apart around the Earth, three deep-space telecommunications stations are located in Goldstone, California (US), Madrid, Spain, and Canberra, Australia. The placement of these ground stations permits constant observation of spacecraft like Cassini as the Earth rotates. Unlike near-earth orbiters which move quickly round the earth, few ground stations are required to support deep space missions since they are visible for long periods of time. As mentioned before, these earth-based DSN ground stations contain steerable, high-gain, parabolic reflector antennas, providing a two-way communications link that tracks robotic interplanetary spacecraft like Cassini, acquiring telemetry data, transmitting commands, uploading software modifications, tracking spacecraft position and velocity, measuring variations in radio waves to support radio science experiments, and collecting science & engineering data. Interplanetary spacecraft such as Cassini, require huge DSN antennas with ultra-sensitive receivers and powerful transmitters in order to transmit/ receive information over the vast earth-planet distances, with the largest antennas of the DSN often called upon during spacecraft emergencies. Nearly all spacecraft are designed to use the smaller DSN antennas (e.g. 34 m diameter) for nominal operations, but for a spacecraft emergency, the largest antennas are typically used (e.g. 70 m diameter) since the onboard FP typically configures low transmitter power, so that recovering any available telemetry is crucial to

the orbiter.

the correct measurement.

Spacecraft are typically equipped with transmitters of relatively low radiating power for communication with earth (20 Watts for Cassini). This telecommunications link must bridge the distance of over a billion kilometers (earth-Saturn distance), which is achieved by employing frequencies in the microwave range using reflectors onboard the spacecraft to concentrate all available power into a narrow beam pointed precisely towards earth. Cassini's HGA is used to achieve this goal (as opposed to the LGA antennas which sacrifice gain but provide relatively uniform coverage over a wide range of spacecraft orientation angles). At the DSN station, large aperture Cassegrain reflectors are used to pick up the spacecraft's signal. These radio antennas use cryogenically cooled (low-noise) amplifiers to first amplify the faint spacecraft signal, followed by sophisticated receivers and decoders which can lock onto and extract the data with virtually with no errors at all.

The signal delivered from the spacecraft to earth's ground station is called a "downlink," and the transmission of commands and sequences from the ground to the spacecraft is called an "uplink." When a D/L signal is received from the spacecraft, the communication is called "one-way" (or if the D/L signal is generated onboard the spacecraft itself, the communication is also called "one-way"). When the U/L signal is being received by the spacecraft at the same time a D/L is being received by the ground station, the communication is called "two-way." Both U/L and D/L consist of a pure Radio Frequency (RF) tone which is called a "carrier." In order to carry information to or from the spacecraft, the carrier signal must be "modulated." A modulated signal may be sent from the ground station to transmit commands to the spacecraft. Likewise, the modulated signal is generated by the spacecraft to transmit science and engineering data to earth on its D/L carrier. The spacecraft's carrier signal is also used for tracking and navigation (as well as some types of science experiments such as radio science or gravity field mapping). Each DSN complex uses a hydrogen-maser-based frequency unit which is maintained in an environmentally controlled room (in the basement), sustained by an uninterruptable power supply. The maser serves as the reference for generating a precisely known U/L frequency. When an U/L signal is received by the spacecraft, it can choose to use the received U/L carrier to control its D/L carrier transmission (called 2-way coherent transmission). This ground-generated reference frequency is multiplied by a predetermined constant (1.1748999 for Cassini) and the transmitted D/L signal is phase coherent with the U/L signal (this multiplier prevents the D/L signal from interfering with the U/L signal which is being received from the ground). Precise tracking of the spacecraft is accomplished through this method, as well as the ability to carry out high precision science experiments onboard the orbiter.

communicates with earth on X-band, and on S-Band with the Huygens probe (and radioscience). It also communicates on Ka-band to support radio science activities, and Ku-band for the imaging RADAR subsystem. The two LGA antennas operate on X-band only, with LGA-1 mounted on the top of the HGA (giving it an unobstructed field of view of 112°), and LGA-2 which is mounted on a boom below the Huygens probe near the bottom of the vehicle, yielding a 120° field of view. The LGA antennas were used for communication with the ground when the HGA could not be configured on earth-point due to thermal constraints (when in close proximity to the sun). In this case, the spacecraft had to be shielded by the HGA, leaving the LGA antennas to transmit and receive data at very low delivery rates. The

Spacecraft are typically equipped with transmitters of relatively low radiating power for communication with earth (20 Watts for Cassini). This telecommunications link must bridge the distance of over a billion kilometers (earth-Saturn distance), which is achieved by employing frequencies in the microwave range using reflectors onboard the spacecraft to concentrate all available power into a narrow beam pointed precisely towards earth. Cassini's HGA is used to achieve this goal (as opposed to the LGA antennas which sacrifice gain but provide relatively uniform coverage over a wide range of spacecraft orientation angles). At the DSN station, large aperture Cassegrain reflectors are used to pick up the spacecraft's signal. These radio antennas use cryogenically cooled (low-noise) amplifiers to first amplify the faint spacecraft signal, followed by sophisticated receivers and decoders which can lock onto and extract

The signal delivered from the spacecraft to earth's ground station is called a "downlink," and the transmission of commands and sequences from the ground to the spacecraft is called an "uplink." When a D/L signal is received from the spacecraft, the communication is called "one-way" (or if the D/L signal is generated onboard the spacecraft itself, the communication is also called "one-way"). When the U/L signal is being received by the spacecraft at the same time a D/L is being received by the ground station, the communication is called "two-way." Both U/L and D/L consist of a pure Radio Frequency (RF) tone which is called a "carrier." In order to carry information to or from the spacecraft, the carrier signal must be "modulated." A modulated signal may be sent from the ground station to transmit commands to the spacecraft. Likewise, the modulated signal is generated by the spacecraft to transmit science and engineering data to earth on its D/L carrier. The spacecraft's carrier signal is also used for tracking and navigation (as well as some types of science experiments such as radio science or gravity field mapping). Each DSN complex uses a hydrogen-maser-based frequency unit which is maintained in an environmentally controlled room (in the basement), sustained by an uninterruptable power supply. The maser serves as the reference for generating a precisely known U/L frequency. When an U/L signal is received by the spacecraft, it can choose to use the received U/L carrier to control its D/L carrier transmission (called 2-way coherent transmission). This ground-generated reference frequency is multiplied by a predetermined constant (1.1748999 for Cassini) and the transmitted D/L signal is phase coherent with the U/L signal (this multiplier prevents the D/L signal from interfering with the U/L signal which is

LGA antennas are configured when FP executes.

230 Space Flight

the data with virtually with no errors at all.

Cassini carries its own Ultra-Stable Oscillator (USO). During the one-way phase when the spacecraft transmits its signal to the ground (before two-way communication is established), the spacecraft must generate its own D/L signal using the on-board USO. Once the ground's U/L signal is acquired by the vehicle, it will abandon its own D/L signal to regenerate the D/L, thus changing the frequency. During this time, the ground station will "lose lock" on the spacecraft and must tune in the new frequency. This "out-of-lock" condition is predetermined by the ground (on the order of a minute or two), so that data delivery to the ground is temporarily halted during this transition period, in order to preserve the precious science data. The USO device is quite reliable in generating a stable D/L signal, more so than the 2-way method with the ground, because the ground U/L signal phase is subject to corruption by atmospheric effects, solar wind, etc. Therefore, the USO is more desirable than the hydrogen maser. However, the USO frequency cannot be precisely known if the D/L frequency changes due to relative motion of the spacecraft (as well as vehicle drifting). Since ranging is fundamentally a phase measurement, the ground must use the hydrogen maser referenced U/L along with phase coherent receivers on the spacecraft and on the ground to determine the correct measurement.

#### **4. Cassini mission telecommunications operations in flight**

NASA's DSN is a part of JPL, consisting of a worldwide network of US spacecraft communication facilities. Placed approximately 120° apart around the Earth, three deep-space telecommunications stations are located in Goldstone, California (US), Madrid, Spain, and Canberra, Australia. The placement of these ground stations permits constant observation of spacecraft like Cassini as the Earth rotates. Unlike near-earth orbiters which move quickly round the earth, few ground stations are required to support deep space missions since they are visible for long periods of time. As mentioned before, these earth-based DSN ground stations contain steerable, high-gain, parabolic reflector antennas, providing a two-way communications link that tracks robotic interplanetary spacecraft like Cassini, acquiring telemetry data, transmitting commands, uploading software modifications, tracking spacecraft position and velocity, measuring variations in radio waves to support radio science experiments, and collecting science & engineering data. Interplanetary spacecraft such as Cassini, require huge DSN antennas with ultra-sensitive receivers and powerful transmitters in order to transmit/ receive information over the vast earth-planet distances, with the largest antennas of the DSN often called upon during spacecraft emergencies. Nearly all spacecraft are designed to use the smaller DSN antennas (e.g. 34 m diameter) for nominal operations, but for a spacecraft emergency, the largest antennas are typically used (e.g. 70 m diameter) since the onboard FP typically configures low transmitter power, so that recovering any available telemetry is crucial to assessing the spacecraft's health in preparation for recovery actions. In the case of Cassini, the LGA is configured by FP with very low U/L & D/L rates.

pass is in progress with Cassini transmitting its telemetry data. Thousands of engineering telemetry measurements (i.e. temperatures, voltages, pressures, computer statuses detailing

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An "out of lock" condition can occur suddenly if Cassini's signal strength drops out (LOS condition). This can be caused by rain at the DSN station from too many water molecules in the vicinity of the antenna which give off an abundance of radio noise that can literally drown out the spacecraft's signal. In this case, the "DSN Receiver Status" on the Cassini ACE's console will light up with an "OUT OF LOCK" reading. The measured system operating noise temperature on the console should rise high enough to indicate that rain is the reason for the signal loss. But if bad weather is not the cause of the LOS condition, or caused by an unforeseen problem in the ground system equipment itself, the ACE will contact the Operations Chief (who is concurrently working with the Cassini ACE at JPL), to request that a second DSN antenna look for the spacecraft's signal, if available. If no signal is detected, the Cassini ACE will declare a "LOS condition" and proceed to follow the "LOS/Anomalous Downlink Contingency Plan" Procedure which requires that he/she contact the appropriate SOFS team members. These are spacecraft subsystem experts who must evaluate the situation and concur with the Cassini ACE that there is no earth-based problem causing the LOS condition (ground station or weather). In this case, the most likely explanation is that an onboard RFS-related FP routine has triggered. Numerous fault monitors are installed into Cassini's FSW that are constantly running to detect faults in spacecraft systems. Upon fault detection, a "canned" response routine(s) is executed autonomously to fix the problem, which is typically followed by an activation of the Safing Response. This response places the spacecraft in a predictable state, configuring lower power consumption with low U/L and D/L rates on LGA, commanding the HGA to sun-point (off earth-point). In the case of a RFS FP routine activation, the RFS device states might be altered, as a swap to a redundant RFS unit is commanded which changes the telecommunications configuration for

