Long-Term Venus Lander

Western Michigan University ScholarWorks at WMU

Honors Theses Lee Honors College

4-20-2021

Scott Miller Western Michigan University, [email protected]

Follow this and additional works at: https://scholarworks.wmich.edu/honors_theses

Part of the Space Vehicles Commons

Recommended Citation Miller, Scott, "Long-Term Venus Lander" (2021). Honors Theses. 3383. https://scholarworks.wmich.edu/honors_theses/3383

This Honors Thesis-Open Access is brought to you for free and open access by the Lee Honors College at ScholarWorks at WMU. It has been accepted for inclusion in Honors Theses by an authorized administrator of ScholarWorks at WMU. For more information, please contact [email protected].

Long-Duration Venus Explorer

Aidan Wales, Ethan Reid, and Scott Miller

Group #04-21-01

Faculty Advisor: Dr. Kristina Lemmer

Western Michigan University

College of Engineering and Applied Sciences

Table of Contents

Abstract ...... pg. 1 Disclaimer ...... pg. 1 Introduction ...... pg. 2 Methodology ...... pg. 2 Preliminary Research ...... pg. 2 Decision Matrices ...... pg. 4 Additional Research and Calculations ...... pg. 10 Discussion of Results ...... pg. 20 Masses ...... pg. 20 Power Draw ...... pg. 22 Concept of Operations ...... pg. 23 Impacts of Project on Engineering Solutions ...... pg. 24 Conclusions and Recommendations ...... pg. 24 Acknowledgements...... pg. 25 References ...... pg. 26 Appendix A: Thermoelectric Generator Calculations ...... pg. 32 Appendix B: Orbit Calculations ...... pg. 36 Appendix C: Concept of Operations ...... pg. 40 Appendix D: MATLAB Codes ...... pg. 49

Table of Figures

Figure 1: Graphic of General Operation of TEG ...... pg. 6 Figure 2: Heat Pipe System Diagram ...... pg. 9 Figure 3: Multilayer Insulation Example ...... pg. 10 Figure 4: Ammonia Internal Flow Through Throttle Valve (100 bar) .... pg. 11 Figure 5: Heat Transfer Insulation Thickness Trade Study ...... pg. 12 Figure 6: Ammonia Piping around Phase Change Thermal Storage Device...... pg. 12 Figure 7: Simscape Thermal Circuit Model ...... pg. 13 Figure 8: Temperature Differential Across Lander Shells vs. Time Following Ammonia Depletion ...... pg. 13 Figure 9: Simplified Solar Panel Rendering...... pg. 15 Figure 10: Simplified Orbital Relay Views...... pg. 16 Figure 11: Heat Shield and Aeroshell ...... pg. 17 Figure 12: Control Fin Stress Concentrations ...... pg. 17 Figure 13: Control Fin Overall Displacements ...... pg. 18

Table of Tables

Table 1: Lander Pressure Management Decision Matrix ...... pg. 4 Table 2: Lander Power Generation Decision Matrix ...... pg. 5 Table 3: Lander Heat Management Decision Matrix ...... pg. 6 Table 4: Lander Terrain Stability Decision Matrix ...... pg. 7 Table 5: Orbital Relay Power Generation Decision Matrix ...... pg. 8 Table 6: Orbital Relay Thermal Control System Decision Matrix ...... pg. 9 Table 7: Lander Scientific Instrument Manifest ...... pg. 14 Table 8: Orbital Parameters and ∆V Summary ...... pg. 19 Table 9: Orbital Relay Mass Table ...... pg. 20 Table 10: Orbital Relay Communications System Mass Table ...... pg. 20 Table 11: Lander, Cruise, and EDL Mass Table ...... pg. 21 Table 12: Final Launch Masses ...... pg. 21 Table 13: Orbital Relay Input Power Table ...... pg. 22 Table 14: Lander Power Draw Table ...... pg. 22