The ACE knows that Cassini will have transitioned from the HGA to the LGA antenna, should the FP activate. The LGA provides an extremely weak D/L signal since its beamwidth is much larger than the HGA beamwidth. At Saturn, the spacecraft's signal is so weak that telemetry delivery is only possible at 5 bps, requiring nearly 18 hours to receive all 30 decks of telemetry data that are needed for the SOFS team members to verify the spacecraft's health and determine its post-fault states. Recovery from any fault is extremely slow, but if no attitude control system problems are present and spacecraft attitude knowledge is preserved (no faults in the AACS computers), a second FP routine called the "High Gain Antenna Swap (HAS) Response" will automatically activate 1 hour after the Safing Response concludes. This FP will increase the U/L and D/L rates (D/L = 1896 bps), followed by a turn of the spacecraft's HGA to earth-point. In this configuration, all 30 decks of telemetry data are delivered to the ground in approximately 10 minutes, making recovery from the fault much more expedient. For typical FP activations, the SOFS team will examine the spacecraft telemetry

the vehicle's health and status) are interleaved with the science data.

**4.2. Anomalous D/L conditions**

D/L signal acquisition.

Ground commands from earth travel at the speed of light (referred to as "One-Way Light Time;" OWLT), reaching Cassini from approximately 1 hour. 15 minute to 1 hour 30 minute, depending on the relative distance between earth and Saturn, given the change in relative distance due to the earth's rotation around the sun and the spacecraft's motion around the Saturnian system. Therefore, the majority of commands sent to the spacecraft for operations and science investigations must be uplinked to the Command & Data processing System (CDS) computers in large "command sequences," which consist of several weeks of planned commanding. These sequences typically consist of commanded turns to point Cassini's 11 operating instruments towards specific targets, providing high precision (down to the sub-milliradian) via two Attitude, Articulation, & Control System (AACS) computers. Captured science data is recorded on two Solid-State Recorders (SSR) during off-earth observation periods. These science activities (e.g. moon and ring encounters) are paused typically once each day (or two) for approximately 9 hours to establish communication with earth (via a scheduled DSN station) to downlink the science & engineering (housekeeping) data.

Once Cassini's earth-pointed attitude is stabilized, its D/L signal is received by the DSN station. Ten minutes later, the ACE initiates the U/L signal for commanding and navigational purposes. The data is transmitted from the spacecraft in the format of "symbols" which are "wiggles" in Cassini's radio signal's phase. The DSN receives the symbols and decodes it into "0" and "1 seconds" in order to reconstruct the telemetry data (engineering housekeeping data, science digital images, etc.). After the 9 hours of telemetry data have been downlinked to earth's DSN ground station, the spacecraft reduces its data rate, suspends its data playback (from the SSRs), and turns to the next science target via the onboard running sequence to collect new science data [4].

#### **4.1. Nominal S/C acquisition**

Prior to spacecraft acquisition at JPL's Space Flight Operations building in Pasadena, California, the "Cassini ACE" Real-time Operations Engineer must prepare to receive the data transmission stream from Cassini, and is in voice contact with the DSN station staff (in California, Australia, or Spain). The Cassini ACE provides their station operator with a 2 minute briefi ng to review the expected events for the day, before the DSN pass starts (any planned Reaction Control System (RCS) burns or Main Engine (ME) maneuvers, Flight Software (FSW) patches or uploads, etc.) and provides any pertinent updates. The DSN station operator, in turn, provides a weather report (clear skies or rain, plus wind conditions) and that all equipment is in working order (green), or has suffered a system breakdown (red). The designated (34 m or 70 m) antenna at the DSN station for the day's 9 hour pass has already been pointed precisely towards Saturn where Cassini's faint signal will be received. Once the spacecraft's signal has been acquired, the DSN station operator reports to the Cassini Ace that the station's receiver is "in lock." The Cassini Ace then acknowledges that the telemetry at his/her workstation is being received and looks nominal. From this point, the 9 hour DSN pass is in progress with Cassini transmitting its telemetry data. Thousands of engineering telemetry measurements (i.e. temperatures, voltages, pressures, computer statuses detailing the vehicle's health and status) are interleaved with the science data.

#### **4.2. Anomalous D/L conditions**

assessing the spacecraft's health in preparation for recovery actions. In the case of Cassini, the

Ground commands from earth travel at the speed of light (referred to as "One-Way Light Time;" OWLT), reaching Cassini from approximately 1 hour. 15 minute to 1 hour 30 minute, depending on the relative distance between earth and Saturn, given the change in relative distance due to the earth's rotation around the sun and the spacecraft's motion around the Saturnian system. Therefore, the majority of commands sent to the spacecraft for operations and science investigations must be uplinked to the Command & Data processing System (CDS) computers in large "command sequences," which consist of several weeks of planned commanding. These sequences typically consist of commanded turns to point Cassini's 11 operating instruments towards specific targets, providing high precision (down to the sub-milliradian) via two Attitude, Articulation, & Control System (AACS) computers. Captured science data is recorded on two Solid-State Recorders (SSR) during off-earth observation periods. These science activities (e.g. moon and ring encounters) are paused typically once each day (or two) for approximately 9 hours to establish communication with earth (via a scheduled DSN station) to

Once Cassini's earth-pointed attitude is stabilized, its D/L signal is received by the DSN station. Ten minutes later, the ACE initiates the U/L signal for commanding and navigational purposes. The data is transmitted from the spacecraft in the format of "symbols" which are "wiggles" in Cassini's radio signal's phase. The DSN receives the symbols and decodes it into "0" and "1 seconds" in order to reconstruct the telemetry data (engineering housekeeping data, science digital images, etc.). After the 9 hours of telemetry data have been downlinked to earth's DSN ground station, the spacecraft reduces its data rate, suspends its data playback (from the SSRs), and turns to the next science target via the onboard running sequence to col-

Prior to spacecraft acquisition at JPL's Space Flight Operations building in Pasadena, California, the "Cassini ACE" Real-time Operations Engineer must prepare to receive the data transmission stream from Cassini, and is in voice contact with the DSN station staff (in California, Australia, or Spain). The Cassini ACE provides their station operator with a 2 minute briefi ng to review the expected events for the day, before the DSN pass starts (any planned Reaction Control System (RCS) burns or Main Engine (ME) maneuvers, Flight Software (FSW) patches or uploads, etc.) and provides any pertinent updates. The DSN station operator, in turn, provides a weather report (clear skies or rain, plus wind conditions) and that all equipment is in working order (green), or has suffered a system breakdown (red). The designated (34 m or 70 m) antenna at the DSN station for the day's 9 hour pass has already been pointed precisely towards Saturn where Cassini's faint signal will be received. Once the spacecraft's signal has been acquired, the DSN station operator reports to the Cassini Ace that the station's receiver is "in lock." The Cassini Ace then acknowledges that the telemetry at his/her workstation is being received and looks nominal. From this point, the 9 hour DSN

LGA is configured by FP with very low U/L & D/L rates.

downlink the science & engineering (housekeeping) data.

lect new science data [4].

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**4.1. Nominal S/C acquisition**

An "out of lock" condition can occur suddenly if Cassini's signal strength drops out (LOS condition). This can be caused by rain at the DSN station from too many water molecules in the vicinity of the antenna which give off an abundance of radio noise that can literally drown out the spacecraft's signal. In this case, the "DSN Receiver Status" on the Cassini ACE's console will light up with an "OUT OF LOCK" reading. The measured system operating noise temperature on the console should rise high enough to indicate that rain is the reason for the signal loss. But if bad weather is not the cause of the LOS condition, or caused by an unforeseen problem in the ground system equipment itself, the ACE will contact the Operations Chief (who is concurrently working with the Cassini ACE at JPL), to request that a second DSN antenna look for the spacecraft's signal, if available. If no signal is detected, the Cassini ACE will declare a "LOS condition" and proceed to follow the "LOS/Anomalous Downlink Contingency Plan" Procedure which requires that he/she contact the appropriate SOFS team members. These are spacecraft subsystem experts who must evaluate the situation and concur with the Cassini ACE that there is no earth-based problem causing the LOS condition (ground station or weather). In this case, the most likely explanation is that an onboard RFS-related FP routine has triggered. Numerous fault monitors are installed into Cassini's FSW that are constantly running to detect faults in spacecraft systems. Upon fault detection, a "canned" response routine(s) is executed autonomously to fix the problem, which is typically followed by an activation of the Safing Response. This response places the spacecraft in a predictable state, configuring lower power consumption with low U/L and D/L rates on LGA, commanding the HGA to sun-point (off earth-point). In the case of a RFS FP routine activation, the RFS device states might be altered, as a swap to a redundant RFS unit is commanded which changes the telecommunications configuration for D/L signal acquisition.

The ACE knows that Cassini will have transitioned from the HGA to the LGA antenna, should the FP activate. The LGA provides an extremely weak D/L signal since its beamwidth is much larger than the HGA beamwidth. At Saturn, the spacecraft's signal is so weak that telemetry delivery is only possible at 5 bps, requiring nearly 18 hours to receive all 30 decks of telemetry data that are needed for the SOFS team members to verify the spacecraft's health and determine its post-fault states. Recovery from any fault is extremely slow, but if no attitude control system problems are present and spacecraft attitude knowledge is preserved (no faults in the AACS computers), a second FP routine called the "High Gain Antenna Swap (HAS) Response" will automatically activate 1 hour after the Safing Response concludes. This FP will increase the U/L and D/L rates (D/L = 1896 bps), followed by a turn of the spacecraft's HGA to earth-point. In this configuration, all 30 decks of telemetry data are delivered to the ground in approximately 10 minutes, making recovery from the fault much more expedient. For typical FP activations, the SOFS team will examine the spacecraft telemetry for off-nominal conditions, sometimes reading out additional sections of Cassini's computer memory to confirm the diagnosis, and then prepare commands for the ACE to send which will recover the spacecraft from the FP activation, and restart the onboard running sequence once again.

**5. Cassini LOS experiences**

command the device on (see **Table 1**).

**Table 1.** SSPS trip FP for USO trip (post-2006).

trip event [7].

Cassini has experienced several LOS events during its mission lifetime. Some events have been caused by relatively minor problems, but two events are of significance. The first occurred on May 1, 2006. At the beginning of the DSN track, the DSN station was unable to acquire the spacecraft's "one-way" carrier signal (i.e. the ground-received spacecraft signal), which in turn, initiated the anomaly response process. However, after Round-Trip-Light-Time (RTLT; twice OWLT) had elapsed, the DSN station was able to lock up on the "two-way" carrier signal and the spacecraft's data. Telemetry indicated that Cassini's USO had suffered an SSPS

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Cassini's power system consists of power control boards which contain 192 SSPS. SSPS trip events occur spuriously and without warning, on average 2–3 times per year due to the unforeseen environmental effects of galactic cosmic ray bombardment [8]. This condition is thought to be caused by one or more photon hits on the voltage comparator of the device, resulting in a false indication that the current load is anomalously high, thus tripping off the switch. Because of this phenomenon, a new "SSPS Trip FP" monitor & response algorithm was uploaded to the Cassini spacecraft's FSW. The monitor examines one SSPS switch state per second, (starting with switch number 1), and proceeds through all 192 SSPS switches. If a SSPS trip is detected, the response contains a table of appropriate actions for FP to act upon, based upon the specific SSPS switch and its function. The actions of the original SSPS FP response table for the USO (uplinked prior to 2006) only recorded the USO trip event (USO SSPS is #68) and cleared the tripped condition by commanding the unit OFF. However, after this USO trip event occurred, the response table was augmented (via uplink command) to

Five years later on December 23, 2011 at the Beginning of the DSN Track (BOT), once again, no D/L signal was seen from the Cassini spacecraft. The DSN station at Canberra was supporting Cassini at the time. Following ACE direction, additional tracking was obtained using a Canberra station antenna, as well as a Goldstone station antenna, but without successful acquisition of the spacecraft's signal. The SOFS Anomaly Team was called together to diagnose the problem. At RTLT, Cassini was once again acquired in 2-way mode, confirming that the problem was with

In certain cases, complete LOS can occur. Resolution of a LOS fault may require extra DSN coverage, depending on the difficulty in determining the fault cause. As mentioned previously, the Cassini ACE also looks for other DSN tracks that can be borrowed from other flight projects or scheduled maintenance for the next few days. If the LOS condition persists, a "spacecraft emergency" will be declared to guarantee continuous DSN coverage to support spacecraft recovery efforts.

#### **4.3. No spacecraft signal acquisition (LOS)**

Unlike most faults that trigger the onboard FP, a fault causing total LOS means no acquisition of the spacecraft's signal at all (i.e. no lock-up on the expected or post-FP RFS configuration) by the DSN station. There are several reasons why a LOS condition can occur. These include DSN station breakdowns, misconfigured lock-up parameters, or even faults which are not detected by the FP design. Unfortunately, not every spacecraft fault case can be precluded by the onboard FP. In spite of the best efforts of pre-launch designers to identify all possible fault scenarios and produce a FP system to support them (detect, isolate, & resolve), certain failure modes are sometimes missed or are very difficult to avoid. Most JPL projects like Cassini strive to meet a "Single Point Failure" (SPF) policy [5], but certain failures cannot be easily detected, or are not identified during the design phase, and some failures can actually occur even though they have been exempted or waived [6]. Other LOS fault possibilities are problems that occur in devices which are intentionally not protected by the onboard FP. These devices include the HGA or LGA antennas, Waveguide Transfer Switches (WTS), and the USO on Cassini. Multiple faults are also a possibility, since they do not fall under FP design guidelines due to the SPF policy.

Hence, LOS can occur from several sources: erroneous ground-generated commands uplinked to the spacecraft, onboard sequence failures, multiple failures which are not typically required to be addressed by the onboard FP, spacecraft pointing errors, failed telecom configurations (via ground commanding), internal FSW errors, computer platform failures, bad weather, or DSN ground equipment failures. Also, not only can RFS FP swap to redundant units due to device faults and malfunctions, thus inhibiting the ground from locking up on Cassini's signal (since the RFS D/L signal path has changed), but environmental effects can also cause a LOS condition. SSPS trip-off of RFS units (caused by cosmic ray bombardment) can also cause temporary loss of the spacecraft's signal. To address this condition, the Cassini ACE must perform several "uplink sweeps" on different variations of the RFS units in an attempt to re-acquire the spacecraft's D/L signal. Once ground problems and weather are ruled out as an LOS cause, the assumption is that hopefully the onboard FP has executed and commanded a RFS device swap to a redundant unit. Otherwise, determination of the fault cause becomes increasingly difficult to diagnose.

## **5. Cassini LOS experiences**

for off-nominal conditions, sometimes reading out additional sections of Cassini's computer memory to confirm the diagnosis, and then prepare commands for the ACE to send which will recover the spacecraft from the FP activation, and restart the onboard running sequence

In certain cases, complete LOS can occur. Resolution of a LOS fault may require extra DSN coverage, depending on the difficulty in determining the fault cause. As mentioned previously, the Cassini ACE also looks for other DSN tracks that can be borrowed from other flight projects or scheduled maintenance for the next few days. If the LOS condition persists, a "spacecraft emergency" will be declared to guarantee continuous DSN coverage to support

Unlike most faults that trigger the onboard FP, a fault causing total LOS means no acquisition of the spacecraft's signal at all (i.e. no lock-up on the expected or post-FP RFS configuration) by the DSN station. There are several reasons why a LOS condition can occur. These include DSN station breakdowns, misconfigured lock-up parameters, or even faults which are not detected by the FP design. Unfortunately, not every spacecraft fault case can be precluded by the onboard FP. In spite of the best efforts of pre-launch designers to identify all possible fault scenarios and produce a FP system to support them (detect, isolate, & resolve), certain failure modes are sometimes missed or are very difficult to avoid. Most JPL projects like Cassini strive to meet a "Single Point Failure" (SPF) policy [5], but certain failures cannot be easily detected, or are not identified during the design phase, and some failures can actually occur even though they have been exempted or waived [6]. Other LOS fault possibilities are problems that occur in devices which are intentionally not protected by the onboard FP. These devices include the HGA or LGA antennas, Waveguide Transfer Switches (WTS), and the USO on Cassini. Multiple faults are also a possibility, since they do not fall under FP design guidelines due to

Hence, LOS can occur from several sources: erroneous ground-generated commands uplinked to the spacecraft, onboard sequence failures, multiple failures which are not typically required to be addressed by the onboard FP, spacecraft pointing errors, failed telecom configurations (via ground commanding), internal FSW errors, computer platform failures, bad weather, or DSN ground equipment failures. Also, not only can RFS FP swap to redundant units due to device faults and malfunctions, thus inhibiting the ground from locking up on Cassini's signal (since the RFS D/L signal path has changed), but environmental effects can also cause a LOS condition. SSPS trip-off of RFS units (caused by cosmic ray bombardment) can also cause temporary loss of the spacecraft's signal. To address this condition, the Cassini ACE must perform several "uplink sweeps" on different variations of the RFS units in an attempt to re-acquire the spacecraft's D/L signal. Once ground problems and weather are ruled out as an LOS cause, the assumption is that hopefully the onboard FP has executed and commanded a RFS device swap to a redundant unit. Otherwise, determination of the fault cause becomes

once again.

234 Space Flight

the SPF policy.

increasingly difficult to diagnose.

spacecraft recovery efforts.

**4.3. No spacecraft signal acquisition (LOS)**

Cassini has experienced several LOS events during its mission lifetime. Some events have been caused by relatively minor problems, but two events are of significance. The first occurred on May 1, 2006. At the beginning of the DSN track, the DSN station was unable to acquire the spacecraft's "one-way" carrier signal (i.e. the ground-received spacecraft signal), which in turn, initiated the anomaly response process. However, after Round-Trip-Light-Time (RTLT; twice OWLT) had elapsed, the DSN station was able to lock up on the "two-way" carrier signal and the spacecraft's data. Telemetry indicated that Cassini's USO had suffered an SSPS trip event [7].

Cassini's power system consists of power control boards which contain 192 SSPS. SSPS trip events occur spuriously and without warning, on average 2–3 times per year due to the unforeseen environmental effects of galactic cosmic ray bombardment [8]. This condition is thought to be caused by one or more photon hits on the voltage comparator of the device, resulting in a false indication that the current load is anomalously high, thus tripping off the switch. Because of this phenomenon, a new "SSPS Trip FP" monitor & response algorithm was uploaded to the Cassini spacecraft's FSW. The monitor examines one SSPS switch state per second, (starting with switch number 1), and proceeds through all 192 SSPS switches. If a SSPS trip is detected, the response contains a table of appropriate actions for FP to act upon, based upon the specific SSPS switch and its function. The actions of the original SSPS FP response table for the USO (uplinked prior to 2006) only recorded the USO trip event (USO SSPS is #68) and cleared the tripped condition by commanding the unit OFF. However, after this USO trip event occurred, the response table was augmented (via uplink command) to command the device on (see **Table 1**).