Abstract

Venus continues to be a target of interest for the scientific community, as the similarities and differences to Earth pose intriguing questions about planetary evolution. The National Aeronautics and Space Administration (NASA) is looking to gain a better understanding of Earth’s sister planet and has presented a challenge to interested undergraduate student teams to develop a in-situ robotic explorer and accompanying relay. The student team representing Western Michigan University has designed a long-term Venusian lander to both survive the harsh environment for an extended period and gather scientific data on the planet’s atmosphere and surface to be communicated back to Earth through the accompanying orbital relay. Subsystems for the lander and orbital relay such as power generation and temperature management were considered. Verification of the systems was obtained through hand calculations and computer software provided by Western Michigan University, such as MATLAB and Simscape. A Concept of Operations was created to reflect the general timeline of the mission and its various tasks being conducted. It was determined that until further technological advancements are made, the 90-day goal set forth by NASA cannot be met. With different approaches taken by the student teams, NASA can potentially use the designs from the competitions to inform the design of their own long-term Venusian lander. *Disclaimer: This project was written by students at Western Michigan University to fulfill an engineering curriculum requirement. Western Michigan University makes no representation that the material contained in this report is error-free or complete in all respects. Persons or organizations who choose to use this material do so at their own risk.

Introduction Venus has long been a subject of interest for the scientific community. While it is similar to Earth in mass, volume, and mean radius, it has a radically different atmospheric composition and amount of volcanic activity (Lopes & Gregg, 2004). Since climate change becomes a more urgent issue with each passing day, it is important to understand whether Earth could eventually evolve to be similarly hostile. Determining this requires taking scientific measurements on the surface of Venus. However, the high temperatures and pressures at the surface have precluded long-duration, in-situ scientific observations. The longest mission on the surface of Venus thus far was the Venera 13 mission, which lasted 127 minutes on the surface (NASA Space Science Data Coordinated Archive [NSSDCA], n.d.e). Thus, lander designs that can last for an extended period on the Venusian surface are needed, and the goal is a design capable of completing a 90-day mission. The mass of the lander must be limited to nor more than 1,500 kg, with at least 200 kg being dedicated to scientific instruments. The measurements taken must be sent to Earth via a relay in orbit around Venus. An overall mission must also be designed for the lander, including factors such as launch, entry/descent/landing (EDL) operations, and the mission timeline. The lander, named Endurance, is designed to satisfy the need for an extended-duration lander on Venus. This design employs new technologies to increase longevity on the surface, such as aerogel insulation, phase-change thermal storage, and a thermoelectric generator. In addition, Endurance features a scientific instrument suite to examine the composition of Venus’s atmosphere and surface, as well as the evolution of the planet.

Methodology The design process began with research on the Venusian environment and past missions to Venus, as well as Mars. This was to determine the nature of the operating environment, identify appropriate scientific instruments to include, and learn from past attempts to mitigate the effects of the high pressures and temperatures on Venus. Following this, methods for managing pressure, generating power, managing heat, and remaining stable on the surface were investigated. Decision matrices were used to make the final selections for critical design features. Further research and calculations were performed to refine these features, and an orbital relay and scientific instrument suite were developed. Finally, masses and power draws were determined, and the orbital trajectories and concept of operations were formulated.

Preliminary Research Before the mission can be designed, it is critical to fully understand the environment for which the mission will be deployed. The biggest challenge to account for is the high temperature experienced at the Venusian surface, averaging at approximately 464 ⁰C (Williams, 2005). This high temperature is a concern for electronics and scientific instruments not specifically designed for high-temperature operation; even with current technology, these high- temperature instruments can only handle such a high temperature for a limited time. The high pressure, on the order of 92 times that experienced at Earth’s surface, has the potential to cause structural failure that would jeopardize the interior of the lander (Williams, 2005). The large amount of volcanic activity on the surface means that the lander must be deployed away from active sites (Lopes & Gregg, 2004). Finally, the mix of chemicals that make up the atmosphere, such as sulfuric acid, is corrosive to most metals, indicating that degradation of the lander’s shell must be considered (Williams, 2005).