Five years later on December 23, 2011 at the Beginning of the DSN Track (BOT), once again, no D/L signal was seen from the Cassini spacecraft. The DSN station at Canberra was supporting Cassini at the time. Following ACE direction, additional tracking was obtained using a Canberra station antenna, as well as a Goldstone station antenna, but without successful acquisition of the spacecraft's signal. The SOFS Anomaly Team was called together to diagnose the problem. At RTLT, Cassini was once again acquired in 2-way mode, confirming that the problem was with


**Table 1.** SSPS trip FP for USO trip (post-2006).

the spacecraft's USO. Commands were sent on Christmas Day to inhibit the USO and swap to the Auxiliary Oscillator as the frequency source for the D/L signal until the fault within the USO device could be evaluated.

has concluded. Attempted spacecraft recovery actions continue through each branch of the

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For a complete LOS condition, the Cassini ACE must perform the "uplink sweep" on the correct RFS device configuration to re-acquire the spacecraft [9]. The assumption is that the activation of a RFS FP response will have swapped to its counterpart unit, possibly changing the polarity of the D/L signal. Depending on the failure (or number of failures), several RFS device combinations are possible with variations on the following components, depending on which FP has activated

• DST-A/CDU-A or DST-B/CDU-B (Deep Space Transponder; Command Detector Unit)

**Figure 6** depicts the RFS Functional diagram for Cassini, whose prime RFS units are: DST-A/CDU-A, TCU-B, and TWTA-B; WTS-A used for U/L, WTS-B used for D/L. The use of these devices are

• WTS: provide switching capability for transmitting or receiving the signal through the

Also included in certain RFS FP response actions is a Power-on-Reset (POR) of the prime TCU and/or the Power subsystem where selected devices are turned off, reset, or reconfigured, which will select spacecraft components according to their own FP protocols. Further complicating the anomalous/LOS condition is the fact that RFS FP algorithms are multi-tiered (address several different fault types), and can activate at any time per their persistence counters (unique for each FP algorithm) which can range from seconds to minutes, further reconfiguring these device states after spacecraft re-acquisition is attempted, so that it is difficult to know which RFS combinations for the ACE to try (or which combinations should be re-tried or eliminated). Therefore, it is very important to keep track of when RFS related FP responses have timed out.

flowcharts until re-acquisition of the vehicle is successful (if possible).

and what the current RFS prime units are:

• TCU-A or TCU-B (Telemetry Control Unit)

listed below:

• WTS-A or WTS-B (Waveguide Transfer Switch)

• DST: is used for both the U/L and D/L function

• TWTA: is an amplifier used in the D/L function

• TCU: controls the RFS system.

HGA, LGA-1, or LGA-2 antennas.

• CDU: is part of the DST and used for the U/L function

• LGA-1 (LGA-2 is no longer in use) or HGA antenna

• Auxiliary Oscillator or DST VCO (Voltage-Controlled Oscillator)

• Auxiliary Oscillator: provides 1-way D/L carrier frequency reference.

• VCO: is part of the DST and provides 2-way D/L carrier frequency reference.

• TWTA-A or TWTA-B (Traveling Wave Tube Amplifire)

The next step for the SOFS team was to evaluate whether one-way operation of the USO was functioning properly (the two-way U/L must be halted in this case). Once configured, the DSN station was unable to lock onto Cassini's one-way signal which indicated that the USO was not operating properly. After a second attempt to establish the one-way link failed, a command was sent to inhibit the USO, allowing the Auxiliary Oscillator to take over again for spacecraft operations. Further tests conducted in January of 2012 confirmed that normal USO operation could not be re-established. After consulting with Radio Science and Applied Physics Laboratory (the builder of the USO), it was decided that the USO would be power cycled in an effort to "reset" the unit, although it was thought unlikely to work since the USO is an analog device. On January 9, 2013 the USO was powered OFF permanently and the Auxiliary Oscillator has been in operation ever since.

#### **6. LOS protocol**

For Cassini, addressing an "anomalous downlink" or LOS condition starts with the RFS Subsystem's "LOS/Anomalous Downlink Contingency Plan" Procedure to help identify possible reasons for the abnormal (or absence of) the spacecraft's D/L signal. This procedure describes possible troubleshooting methods and recovery actions needed for both offnominal signal levels (e.g. carrier power is too low or too high) as well as partial lock-up conditions (e.g. no subcarrier, symbol, telemetry, or frame lock-up), and complete LOS. The procedure provides diagnoses & recovery actions in the form of flowcharts for the ACE and SOFS Anomaly team members to follow. Five partial signal loss/LOS candidate faults are considered when determining required anomaly resolution actions:


RFS FP response actions are also noted in the recovery strategy flowcharts of the procedure and specify the expected post-fault RFS device states. Any attempt to re-acquire the spacecraft on the newly commanded RFS configuration is directly dependent on when the FP response has concluded. Attempted spacecraft recovery actions continue through each branch of the flowcharts until re-acquisition of the vehicle is successful (if possible).

For a complete LOS condition, the Cassini ACE must perform the "uplink sweep" on the correct RFS device configuration to re-acquire the spacecraft [9]. The assumption is that the activation of a RFS FP response will have swapped to its counterpart unit, possibly changing the polarity of the D/L signal. Depending on the failure (or number of failures), several RFS device combinations are possible with variations on the following components, depending on which FP has activated and what the current RFS prime units are:


the spacecraft's USO. Commands were sent on Christmas Day to inhibit the USO and swap to the Auxiliary Oscillator as the frequency source for the D/L signal until the fault within the USO

The next step for the SOFS team was to evaluate whether one-way operation of the USO was functioning properly (the two-way U/L must be halted in this case). Once configured, the DSN station was unable to lock onto Cassini's one-way signal which indicated that the USO was not operating properly. After a second attempt to establish the one-way link failed, a command was sent to inhibit the USO, allowing the Auxiliary Oscillator to take over again for spacecraft operations. Further tests conducted in January of 2012 confirmed that normal USO operation could not be re-established. After consulting with Radio Science and Applied Physics Laboratory (the builder of the USO), it was decided that the USO would be power cycled in an effort to "reset" the unit, although it was thought unlikely to work since the USO is an analog device. On January 9, 2013 the USO was powered OFF permanently and the Auxiliary Oscillator has

For Cassini, addressing an "anomalous downlink" or LOS condition starts with the RFS Subsystem's "LOS/Anomalous Downlink Contingency Plan" Procedure to help identify possible reasons for the abnormal (or absence of) the spacecraft's D/L signal. This procedure describes possible troubleshooting methods and recovery actions needed for both offnominal signal levels (e.g. carrier power is too low or too high) as well as partial lock-up conditions (e.g. no subcarrier, symbol, telemetry, or frame lock-up), and complete LOS. The procedure provides diagnoses & recovery actions in the form of flowcharts for the ACE and SOFS Anomaly team members to follow. Five partial signal loss/LOS candidate faults are

**1.** Spacecraft is not on earth-point when expected due to an incomplete turn, a fault in the AACS

**2.** DSN ground-station problem: station is not tracking the spacecraft properly, station receiver

**3.** Spacecraft telecom problem: there is a problem in the telecommunications system (error caused by the onboard sequence commanding, ground U/L commanding, or the FP has executed)

RFS FP response actions are also noted in the recovery strategy flowcharts of the procedure and specify the expected post-fault RFS device states. Any attempt to re-acquire the spacecraft on the newly commanded RFS configuration is directly dependent on when the FP response

considered when determining required anomaly resolution actions:

device could be evaluated.

236 Space Flight

been in operation ever since.

system, or FP activation.

is down, breakdowns, weather, etc.

**5.** Multiple faults or a catastrophic failure

**4.** Loss of the CDS (most likely a multi-fault condition)

**6. LOS protocol**


**Figure 6** depicts the RFS Functional diagram for Cassini, whose prime RFS units are: DST-A/CDU-A, TCU-B, and TWTA-B; WTS-A used for U/L, WTS-B used for D/L. The use of these devices are listed below:


Also included in certain RFS FP response actions is a Power-on-Reset (POR) of the prime TCU and/or the Power subsystem where selected devices are turned off, reset, or reconfigured, which will select spacecraft components according to their own FP protocols. Further complicating the anomalous/LOS condition is the fact that RFS FP algorithms are multi-tiered (address several different fault types), and can activate at any time per their persistence counters (unique for each FP algorithm) which can range from seconds to minutes, further reconfiguring these device states after spacecraft re-acquisition is attempted, so that it is difficult to know which RFS combinations for the ACE to try (or which combinations should be re-tried or eliminated). Therefore, it is very important to keep track of when RFS related FP responses have timed out.

of time during which no commands have been received by the spacecraft from the ground. The Command Loss Monitor is configured with a timer which counts down from a programmable value (usually days) until it reaches "0" seconds or is reset via ground command (on Cassini, this "Command Loss Timer" (CLT) is currently set to 115 hours). The receipt of a valid U/L command by the spacecraft will reset the timer to its original value and restart the countdown. This provides an end-to-end check on command functionality between the vehicle and the ground. If triggered (timer reaches "0"), the Command Loss Response will initiate an extended series of actions which are designed to re-enable ground commandability onboard the spacecraft. The response will attempt to command various telecom configurations and spacecraft attitudes in an attempt to find a viable uplink path. Each reconfiguration of a new uplink path is separated by an appropriate ground response interval for the SOFS

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**Figure 7** illustrates Cassini's Command Loss Response chain. Once triggered, it progresses through a series of "Command Groups" divided by multi-hour "Command Pauses" which allow the SOFS team to react by sending an U/L command to halt the response. The Command Groups consist of actions to reconfigure redundant hardware and re-command spacecraft attitude and antennas. Each Command Pause allow several hours for the SOFS team to attempt re-acquisition of the spacecraft upon the newly commanded spacecraft configuration (the pause durations are set to a minimum of two RTLT periods). As shown in the figure, the first Command Group will select the Auxiliary Oscillator and execute the Safing Response which will turn off non-essential spacecraft loads, place the spacecraft in a lower power state, and re-direct the spacecraft's High Gain Antenna to sun-point, placing the spacecraft in a low U/L & D/L state through the LGA-1 antenna. After the first Command Group has executed, a 15 hour wait period (Command Pause) allows sufficient time for the SOFS Anomaly team to assemble at JPL and attempt re-establishment of the U/L, if possible, before RFS hardware swaps begin in successive Command Groups. If the re-acquisition attempt fails after Command Group #1 execution, the response will proceed with the next

> **Cmd Group #6 Power On Redundant BU CDS Computer**

> **Cmd Group #6 Power On Redundant BU CDS Computer**

**Cmd Group #7 Swap Antennas & Point HGA to Earth**

*7h 7hrs 7hrs rs*

**Cmd Group #7 Swap Antennas & Point HGA to Earth**

*7h 7hrs 7hrs rs* **Swap CDS**

**Cmd Group #8 Swap to LGA & Power offUSO**

**Cmd Group #8 Swap to LGA & Power offUSO**

*7hrs*

**Computers**

*7days 15hrs 29min*

team to re-acquire the spacecraft.