To determine what methods have previously been used to address the challenging environment on Venus and what scientific instruments should be included, prior missions to Venus were researched. Most of the atmospheric vehicles previously sent to the Venusian surface were part of the Venera program, with the exception being the Pioneer Venus probes (Krebs, 2021). Each Venera lander utilized a spherical pressure vessel as its main body, landed on a toroidal landing platform with a suspension, used a combination of a disk aerobrake and parachute to descend to the surface, and a helical antenna wrapped around a cylinder on top of the lander (NSSDCA, n.d.e). Prior to entry, the landers were precooled using circulating fluid to increase the amount of time they would operate in the atmosphere (NSSDCA, n.d.d). The Pioneer Venus probes all flew on the same mission as part of a multiprobe, which separated into three small probes, a large probe, and the spacecraft bus (Fimmel et al., 1983, p. 46). These probes each utilized a spherical pressure vessel, an aeroshell that protected the lander from heating during descent, a radio-transparent aft cover to allow the antenna to communicate during descent and relied on aerodynamic drag to slow down; the large probe also utilized a parachute (Fimmel et al., 1983, pp. 48–49, 53). The scientific instruments of the Venera missions varied; however, Venera landers 9-12 all included temperature, pressure, acceleration, and wind sensors (NSSDCA, n.d.c., n.d.d). Venera 9 and 10 included an imaging system, nephelometer, mass spectrometer, gamma-ray spectrometer, photometers, and a radiation densitometer (NSSDCA, n.d.c). Venera 11 and 12 contained a gas chromatograph, solar radiation sensors, soil composition sensors, a soil penetrator, and an atmospheric electrical discharge detector (NSSDCA, n.d.d). Venera 13 and 14 carried a camera system with filters, an X-ray fluorescence spectrometer, a drill and surface sampler, a dynamic penetrometer, a seismometer, and an acoustic detector (NSSDCA, n.d.e). It was found that a nephelometer performs analyses on clouds (NSSDCA, n.d.a), spectrometers and gas chromatographs help determine material and gas composition (NASA Atmospheric Experiments Laboratory, n.d.), and dynamic penetrometers allow mechanical properties of soil to be determined (NSSDCA, n.d.e). The Pioneer Venus mission included similar instruments. The bus contained neutron and ion mass spectrometers, and all probes contained temperature/pressure/acceleration sensors and a nephelometer (Fimmel et al., 1983, pp. 56, 79). The small probes carried net flux radiometers to aid in analysis of atmospheric circulation (Fimmel et al., 1983, p. 83). Finally, the large probe carried a gas chromatograph, a mass spectrometer, a solar flux radiometer to determine where solar radiation was deposited in the atmosphere, an infrared radiometer to observe heat variations with altitude, and a cloud particle size spectrometer (Fimmel et al., 1983, pp. 72, 76– 79). Past NASA missions to Mars were also examined to determine additional scientific instruments to include. The Curiosity rover features an alpha particle X-ray spectrometer and laser-camera combination to analyze surface composition, a gas chromatograph, and a variety of additional spectrometers (NASA Mars Exploration Program [NASA MEP], n.d.a, n.d.b, n.d.i). The Mars Odyssey orbiter carries a gamma ray spectrometer to examine elemental composition of the surface from space (NASA MEP, n.d.c). The Mars Exploration Rovers used Mössbauer spectrometers to identify iron-bearing materials on the surface (NASA MEP, n.d.f). Finally, the Perseverance rover features a subsurface radar, an X-ray spectrometer to examine elements and search for microbial life, a weather measurement suite, a camera system, and a Raman spectrometer to search for organic materials (NASA MEP, n.d.d, n.d.e, n.d.g, n.d.h, n.d.j).

Decision Matrices

Lander Pressure Management Potential systems for pressure management included a traditional lander body, an at- pressure lander body, syntactic foam, a spherical lander body, and multiple spherical metal shells. The at-pressure lander body is characterized by an internal lander pressure equal to that of the ambient atmospheric pressure. The syntactic foam is a material often employed by submersibles for pressure management underneath the oceans of Earth. Multiple spherical metal shells would be a way to lessen the pressure differential across each shell until the pressure differential would be manageable for the scientific instrument suite. The design matrix for pressure management is found in Table 1 below.

Table 1: Lander Pressure Management Decision Matrix

The two lowest scoring options were the at-pressure lander body and multiple spherical shells, due to the high system support that would be needed to maintain optimal conditions and high mass. The two highest scoring options were the spherical lander body and the syntactic foam, due to the flight heritage of both systems and the negligible power draw required to manage pressure. The spherical lander body evenly distributes the pressure from the Venusian atmosphere. When used as a sandwich core, the syntactic foam can withstand the high pressures involved (Hodge et al., 2000). An additional factor that syntactic foam exhibits is its low thermal conductivity at high temperatures (Hodge et al., 2000). After all factors were considered, it was decided that a combination of a spherical lander body and syntactic foam was the best option. The combination of the spherical body and syntactic foam will ensure the structural safety of the lander. Along with the pressure management system, a material must be selected for the lander body. The material must withstand both the high pressure and temperature on the surface and experience minimal degradation due to the corrosive atmospheric conditions. To accomplish this, the metal needs a low thermal conductivity, a high melting point, and a high compressive strength. Research into various metals suggests that the titanium alloy Ti-6Al-4V