**Cmd Grps: 2-4 Perform RFS Uni t Swaps (DSTs, TWTAs, TCUs)**

*15hrs 7hrs*

**Cmd Grps:2-4 Perform RFS Uni t Swaps (DSTs, TWTAs, TCUs)**

*49hrs*

*7hrs*

**Figure 7.** Cmdloss response actions.

**Cmd Group #5 Point HGAto the Sun with Constant Roll Rate**

**Cmd Group #5 Point HGAto the Sun with Constant Roll Rate**

*7hrs*

**Cmd Group #1 Swap to Auxiliary Oscillator; Run Safing**

**Enter (CLT=0)**

**Figure 6.** Cassini's RFS functional diagram.

#### **7. Command loss FP**

An unresolvable LOS condition where the ground is no longer able to deliver commands to the spacecraft will eventually lead to the activation of a LOS FP response. The actions of this response can help to re-establish the U/L. In Cassini's FP design, loss of D/L fault coverage is not protected in an "end-to-end" manner since the D/L is not considered to be a critical spacecraft function which requires autonomous restoration. But restoration of the U/L however, is considered crucial to mission success and is therefore allocated "endto-end" protection through a "Loss of Commandability" algorithm [10]. Although several other (higher priority) FP routines are installed into Cassini's FP suite to protect against these same type of failures in the U/L path (which provide more timely action), the Loss of Commandability algorithm provides a "safety net" type of FP which has the potential to restore both U/L and D/L. With this scheme in place, multiple levels of FP defense are provided (covering up to 3 faults).

This catch-all type of FP is referred to as a "Command Loss FP" (from the perspective of the spacecraft since it is no longer receiving ground commands) and is typically an "endlessloop" response. The Command Loss Monitor aboard Cassini will detect an extended period of time during which no commands have been received by the spacecraft from the ground. The Command Loss Monitor is configured with a timer which counts down from a programmable value (usually days) until it reaches "0" seconds or is reset via ground command (on Cassini, this "Command Loss Timer" (CLT) is currently set to 115 hours). The receipt of a valid U/L command by the spacecraft will reset the timer to its original value and restart the countdown. This provides an end-to-end check on command functionality between the vehicle and the ground. If triggered (timer reaches "0"), the Command Loss Response will initiate an extended series of actions which are designed to re-enable ground commandability onboard the spacecraft. The response will attempt to command various telecom configurations and spacecraft attitudes in an attempt to find a viable uplink path. Each reconfiguration of a new uplink path is separated by an appropriate ground response interval for the SOFS team to re-acquire the spacecraft.

**Figure 7** illustrates Cassini's Command Loss Response chain. Once triggered, it progresses through a series of "Command Groups" divided by multi-hour "Command Pauses" which allow the SOFS team to react by sending an U/L command to halt the response. The Command Groups consist of actions to reconfigure redundant hardware and re-command spacecraft attitude and antennas. Each Command Pause allow several hours for the SOFS team to attempt re-acquisition of the spacecraft upon the newly commanded spacecraft configuration (the pause durations are set to a minimum of two RTLT periods). As shown in the figure, the first Command Group will select the Auxiliary Oscillator and execute the Safing Response which will turn off non-essential spacecraft loads, place the spacecraft in a lower power state, and re-direct the spacecraft's High Gain Antenna to sun-point, placing the spacecraft in a low U/L & D/L state through the LGA-1 antenna. After the first Command Group has executed, a 15 hour wait period (Command Pause) allows sufficient time for the SOFS Anomaly team to assemble at JPL and attempt re-establishment of the U/L, if possible, before RFS hardware swaps begin in successive Command Groups. If the re-acquisition attempt fails after Command Group #1 execution, the response will proceed with the next

**Figure 7.** Cmdloss response actions.

**7. Command loss FP**

238 Space Flight

**Figure 6.** Cassini's RFS functional diagram.

provided (covering up to 3 faults).

An unresolvable LOS condition where the ground is no longer able to deliver commands to the spacecraft will eventually lead to the activation of a LOS FP response. The actions of this response can help to re-establish the U/L. In Cassini's FP design, loss of D/L fault coverage is not protected in an "end-to-end" manner since the D/L is not considered to be a critical spacecraft function which requires autonomous restoration. But restoration of the U/L however, is considered crucial to mission success and is therefore allocated "endto-end" protection through a "Loss of Commandability" algorithm [10]. Although several other (higher priority) FP routines are installed into Cassini's FP suite to protect against these same type of failures in the U/L path (which provide more timely action), the Loss of Commandability algorithm provides a "safety net" type of FP which has the potential to restore both U/L and D/L. With this scheme in place, multiple levels of FP defense are

This catch-all type of FP is referred to as a "Command Loss FP" (from the perspective of the spacecraft since it is no longer receiving ground commands) and is typically an "endlessloop" response. The Command Loss Monitor aboard Cassini will detect an extended period course of actions specified in Command Group #2, which starts the series of RFS hardware unit swaps. Seven hour Command Pauses are installed between each subsequent Command Group to allow the SOFS team sufficient time to re-acquire the spacecraft on the newly commanded configuration. If the SOFS Anomaly team is able to re-acquire the vehicle within the first 71 hours (during the RFS unit swap phase), it is permissible for the HAS Response FP to execute (1 hour after the Command Loss Response has been terminated) via the selected (6NOP) U/L command which halts the response. Faults resolved during this first 71 hours are deemed to be "non-severe," since they are associated with RFS device failures. The HAS Response will increase the post-Safing U/L & D/L rates and swap from LGA-1 to the HGA antenna. However, if the Command Loss Response proceeds to Command Groups #5, it must be halted using the HAS FP "disable" command to keep the spacecraft on LGA-1 with the lower U/L & D/L rates, since the fault is considered to be too severe to transition to the higher rates. At the end of the Command Loss Response chain (approx. 7 days 15 hours), a swap to the redundant CDS is commanded and the Command Loss Response will start all over again on the redundant backup computer. The response will run endlessly until an U/L command is received by the ground. Once the spacecraft receives a ground command which restores the uplink successfully, the response will terminate and reset its Command Loss Timer, thus leaving the spacecraft on the last (successfully) commanded RFS/antenna configuration.

**9. EXCEL tool example: 2011 USO Failure**

is provided here for this USO failure event.

**1.** Time of LOS = > 17:15:00 UTC

**Figure 8.** SFOS file for USO failure event.

**2.** OWLT = > 1 hour 23 minute 51 seconds

**4.** Last time CLT was reset = > DOY357 @ 02:15:00 UTC

**Figure 8**:

**3.** Year = > 2011

Experience gained from the failed USO/LOS event on December 23, 2011 at BOT led to the development of this LOS/Cmdloss Timeline EXCEL Tool. To demonstrate its use, an example

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Once no signal was detected from Cassini on Day of Year (DOY) 357 of 2011, the ACE proceeded to follow the "LOS/Anomalous Downlink Contingency Plan" Procedure, performing sweeps of the spacecraft on different RFS configurations to attempt re-acquisition of the vehicle. A second DSN station was requested and confirmed no acquisition of Cassini's signal (ruling out weather and station configuration problems). Had the EXCEL tool been available at the time, the following data would have been collected from the SFOS file as noted in

#### **8. The LOS/Cmdloss timeline EXCEL tool**

In all cases, it is desirable to re-acquire the spacecraft before the Command Loss algorithm times out and triggers its response, if at all possible, since this FP routine will configure the LGA antenna, which yields extremely slow data delivery. Should this response trigger, the Command Group actions (device swaps, etc.) most likely cannot be confirmed in telemetry with the very slow D/L rate of 5 bits per second. Therefore, it was determined that two timelines were needed to provide visibility into fault possibilities and to supplement the LOS/Anomalous Downlink Contingency Plan Procedure recovery efforts: 1) a pre-Command Loss Response "LOS Timeline" containing FP expiration times (and the corresponding RFS configurations) to eliminate the numerous fault possibilities, 2) a timeline to track the Command Loss Response actions if activated. This goal was accomplished through the development of an EXCEL tool which receives minimal user inputs, utilizing the Space Flight Operations Schedule (SFOS) file which is used daily by both the ACE and SOFS teams. The "LOS/Cmdloss EXCEL Tool" provides the following:


## **9. EXCEL tool example: 2011 USO Failure**

Experience gained from the failed USO/LOS event on December 23, 2011 at BOT led to the development of this LOS/Cmdloss Timeline EXCEL Tool. To demonstrate its use, an example is provided here for this USO failure event.

Once no signal was detected from Cassini on Day of Year (DOY) 357 of 2011, the ACE proceeded to follow the "LOS/Anomalous Downlink Contingency Plan" Procedure, performing sweeps of the spacecraft on different RFS configurations to attempt re-acquisition of the vehicle. A second DSN station was requested and confirmed no acquisition of Cassini's signal (ruling out weather and station configuration problems). Had the EXCEL tool been available at the time, the following data would have been collected from the SFOS file as noted in **Figure 8**:


course of actions specified in Command Group #2, which starts the series of RFS hardware unit swaps. Seven hour Command Pauses are installed between each subsequent Command Group to allow the SOFS team sufficient time to re-acquire the spacecraft on the newly commanded configuration. If the SOFS Anomaly team is able to re-acquire the vehicle within the first 71 hours (during the RFS unit swap phase), it is permissible for the HAS Response FP to execute (1 hour after the Command Loss Response has been terminated) via the selected (6NOP) U/L command which halts the response. Faults resolved during this first 71 hours are deemed to be "non-severe," since they are associated with RFS device failures. The HAS Response will increase the post-Safing U/L & D/L rates and swap from LGA-1 to the HGA antenna. However, if the Command Loss Response proceeds to Command Groups #5, it must be halted using the HAS FP "disable" command to keep the spacecraft on LGA-1 with the lower U/L & D/L rates, since the fault is considered to be too severe to transition to the higher rates. At the end of the Command Loss Response chain (approx. 7 days 15 hours), a swap to the redundant CDS is commanded and the Command Loss Response will start all over again on the redundant backup computer. The response will run endlessly until an U/L command is received by the ground. Once the spacecraft receives a ground command which restores the uplink successfully, the response will terminate and reset its Command Loss Timer, thus leaving the spacecraft on the last (successfully) commanded RFS/antenna