can withstand the atmospheric composition of Venus with minimal degradation (Lukco et al., 2020). In addition, the alloy’s melting point is higher than the temperature experienced at the surface, the alloy has a low thermal conductivity, and the compressive strength is above the pressure experienced (Titanium Alloys Ti6Al4V Grade 5, 2002). Thus, Ti-6Al-4V was selected as the material of the spherical lander body.

Lander Power Generation To generate the electrical power required for scientific operations on the surface of Venus, several methods were considered. A design matrix quantifying the reasoning summarized here can be found in Table 2 below.

Table 2: Lander Power Generation Decision Matrix

Using batteries alone limits the possible mission time significantly, as the energy used cannot be counterbalanced by recharging. Solar panels are not viable due to surface light levels being comparable to an overcast day on Earth (Lumen Learning, n.d.) and the need to design the panels to withstand the high surface pressures. A radioisotope thermoelectric generator requires bringing a heat source to the surface of Venus (NASA Jet Propulsion Laboratory, n.d.a); since heat management is already a major mission challenge, this method was not selected. Thermogalvanic cells are designed to recover waste heat from a power generation process (Borghino, 2014); therefore, they will not generate a useful amount of energy alone. The remaining options considered are all forms of heat engines, which have the advantage of utilizing the existing temperature gradient between the lander’s exterior and interior to generate power. The Stirling engine and Rankine engine both require moving parts, which increases complexity (Afework, Hanania, Heffernan, Jenden, Stenhouse, & Donev, 2018; Afework, Hanania, Stenhouse, & Donev, 2018). However, the thermoelectric generator (also known as a Seeback generator or TEG) has no moving parts, has a relatively low mass, and is extremely reliable (Piggott, n.d.), so it was selected to serve as the power source in conjunction with the onboard batteries. Figure 1 depicts the general operation of the TEG.

Figure 1: Graphic of General Operation for TEG

Lander Thermal Regulation Several options were explored for heat management, including both active and passive solutions. Active systems included a heat pump, external radiation, fans, and using the interior of the lander as a heat engine cold sink. Passive systems include using insulation, using run-hot electronics, and using a phase change thermal storage; this is referred to as a thermal battery in Table 3 below.

Table 3: Lander Heat Management Decision Matrix

The heat pump did not score well because of the power required to operate at the temperature difference targeted. Passive insulation scored well due to the many suitable options available. Run-hot electronics scored well because of the simplicity of the solution and the lack of

additional power draw. External radiation did not score well because of the mass penalty and the requirement for the radiation device to be hotter than the Venusian atmosphere. Using the lander interior as a heat engine cold sink (specifically a Rankine engine) did not score well; upon further research, it was found that the heat engine would add thermal energy rather than removing it. Fans scored well because of their small mass and moderate heat distribution score. The phase change thermal storage device (thermal battery) scored well because of the mechanical simplicity of the system as well as the lack of power draw for the system. In the end, a combination of passive insulation, run-hot electronics, and a phase change thermal storage device was chosen. These were chosen over other similarly scoring options largely because of their low power draw requirements. This is important because power generation on the Venusian surface will be difficult, so it is desirable to reserve power for other systems. The passive insulation is a reinforced aerogel composite developed by the NASA Glenn Research Center (Aerogel Reinforced Composites, 2020). The insulation is sandwiched between the outer and inner titanium shells of the lander, both with a thickness of 10mm. The run-hot electronics are composed of silicon carbide (NASA SME, Personal Communication, February 2020). The phase change thermal storage device is a titanium cylinder with a rounded top, a diameter of 240 mm, and a total height of 815 mm. The phase change thermal storage device will allow for 0.035 m3 of phase change material. Several phase change materials (PCMs) were researched, primarily miscible gap alloys (Kisi et al., 2018); however, the specific heat capacities, phase change temperatures, densities, and volumes required did not align with the lander requirements. Water was chosen as the PCM for the phase change thermal storage device because of its high specific heat capacity and phase change temperature at the lander’s internal pressure.

Lander Terrain Stability To ensure the lander dampens its touchdown and remains stable on the surface, several systems were considered. A design matrix quantifying the reasoning summarized here is shown in Table 4 below.

Table 4: Lander Terrain Stability Decision Matrix

As previously mentioned, the Venera landers utilized a toroidal landing platform with a suspension system to absorb the shock of touchdown. Traditional landing legs (using either hydraulics, pneumatics, or a crush core) have been used on previous missions to other bodies as well. Several unconventional options were considered as well, such as a folding leg system mimicking that of a daddy long-legs spider, spherical airbags surrounding the lander body, a flat lower surface with high friction against the Venusian surface, and a body designed to withstand multiple impacts. The daddy long-legs system would enable a larger leg span and stable landings on unideal terrain. However, it also introduces complexity, and the leg thinness relative to the length increases the chance of a structural failure. Spherical airbags were not a practical choice; the high atmospheric pressure on Venus means that airbag inflation requires an even higher pressure. A flat lander bottom would prevent sliding and minimizes the number of parts on the lander, but it introduces a large surface for conductive heat transfer and cannot account for landings on rocky areas. A spherical body designed to withstand impact damage would ensure the lander’s survival upon landing, but it does not provide any stability (the lander would be free to roll down an incline). The toroidal platform and traditional landing legs were the two final options considered due to their simplicity, flight heritage, stabilizing capabilities, and impact damping capabilities. The traditional landing legs were selected as the final design feature due to their extensive prior use (both on space missions and on Earth), smaller total contact area for conductive heat transfer from the surface, and stability on inclines. Traditional legs have a larger span and allow the lander to lean onto one or more of the legs if it lands on a steep incline. The toroidal platform is a rigid unit with a smaller span; thus, it is more difficult to keep the center of gravity within the lander’s footprint.

Orbital Relay Power Generation Two systems were considered for the orbital relay power generation: a solar cell array and a radioisotope generator. The solar cell array harnesses the solar energy of the Sun to generate electricity, while the radioisotope generator harnesses the energy present from a decaying radioactive material. The design matrix for the orbital relay power generation is shown below in Table 5 below.

Table 5: Orbital Relay Power Generation Decision Matrix

Both systems scored similarly for different reasons. The solar panels have slightly better power generation and have more flight heritage, but the radioisotope generator has less mass involved. Based on these factors, the solar cell array was selected since the close proximity to the Sun results in excellent power generation.

Orbital Relay Thermal Control A variety of systems were researched, including multilayer insulation, special coatings, radioisotope heater units, patch heaters, cartridge heaters, louvers, heat pipes, and a thermoelectric cooler. To provide redundancy, a passive and active thermal control system will be selected. The design matrix for the orbital relay thermal control is shown in Table 6 below, where the first three systems are classified as passive systems and the rest are classified as active systems.

Table 6: Orbital Relay Thermal Control System Decision Matrix

The factors considered for the system include heritage, power draw, and mass. All passive systems scored well, but the multilayer insulation fared better because of its lower mass. As for the active systems, the two best contenders were the heat pipe system and the patch heater. Preference was given to the heat pipe system, as there is no power draw associated with the system. Based on these factors, a combination of multilayer insulation and a heat pipe were selected. Both choices have no power draw, which will allow for less power that needs to be generated. Figure 2 and Figure 3 showcase the heat pipe system and an example of multilayer insulation, respectively.

Figure 2: Heat Pipe System Diagram

Figure 3: Multilayer Insulation Example

Additional Research and Calculations

Lander Power Generation By assuming values of the lander shell thickness and internal lander temperature, preliminary calculations for the heat transferred into the lander were completed. The convection heat transfer coefficient outside of the lander was estimated using methods given by Bengtson (2010) for natural convection from a sphere. An upper bound for the thermal efficiency was found using the Carnot efficiency for the temperature gradient between the inside and outside of the lander. Once the shell thicknesses and internal lander temperature were determined, a MATLAB code was used to calculate the masses of the layers, the TEG surface area, and the TEG mass. The MATLAB code is presented in Appendix D. These calculations assume a hypothetical thermoelectric material with a density equal to 54% that of lead telluride, which is a thermoelectric material currently being used (Quick, 2012). In addition, 25 W of power are assumed to be supplied by the TEG, and the hypothetical thermoelectric material has a figure of merit equal to 20, where the figure of merit is defined as by G. Snyder & A. Snyder (2017). Without technology advancements in battery energy density, refrigeration technology, and thermoelectric material density, it is not possible to generate large amounts of power with a thermoelectric generator without incurring significant mass penalties. A low-powered (roughly 100 W) active refrigeration system that would maintain a steady temperature differential indefinitely was considered; however, the coefficient of performance of the required system was significantly higher than the maximum possible value, making such a system impossible. Thus, the TEG does not generate a substantial amount of the lander’s power; rather, its purpose is to keep the batteries charged so that they are at full capacity once thermal regulation capability expires. The calculations conducted in this analysis can be found in Appendix A.