In all cases, it is desirable to re-acquire the spacecraft before the Command Loss algorithm times out and triggers its response, if at all possible, since this FP routine will configure the LGA antenna, which yields extremely slow data delivery. Should this response trigger, the Command Group actions (device swaps, etc.) most likely cannot be confirmed in telemetry with the very slow D/L rate of 5 bits per second. Therefore, it was determined that two timelines were needed to provide visibility into fault possibilities and to supplement the LOS/Anomalous Downlink Contingency Plan Procedure recovery efforts: 1) a pre-Command Loss Response "LOS Timeline" containing FP expiration times (and the corresponding RFS configurations) to eliminate the numerous fault possibilities, 2) a timeline to track the Command Loss Response actions if activated. This goal was accomplished through the development of an EXCEL tool which receives minimal user inputs, utilizing the Space Flight Operations Schedule (SFOS) file which is used daily by both the ACE

• Sheet #1: instructions for using the EXCEL Tool & required inputs taken from the SFOS file • Sheet #2: Timeline #1 starting from LOS occurrence = > CLT = 0 seconds (Command Loss Re-

• Sheet #3: Timeline #2 detailing the Command Loss Response actions from CLT = 0 seconds

• Sheet #4: all corresponding end conditions for each FP response activation in Timeline #1 with

and SOFS teams. The "LOS/Cmdloss EXCEL Tool" provides the following:

configuration.

240 Space Flight

sponse trigger time)

through one entire CDS response cycle

the required recovery actions

**8. The LOS/Cmdloss timeline EXCEL tool**

**4.** Last time CLT was reset = > DOY357 @ 02:15:00 UTC


**Figure 8.** SFOS file for USO failure event.

#### **5.** Command Loss Defaults = > 115 hours

#### **6.** Prime RFS Devices set to: DST-A, TCU-B, TWTA-B

Once these data had been collected per the instructions listed in Sheet #1, EXCEL Sheet #2 inputs would be entered in the YELLOW spaces as shown in **Figure 9**, which in turn, will cause Sheet #2 through Sheet #4 to be populated with desired timing/post-fault configuration data. Copies of the SFOS and Sheet #2 though Sheet #4 would then be printed and distributed to each subsystem once the Anomaly team gathered to determine the cause and resolution of the LOS condition. As the group followed along with the SFOS file in LOS Timeline #1, spacecraft recovery efforts would have been coordinated with the Cassini ACE via telecom. All system-level FP responses are included in the LOS timeline for completeness (RFS-related responses are shown in red). These are the LATEST times that the FP responses would conclude, assuming that each activation started at BOT. Fault cases would be eliminated by the SOFS Anomaly team once re-acquisition for each completed FP response failed to re-establish the earth-spacecraft link.

1 Response acons contain unitswap(s)

<sup>2</sup> Exact response me is variable: SSPS FP Filter contains 3 cycles (192 switches \*3); this trip occurrence can occur any me within the last 192sec cycle (i.e. +/-3.2min)

3 RFS POR

**Note 2:** Failure to acquire S/C aer OWLT has elapsed could denote a problem with the 1-way oscillator (Aux Osc) **Note 3: RED**-LOS related faults; **BLACK** non-LOS faults *Not to Scale*

In the figure, each completed response notes whether a RFS POR occurs, as well as RFS device swap occurrences. The end of the timeline calculates when the Command Loss Timer will decrement to "0" seconds. For each response case, the resulting antenna selected (LGA or HGA if the HAS response is executed for that particular response) is noted in the timeline. Corresponding

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Although there are eight possible RFS combinations (see **Table 3**), there are only three DST/ TCU/TWTA combinations of interest due to the selection of RFS prime units in the FP (i.e. the FP will never command the alternate combinations). Also, telemetry delivery on the post-Safing commanded LGA is minimal at best, so that the recommendation to the Cassini ACE would be to attempt re-acquisition with the FP commanded RFS combinations after the HAS response had concluded (since all RFS-related responses will execute the HAS response to swap to the HGA antenna and increase the D/L rate). According to the LOS timeline, no new RFS configurations will be commanded after 3 hours 20 minute (so that the nominal DST-A/TCU-B/TWTA-B arrangement is assumed), since all RFS-related FP responses will have executed. Problems to focus on from this point forward would be an onboard sequencing error, an activation of the AACS FP, undetected RFS failures not protected by FP, a LGA-1 or HGA antenna failure, WTS-B failure, multiple faults, or possibly a waived failure; all which will most likely leave the spacecraft on the LGA-1 antenna (note: for a USO failure, the DST's

RFS post-response states and end conditions of interest are listed in Sheet #4 (**Table 2**).

VCO will take over the D/L delivery once 2-way communication is established).

**Table 2.** Post-response concluding end conditions (sheet #4).

**Figure 9.** LOS timeline of SFP response expiration times.

**Note 1:** All mes in ERT (UTC)

In the figure, each completed response notes whether a RFS POR occurs, as well as RFS device swap occurrences. The end of the timeline calculates when the Command Loss Timer will decrement to "0" seconds. For each response case, the resulting antenna selected (LGA or HGA if the HAS response is executed for that particular response) is noted in the timeline. Corresponding RFS post-response states and end conditions of interest are listed in Sheet #4 (**Table 2**).

Although there are eight possible RFS combinations (see **Table 3**), there are only three DST/ TCU/TWTA combinations of interest due to the selection of RFS prime units in the FP (i.e. the FP will never command the alternate combinations). Also, telemetry delivery on the post-Safing commanded LGA is minimal at best, so that the recommendation to the Cassini ACE would be to attempt re-acquisition with the FP commanded RFS combinations after the HAS response had concluded (since all RFS-related responses will execute the HAS response to swap to the HGA antenna and increase the D/L rate). According to the LOS timeline, no new RFS configurations will be commanded after 3 hours 20 minute (so that the nominal DST-A/TCU-B/TWTA-B arrangement is assumed), since all RFS-related FP responses will have executed. Problems to focus on from this point forward would be an onboard sequencing error, an activation of the AACS FP, undetected RFS failures not protected by FP, a LGA-1 or HGA antenna failure, WTS-B failure, multiple faults, or possibly a waived failure; all which will most likely leave the spacecraft on the LGA-1 antenna (note: for a USO failure, the DST's VCO will take over the D/L delivery once 2-way communication is established).


**5.** Command Loss Defaults = > 115 hours

the earth-spacecraft link.

**OWLT: Hr Min Sec CLT Default = 115** (in hours) **1 23 51 Last Reset @ 2:15:00** (HH:MM:SS)

**INPUT DATA:** Fill in Yellow slots only (ERT)

**BOT = 17:15:00** (HH:MM:SS) PST

**DST- A TCU- B TWTA- B DOY = 357**

242 Space Flight

1

3 RFS POR

Response acons contain unitswap(s)

**Note 1:** All mes in ERT (UTC)

**6.** Prime RFS Devices set to: DST-A, TCU-B, TWTA-B

**Year = 2011 RTLT**

**DST SSPS Trip (HGA) 2**

**& TWTA SSPS Trip (HGA)**

**Alert Msg 1/Safing (LGA) TWTA Fail (LGA) 1,3 Alert Msg 1/Safing (HGA)**

OWLT ~10 min ~38 min ~10 min ~50 min

**2**

**DST Fail (LGA) 1,3 19:10:17**

**Deep UV (LGA) 1,3**

**18:48:27**

<sup>2</sup> Exact response me is variable: SSPS FP Filter contains 3 cycles (192 switches \*3); this trip occurrence can occur any me within the last 192sec cycle (i.e. +/-3.2min)

**Note 2:** Failure to acquire S/C aer OWLT has elapsed could denote a problem with the 1-way oscillator (Aux Osc)

**Figure 9.** LOS timeline of SFP response expiration times.

**TCU SSPS Trip or Fail (LGA) 1,3 Shallow UV (LGA)**

**18:39:37 18:49:05**

**OP-2 (LGA) 18:40:39**

**18:39:30**

**Alert Msg 2/CDS Loss (LGA)**

**18:39:41**

Once these data had been collected per the instructions listed in Sheet #1, EXCEL Sheet #2 inputs would be entered in the YELLOW spaces as shown in **Figure 9**, which in turn, will cause Sheet #2 through Sheet #4 to be populated with desired timing/post-fault configuration data. Copies of the SFOS and Sheet #2 though Sheet #4 would then be printed and distributed to each subsystem once the Anomaly team gathered to determine the cause and resolution of the LOS condition. As the group followed along with the SFOS file in LOS Timeline #1, spacecraft recovery efforts would have been coordinated with the Cassini ACE via telecom. All system-level FP responses are included in the LOS timeline for completeness (RFS-related responses are shown in red). These are the LATEST times that the FP responses would conclude, assuming that each activation started at BOT. Fault cases would be eliminated by the SOFS Anomaly team once re-acquisition for each completed FP response failed to re-establish

**20:02:42**

**USO Failure! 20:02:42**

**OP-2 (HGA) 19:40:51**

**DST Fail (HGA) 1,3**

**20:26:14 TWTA Fail (HGA) 1,3**

**20:10:29**

. . . . . .

**Longest AHBL (LGA) <sup>1</sup> 22:10:11**

. . . . . .

**361 T 21:15:00**

**CmdLoss = 0 @**

 << HGA Swap Response >> @ OWLT + 1hr.