Lander Thermal Regulation A form of heat removal is necessary to survive for an extended period on the Venusian surface. A system including a high-pressure storage tank of ammonia and an expansion valve was designed to meet this need. The ammonia is stored as a liquid at 800 bar, and when the internal temperature of the lander reaches 300 ⁰C, the ammonia passes through an expansion valve to become a gas at 100 bar and 189 ⁰C. The ammonia is piped in a helical fashion around the phase change thermal storage device and expelled from the lander. The ammonia expulsion is released in short bursts to allow the expanded ammonia to reach thermal equilibrium. When new ammonia is required to keep equilibrium inside the lander, it is released from the storage tank and expanded while the previously expanded ammonia is released through an exterior valve. It is estimated that 39.19 kg of ammonia will expel 35 MJ of thermal energy. By maintaining the internal temperature of the lander at 300 ⁰C, it is estimated that 39.19 kg of ammonia yields an extra 10 hours of thermal equilibrium. This estimation is based on the insulation trade study of heat transfer, shown in Figure 4 below. The MATLAB scripts used in this analysis can be found in Appendix D.

Figure 4: Ammonia Internal Flow Through Throttle Valve (100 bar)

The pipe diameter for the ammonia expulsion was analyzed in Figure 5, and a pipe diameter of 2 cm was chosen to balance heat transfer and mass.

Figure 5: Heat Transfer Insulation Thickness Trade Study

After expansion, the piping wraps around the exterior of the phase change thermal storage device in a helical fashion starting at the bottom and then travels outward to the lander body wall for expulsion, shown in Figure 6.

Figure 6: Ammonia Piping Around Phase Change Thermal Storage Device

78.38 kg of ammonia will be used for thermal management within the lander. It is estimated that a storage tank of 20 kg is necessary to store this quantity of ammonia. The mass of the piping for the expanded ammonia is estimated to be about 5 kg. The throat radius of the throttle valve has been calculated to be 4.3mm. The temperature change in the lander following ammonia depletion was modeled using a thermal circuit in Simscape. It was found that the lander would reach thermal equilibrium with the Venusian surface temperature within 24 hours after the loss of phase change thermal storage capability. The thermal circuit and temperature vs. time plot (in seconds) are shown in Figure 7 and Figure 8, respectively. The initial transient on the plot is the result of an arbitrary initial condition between the external and internal temperatures being selected. This quickly dissipates as the actual temperature states are modeled, but such a condition was required to ensure convergence on a reasonable solution.

Figure 7: Simscape Thermal Circuit Model

Figure 8: Temperature Differential Across Lander Shells vs. Time Following Ammonia Depletion

Scientific Instrumentation Suite The lander includes scientific instruments inspired by those included on the Venera landers and Pioneer Venus probes, missions to Mars’s surface, and the hypothetical Venera-D mission. The emphases of the science suite include atmospheric composition, surface composition, atmosphere and surface evolution, and detection of biosignatures. The scientific instruments selected are summarized in Table 7 below, along with their previous inclusion on missions and their purposes.

Table 7: Lander Scientific Instrument Manifest

To further enable scientific data collection, some scientific instruments were selected to be added to the orbital relay. The instruments selected were a gas chromatograph, a camera, and a Raman spectrometer. The instrument masses will be the same as those employed on the lander. Of these instruments, the spectrometers are the most relevant for studying the environment on Venus as it relates to Earth. They measure the composition and the evolution of the surface and atmosphere, which can help scientists determine whether Earth may one day undergo similar changes. In particular, the tunable laser spectrometer can examine greenhouse effects in the Venusian atmosphere, which are a growing concern on Earth as climate change approaches the point of no return.

Orbital Relay The orbital relay communications system is based off the system used for the Mars Reconnaissance Orbiter (MRO). This system is still in use today on the MRO, making it a suitable basis for the Venusian orbital relay. However, the thick atmosphere of Venus will make communications between the lander and relay more difficult and require more power to amplify the signal. Therefore, the system discussed will need to have a larger power draw than presented in the table regarding the communication system input power. This larger power draw from the communications system will require an increase in solar panel surface area to generate more power.

The solar panel array will consist of two panels with gallium arsenide solar cells and sub- strate. Gallium arsenide has an efficiency factor of 0.2432 for converting solar energy into electricity (Li et al, 2019). Gimbals are needed to make sure that the solar panels are directly facing the sun to generate the most electricity. 6 mm of cover glass will be implemented to further promote efficiency of the panels. Each panel has dimensions of 0.8 x 0.8 x 0.0064 m. Based on the surface area, the efficiency of gallium arsenide, the assumption of zero-degree angle between the panels and the sun, and the solar constant of Venus as 2,636 W/m2, the power generation will be equal to 410.3 W. The surface area will need to be increased to generate the power necessary for boosted communication, as previously mentioned. An Autodesk In- ventor simplified rendering of a solar panel is represented in Figure 9 below.

Figure 9: Simplified Solar Panel Rendering

The thermal control system consists of a heat pipe system and multilayer insulation. The piping is made from copper, and the working fluid is water. Assuming that the piping has a length of 1 m, an inner radius of 18 mm, and a thickness of 2 mm, its mass is 2.13 kg. 1.01 kg of water is required to fill this internal volume. The multilayer insulation weighs an additional 1.2 kg and further protects the onboard electronics from temperature changes. Due to the proximity to the Sun, radiation protection must be considered for the onboard instruments. The relay material must have a low mass and offer decent protection for the mission. Of the common radiation protection materials, aluminum was chosen. A thickness of 2 cm is sufficient to shield against high energy particles (Zeynali et al., 2012). The body is a cube with side lengths of 1.3 m. Based on the thickness and side dimensions, the mass of the aluminum is 549.6 kg.

Finally, a propulsion system is needed for the relay to maintain attitude control. The system consists of 10 maneuvering thrusters rated at 20 N. Due to its high density-specific impulse, monopropellant hydrazine was selected as the propellant. 40 kg of hydrazine propellant are needed for this attitude control; a spherical propellant tank made of Ti-6Al-4V houses this hydrazine. The mass of the tank is 28.2 kg for a thickness of 1 cm and an inner radius of 0.22 m. Information for the hydrazine thrusters is derived from ArianeGroup (2020). The overall system mass is approximately 74.7 kg. Overall renderings of the orbital relay are given in Figure 10 below; the rectangular pro- trusion on the side opposite the high gain antenna represents the camera.

Figure 10: Simplified Orbital Relay Views

Entry, Descent, and Landing The Entry, Descent, and Landing timeline is as follows. The insertion vehicle will perform aerodynamic braking over 64 km above the surface of Venus. At 64 km, the parachute will deploy, and it will then be jettisoned at 50 km above the surface after the entry vehicle has slowed down. Along with the parachute, the heat shield, and aeroshell will be released at 50 km. From 50 km above the surface, the lander will begin a control descent with control fins to the landing site. This is a similar flight profile to the Venera lander missions (Siddiqi, 2018).

The heat shield will be a titanium honeycomb filled with Avcoat ablative. Estimated to be 30% titanium and 70% ablative, the heat shield weighs 170 kg. The aeroshell is an aluminum honeycomb with Avcoat fill. Estimated to be 30% aluminum and 70% ablative, the aeroshell weighs 250 kg. Figure 11 shows the heat shield and aeroshell together.

Figure 11: Heat Shield and Aeroshell

The heat shield and aeroshell are sized to fit the lander inside while minimizing empty space within. The parachute was sized using a mass ratio comparing the Endurance lander to the Perseverance lander. The parachute has a diameter of 14.95 m and a mass of 56 kg. The parachute release speed was determined by performing load analysis on the control fins with a safety factor of 1.5. The fins are connected to the lander via a hinged mount with a back plate. This allows the fin to droop down while within the aeroshell. Once the lander is released from the aeroshell, the dynamic pressure pushes the fin upwards into the proper position. Analysis of the fin and connector was performed using Abaqus CAE. Safety allowances were included by multiplying the loads of the fins by the safety factor. Loads were estimated by altering the drag coefficient for a half Rankine body. The maximum stresses (819 MPa) occur at the hinge joint, shown in Figure 12 below.