**19:26:02 19:39:42**

**TCU SSPS Trip or Fail (HGA) 1,3**

**19:39:53**

**19:39:49**

**Shallow UV (HGA)**

*1 hr 2 hr* **Elapsed Time Since BOT (hrs.)** *3 hr 5 hr*

**Note 3: RED**-LOS related faults; **BLACK** non-LOS faults *Not to Scale*

**Table 2.** Post-response concluding end conditions (sheet #4).


Time Coordinated), which is consistent with the SFOS file timeline (in successive pages to DOY357 which are not shown in this article). The timeline is also quoted in terms of DOY and elapsed time since the Command Loss Response triggered, showing each upcoming Command Group execution time. As mentioned before, the Command Groups consist of actions which reconfigure redundant hardware, eventually commanding spacecraft attitude and antennas in later Command Groups. Once a ground command is successfully received by the spacecraft, the response will be terminated, the CLT reset (to 115 hours), leaving the vehicle on the successfully

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The Command Loss Timeline is listed for one "CDS cycle" of the response. If all attempts to re-acquire the spacecraft have failed on the first response cycle of Command Groups on the prime CDS unit, the backup CDS computer will take over at the end of this response chain (after 7 days 15 hours 29 minute), so that the cycle is repeated on the redundant computer. As stated above, the Command Loss Response is an endless loop algorithm; below are the actions of the

• 1st Response Cycle: The Prime CDS uses its RAM load; it is then re-booted with a FSW load

• 2nd Response Cycle: The BU CDS takes over immediately using its RAM load; it is re-booted

stored on the SSR (at the end of the response cycle).

**Figure 11.** LOS/Cmdloss response info for SOFS team & ACE.

with a FSW load stored on the SSR (at the end of the response cycle).

commanded configuration.

response cycles:

**Table 3.** Possible RFS combinations.

#### **9.1. Command loss response activation**

If the SOFS Anomaly team was unable to re-acquire the spacecraft before the Command Loss Timer decremented to "0" seconds, the Command Loss Timeline in **Figure 10** would have been followed in synchrony with the SFOS file. In Sheet #3, the event times are listed in UTC (Universal

**Figure 10.** One command loss response cycle (sheet #3).

Time Coordinated), which is consistent with the SFOS file timeline (in successive pages to DOY357 which are not shown in this article). The timeline is also quoted in terms of DOY and elapsed time since the Command Loss Response triggered, showing each upcoming Command Group execution time. As mentioned before, the Command Groups consist of actions which reconfigure redundant hardware, eventually commanding spacecraft attitude and antennas in later Command Groups. Once a ground command is successfully received by the spacecraft, the response will be terminated, the CLT reset (to 115 hours), leaving the vehicle on the successfully commanded configuration.

The Command Loss Timeline is listed for one "CDS cycle" of the response. If all attempts to re-acquire the spacecraft have failed on the first response cycle of Command Groups on the prime CDS unit, the backup CDS computer will take over at the end of this response chain (after 7 days 15 hours 29 minute), so that the cycle is repeated on the redundant computer. As stated above, the Command Loss Response is an endless loop algorithm; below are the actions of the response cycles:


**Figure 11.** LOS/Cmdloss response info for SOFS team & ACE.

**9.1. Command loss response activation**

**Table 3.** Possible RFS combinations.

**Time (UTC) HH:MM:SS => DOY =>**

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*T3*

**1 106.27**

**Cmd Grp 4: Swap TWTAs**

**with 6NOP cmd**

**361**

**Cmd Grp 3: Swap TCUs CDS Safing**

**Cmd Grp 1: Terminate Sequence Reset CDU Select Aux Osc. CDS Safing**

**21:15:00 12:15:03**

**Recover S/C with HGA Disable cmd** 

**Figure 10.** One command loss response cycle (sheet #3).

**Response with:**

If the SOFS Anomaly team was unable to re-acquire the spacecraft before the Command Loss Timer decremented to "0" seconds, the Command Loss Timeline in **Figure 10** would have been followed in synchrony with the SFOS file. In Sheet #3, the event times are listed in UTC (Universal

**CLT=0 seconds =>** *T1 T3 T3 T3 T3 T3 T3*

**0.00 43.01**

**Cmd Loss Pauses: T1 = 15 hrs Prime RFS Unit Commanded: BLUE: Prime DST**

**19:15:03 2:15:11 9:15:11 16:15:19 23:15:19 363 36.01**

> **Cmd Grp 7: Configure alt ant & point to Earth**

**Cmd Grp 3: Swap TCUs CDS Safing**

**T4 = 7 hrs**

**A B B B B B B A B B A A B B A A B A A A A A A B**

*T3 T3 T2 T3 T4 T3* **A B A A A A A A B A B B A B B A B B A B B A B B**

**7:30:49 0:30:49 0:28:41 17:28:38 10:28:35 3:28:35 20:15:35 13:15:27**

*T3 T3 T3 T3 T3 T3 T4* **A B B A A B B A B B B B B B A B A A B A A B A A**

**2**

**365 92.23**

**Cmd Grp 3: Swap TCUs CDS Safing**

> **2 141.27**

**Cmd Grp 4: Swap TWTAs**

*(7days 15hr 29min 27sec)*

**Terminate** *T3 T2* 

**127.27**

**Re Response Acons cover S/C** 

**1 99.23**

**Cmd Grp 2: Swap DSTs (&CDUs)**

**362 22.00**

**Cmd Grp 3: Swap TCUs CDS Safing**

**Cmd Grp 3: Swap TCUs CDS Safing**

**113.27 134.27**

**14:30:57 21:30:57 4:31:05 11:31:05**

*2 min* 

**1 120.27 1 2**

**1 99.26**

**Cmd Grp 3: Swap TCUs CDS Safing**

**15.00 362**

**Cmd Grp 4: Swap TWTAs**

**Cmd Grp 2: Swap DSTs (&CDUs)**

**Elapsed Time (hr) =>** *T3*

**Cmd Grp 8: LGA Issue 7SAFE (Nom) power off USO**

**Cmd Grp 4: Swap TWTAs**

> **363 29.00**

**T2 = 7 hrs RED: Prime TCU T3 = 7 hrs GREEN: Prime TWTA**

*T3*

**3 162.49**

**Cmd Grp 7: Configure Alt ant & point to Earth**

**64.01**

**Cmd Grp 6: Power on B/U CDS**

**364**

**3 169.49**

**6:15:27**

**Cmd Grp 4: Swap TWTAs**

**364**

**57.01**

**Cmd Grp 3: Swap TCUs CDS Safing**

**B A A B A A B A A**

**3 155.49**

**Cmd Grp 8: LGA Issue 7SAFE (Nom) power off USO**

**364 71.23**

**Cmd Grp 5: Point -Z => sun (Constant roll)**

**363 50.01**

**Cmd Grp 3: Swap TCUs CDS Safing**

**Cmd Grp 5: Point -Z => sun (Constant roll)**

**12:44:27 5:44:24 22:44:21**

**4 176.49**

**18:31:13 1:31:13 8:31:21 15:44:21**

**3 148.27**

**Cmd Grp 3: Swap TCUs CDS Safing**

**363**

**Cmd Grp 6: Power on B/U CDS**

**Cmd Grp 2: Swap DSTs (&CDUs)**

**4 183.49**

**Swap CDS Units and Repeat**

**85.23 78.23**

**365 365**

• 3rd Response Cycle: The Prime CDS uses the default SSR FSW load from the previous reset; the Command Loss Timer is set to the FSW default value of CLT = 5 days; at the end of this cycle, the CDS is re-booted with the same FSW load stored on the SSR (at the end of the response cycle), but must wait 5 days before continuing the response.

be permanently lost (since there is no WTS FP on Cassini). In this case, the Command Loss Timer default of 115 hours would cause the Orbital Trim Maneuver (OTM) #407 to be missed on DOY078 should WTS-A fail (as well as the planned OTM backup opportunity on DOY079). To protect against loss of U/L after the WTS switch is commanded, the EXCEL tool was used to predict actions from the Command Loss Response which can provide a different U/L path through DST-B/WTS-B should WTS-A fail. The strategy shown in the figure depicts a reduced Command Loss Timer default of CLT = 72 hours with a "command moratorium" period implemented (no commanding allowed), which allows a controlled decrementation of the CLT timer during the RSS LGA Gravity Experiment. Once the test is complete on DOY075, an attempt to verify the telecom state by uplinking the original CLT default value of 115 hours is performed on DOY076. Should this U/L command fail to execute on the spacecraft, the command moratorium will continue until the CLT clocks down to "0" seconds, allowing the Command Loss Response to execute through to Command Group #2 which swaps DST-A= > DST-B, placing the U/L and D/L on WTS-B, just before the DSN track starts. The spacecraft would then be acquired on this new RFS configuration. The OTM would then proceed on the backup DSN pass. For Cassini, a failure of WTS-A would have meant that WTS-B must be used for the remainder of the mission, since the WTS-A switch is henceforth unusable. The actual execution of the

Cassini Spacecraft-DSN Communications, Handling Anomalous Link Conditions, and Complete…

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247

RSS Gravity Experiment was successful without the need for FP intervention.

Overall, anomalous D/L and LOS occurrences are very challenging and can be difficult for the SOFS Anomaly team to diagnose and resolve. Once the spacecraft's D/L signal is lost, an expedient, accurate resolution process is needed for quick re-acquisition of the vehicle. Identification of FP responses, their conclusion times and corresponding end states, as well as plausible LOS causes, is extremely helpful in eliminating fault cases systematically, thus allowing the SOFS Anomaly team to focus on the actual cause of the LOS problem. Unfortunately, pre-launch FP analyses do not always protect against all LOS-related fault possibilities since design oversites, lack of schedule or funding in implementing FP algorithms, errors within the FSW, or even false assumptions made during the pre-launch testing phase (waived failures) can occur. In all cases, it is highly desirable to address a LOS condition before the Command Loss FP response activates. But if not, a concise timeline of this response and its actions is essential in order to coordinate team efforts in attempting to re-acquire the vehicle; especially since the LGA-1 antenna is commanded, configuring the very low D/L rate which must be delivered through Cassini's very noisy Auxiliary Oscillator (backup device used since the primary USO failed). Therefore, the "LOS/Anomalous Downlink Contingency Plan" Procedure in combination with "LOS/Cmdloss" EXCEL tool is expected to be very useful when supporting this challenging class of faults during the remainder of Cassini's highly successful 20 year mission, until its final plunge into Saturn's

**11. Conclusions & lessons learned**

atmosphere on September 15, 2017.


For the 2011 USO failure event, the EXCEL LOS/Cmdloss Tool would have been used to generate the supporting data needed for trouble-shooting the anomaly for the SFOS Anomaly team, with recommendations included for the Cassini ACE as shown in **Figure 11**.

#### **10. Other uses for the Excel tool**

Cassini also relies upon the Command Loss Response to protect events of significant importance should a loss of U/L occur during science experiments and other selected spacecraft activities. **Figure 12** provides an example of this type of "Command Loss Response strategy" used to support the RSS LGA Gravity Experiment performed in 2015, where the HGA must be swapped to LGA-1 and then back again to HGA. The risk associated with this experiment was commanding the WTS switch during the HGA/LGA-1/HGA antenna swap series, where if a malfunction occurred on WTS-A, the U/L capability would

**Figure 12.** EXCEL tool support of 2015 RSS LGA gravity experiment.

be permanently lost (since there is no WTS FP on Cassini). In this case, the Command Loss Timer default of 115 hours would cause the Orbital Trim Maneuver (OTM) #407 to be missed on DOY078 should WTS-A fail (as well as the planned OTM backup opportunity on DOY079). To protect against loss of U/L after the WTS switch is commanded, the EXCEL tool was used to predict actions from the Command Loss Response which can provide a different U/L path through DST-B/WTS-B should WTS-A fail. The strategy shown in the figure depicts a reduced Command Loss Timer default of CLT = 72 hours with a "command moratorium" period implemented (no commanding allowed), which allows a controlled decrementation of the CLT timer during the RSS LGA Gravity Experiment. Once the test is complete on DOY075, an attempt to verify the telecom state by uplinking the original CLT default value of 115 hours is performed on DOY076. Should this U/L command fail to execute on the spacecraft, the command moratorium will continue until the CLT clocks down to "0" seconds, allowing the Command Loss Response to execute through to Command Group #2 which swaps DST-A= > DST-B, placing the U/L and D/L on WTS-B, just before the DSN track starts. The spacecraft would then be acquired on this new RFS configuration. The OTM would then proceed on the backup DSN pass. For Cassini, a failure of WTS-A would have meant that WTS-B must be used for the remainder of the mission, since the WTS-A switch is henceforth unusable. The actual execution of the RSS Gravity Experiment was successful without the need for FP intervention.

#### **11. Conclusions & lessons learned**

• 3rd Response Cycle: The Prime CDS uses the default SSR FSW load from the previous reset; the Command Loss Timer is set to the FSW default value of CLT = 5 days; at the end of this cycle, the CDS is re-booted with the same FSW load stored on the SSR (at the end of the re-

• 4th Response Cycle: The BU CDS uses the default SSR FSW load from the previous reset; the Command Loss Timer is set to the FSW default value of CLT = 5 days; at the end of this cycle, the CDS is re-booted with the same FSW load stored on the SSR (at the end of the response

For the 2011 USO failure event, the EXCEL LOS/Cmdloss Tool would have been used to generate the supporting data needed for trouble-shooting the anomaly for the SFOS Anomaly

Cassini also relies upon the Command Loss Response to protect events of significant importance should a loss of U/L occur during science experiments and other selected spacecraft activities. **Figure 12** provides an example of this type of "Command Loss Response strategy" used to support the RSS LGA Gravity Experiment performed in 2015, where the HGA must be swapped to LGA-1 and then back again to HGA. The risk associated with this experiment was commanding the WTS switch during the HGA/LGA-1/HGA antenna swap series, where if a malfunction occurred on WTS-A, the U/L capability would

team, with recommendations included for the Cassini ACE as shown in **Figure 11**.

sponse cycle), but must wait 5 days before continuing the response.

cycle), but must wait 5 days before continuing the response. • 5th Response Cycle - ∞: Repeat cycles 3 & 4 above indefinitely.

**10. Other uses for the Excel tool**

246 Space Flight

**Figure 12.** EXCEL tool support of 2015 RSS LGA gravity experiment.

Overall, anomalous D/L and LOS occurrences are very challenging and can be difficult for the SOFS Anomaly team to diagnose and resolve. Once the spacecraft's D/L signal is lost, an expedient, accurate resolution process is needed for quick re-acquisition of the vehicle. Identification of FP responses, their conclusion times and corresponding end states, as well as plausible LOS causes, is extremely helpful in eliminating fault cases systematically, thus allowing the SOFS Anomaly team to focus on the actual cause of the LOS problem. Unfortunately, pre-launch FP analyses do not always protect against all LOS-related fault possibilities since design oversites, lack of schedule or funding in implementing FP algorithms, errors within the FSW, or even false assumptions made during the pre-launch testing phase (waived failures) can occur. In all cases, it is highly desirable to address a LOS condition before the Command Loss FP response activates. But if not, a concise timeline of this response and its actions is essential in order to coordinate team efforts in attempting to re-acquire the vehicle; especially since the LGA-1 antenna is commanded, configuring the very low D/L rate which must be delivered through Cassini's very noisy Auxiliary Oscillator (backup device used since the primary USO failed). Therefore, the "LOS/Anomalous Downlink Contingency Plan" Procedure in combination with "LOS/Cmdloss" EXCEL tool is expected to be very useful when supporting this challenging class of faults during the remainder of Cassini's highly successful 20 year mission, until its final plunge into Saturn's atmosphere on September 15, 2017.

#### **Acknowledgements**

This research was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under a contract with the National Aeronautics and Space Administration.

RCS Reaction control system

RFS Radio frequency system

RSS Radio science subsystem

SOFS Spacecraft operations flight team

RTLT Round trip light time

SOI Saturn orbit insertion

SSPS Solid state power switch

SPF Single point failure

SSR Solid-state recorder

USO Ultra-stable oscillator

UTC Universal time coordinated

WTS Waveguide transfer switch

UVIS Ultraviolet imaging spectrograph

VIMS Visible & infrared mapping spectrometer

Address all correspondence to: paula.s.morgan@jpl.nasa.gov

Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California,

[1] Morgan P. Cassini. Mission-to-saturn spacecraft overview & cds preparations for endof-mission proximal orbits. Jet Propulsion Laboratory/California Institute of Technology.

In: Proceedings of the IEEE/AIAA Conference. Montana: Big Sky; March 2015

U/L Uplink

**Author details**

Paula S. Morgan

**References**

United States of America

RFIS Radio frequency instrument subsystem

Cassini Spacecraft-DSN Communications, Handling Anomalous Link Conditions, and Complete…

http://dx.doi.org/10.5772/intechopen.72075

249

RPWS Radio & plasma wave science instrument

RF Radio frequency

#### **Nomenclature**


AACS Attitude, articulation, & control system


**Acknowledgements**

248 Space Flight

**Nomenclature**

D/L Downlink DOY Day of year

This research was carried out at the Jet Propulsion Laboratory, California Institute of Technology,

under a contract with the National Aeronautics and Space Administration.

AACS Attitude, articulation, & control system

CDS Command & data processing system

HAS High gain antenna swap (algorithm)

INMS Ion & neutral mass spectrometer

MAG Dual technique magnetometer

OTM Orbital trim maneuver

OWLT One-way light time

MIMI Magnetospheric imaging instrument

JPL Jet propulsion laboratory

CIRS Composite infrared spectrometer

BOT Beginning of (DSN) track

CDA Cosmic dust analyzer

DSN Deep space network

FP Fault protection FSW Flight software

HGA High gain antenna

LGA Low gain antenna

LOS Los of signal

ME Main engine

ESA European space agency

ASI Agenzia Spaziale Italiana (Italian space agency)


#### **Author details**

Paula S. Morgan

Address all correspondence to: paula.s.morgan@jpl.nasa.gov

Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California, United States of America

#### **References**

[1] Morgan P. Cassini. Mission-to-saturn spacecraft overview & cds preparations for endof-mission proximal orbits. Jet Propulsion Laboratory/California Institute of Technology. In: Proceedings of the IEEE/AIAA Conference. Montana: Big Sky; March 2015


[2] Jet Propulsion Laboratory/California Institute of Technology Saturn Tour Highlights. In: Cassini-Huygens Website [Internet]. 2016. Available from: https://saturn.jpl.nasa.gov/

[3] National Aeronautics and Space Administration's Office of Planetary Protection [Internet].

[4] Doody D. Deep Space Craft: An Overview of Interplanetary Flight. Chichester: Praxis; 2009 [5] Jones C. Cassini project pre-ship review/single point failures. Jet Propulsion Laboratory/

[6] Morgan P. Cassini spacecraft's in-flight fault protection redesign for unexpected regulator malfunction. Jet Propulsion Laboratory/California Institute of Technology. In:

[7] Jet Propulsion Laboratory/California Institute of Technology Cassini Significant Events, Cassini-Huygens News & Features. 2011. Available from: https://saturn.jpl.nasa.gov/

[8] Morgan P. Resolving the difficulties encountered by JPL interplanetary robotic spacecraft in flight. In: Ghadawala R, editor. Advances in Spacecraft Systems and Orbit Determination.

[9] Taylor J, Sakamoto L, Wong C. Cassini Orbiter/Huygens Probe Telecommunications Deep Space Communications and Navigation Systems Center of Excellence (Descanso)

[10] Morgan P. Robotic spacecraft health management. In: Johnson S, Gormley T, Kessler S, Mott C, Patterson-Hine A, Reichard K, Scandura P, editors. System Health Management:

Proceedings of the IEEE/AIAA Conference. Montana: Big Sky; March 2010

news/2861/2016-saturn-tour-highlights/

250 Space Flight

California Institute of Technology. 1997

news/1935/cassini-significant-events-122111-1312/

Design and Performance Summary Series. 2002

1st ed. Croatia: InTech Open Access; 2012. p. 235-264. ch11

With Aerospace Applications. 1st ed. Wiley; 2011. p. 543-554. ch34

2014. Available from: http://planetaryprotection.nasa.gov/

## *Edited by George Dekoulis*

Space has always been intriguing people's imagination. However, space flight has only been feasible over the last 60 years. The collective effort of distinguished international researchers, within the field of space flight, has been incorporated into this book suitable to the broader audience. The book has been edited by Prof. George Dekoulis, Aerospace Engineering Institute (AEI), Cyprus, an expert on the state-of-the-art implementations of reconfigurable space physics systems. The book consists of six sections, namely, "Introduction," "Spacecraft Simulators," "Spacecraft Navigation," "Spacecraft Propulsion," "Suborbital Flight," and "Deep-Space Flight." We hope that this book will be beneficial for professionals, researchers, and academicians and inspires the younger generations into pursuing relevant academic studies and professional careers within the space industry.

Published in London, UK © 2018 IntechOpen © Zenobillis / iStock

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