Figure 12: Control Fin Stress Concentrations

These are slightly less than the yield strength of the titanium used. When accounting for the factor of safety, 360 m/s is the maximum allowable speed for parachute release at 50 km. Under these conditions, the maximum displacement, shown in Figure 13, of the fin tip is 1.94 cm, which is 2.5% of the total length of the fin; this was deemed allowable. The mass of each fin was determined to be 2.56 kg.

Figure 13: Control Fin Overall Displacements

Launch, Orbits, and Propellant Masses The launch date of the mission was determined based on mass and orbit data for Venus and Earth found in Curtis (2014). The synodic period between Earth and Venus was found to be 584 days, meaning that the optimum transfer window occurs once every 1.5 years. Using data for the most recent perihelion of both Earth and Venus as well as the orbit parameters, the first transfer available in the future was found to be in January 2022. Additional windows were found by moving forward in 584-day increments. To ensure time for the newer technologies on the lander to be developed and tested to a technology readiness level of 6, a launch window in August 2031 was selected. The Hohmann transfer orbit to Venus and the approach hyperbola were determined next. It was decided that the orbital relay would occupy an orbit ranging between 500 km and 15,000 km in altitude, and the vehicle would spend approximately 2.5 hours within a 45° cone over the landing site per orbit to enable data transmission. The lander would then enter a circular orbit 500 km in altitude to decrease velocity prior to the deorbit burn (which would place the pericythe within the Venusian atmosphere). Once the velocity changes resulting from these maneuvers were found, the required propellant mass was calculated using the system mass and a specific impulse of 230 s, which is an approximate value for monopropellant hydrazine (Curtis, 2014). A summary of the orbit parameters and the delta-V between them can be found in Table 8. The calculations conducted in this analysis can be found in Appendix B.

Orbit Parameter Value Approach Hyperbola Excess Velocity (km/s) 2.707 Aiming Radius (km) 24,974 Eccentricity 1.148 Semi-Major Axis (km) 44,323 Angular Momentum (km2/s) 67,618 Pericythe (km) 6,552 Delta-V to Relay Ellipse (km/s) 1.559 Relay Ellipse Eccentricity 0.534 Semi-Major Axis (km) 15,000 Angular Momentum (km2/s) 59,035 Pericythe (km) 6,552 Apocythe (km) 21,552 Delta-V to Lander Parking 1.679 Lander Parking Eccentricity 1.000 Radius (km) 6,552 Angular Momentum (km2/s) 46,138 Delta-V to Deorbit (km/s) 0.119 Table 8: Orbit Parameters and ∆V Summary

Following these calculations, the total spacecraft launch mass was found to be high enough to require launching on a Falcon Heavy. Since Endurance would be a NASA mission, it was assumed that only domestically produced launch vehicles would be allowed. The Delta IV Heavy is the only other currently operational launch vehicle with the necessary capabilities; however, it is no longer in production, and all existing vehicles are assigned to upcoming NRO launches (Henry, 2019). Other vehicles, such as the New Glenn, Vulcan Centaur, and Starship, do not yet have flight heritage. Therefore, the Falcon Heavy was the vehicle selected.

Landing Site It was determined that to minimize the temperature experienced by the lander, a high- altitude surface on Venus needed to be selected. The eastern slopes of Maxwell Montes in the Ishtar Terra region provide a relatively flat surface while also decreasing the average temperature (NASA JPL, 1996). Additionally, the region provides an interesting target for further research, as it is one of the largest mountains in the solar system.

Discussion of Results

Masses

Orbital Relay Mass for the orbital relay is a combination of the solar panel array, the scientific instruments onboard, the propulsion system, the communications system, the thermal control system, and the radiation shielding. There were no constraints placed on the mass of the orbital relay, but it was still designed to minimize unnecessary mass. The masses follow the AIAA Mass Properties Control for Space Systems standards. All systems are preexisting; therefore, they have a maximum percentage mass growth allowance of 3%. Table 9 below details the masses of the orbital relay, divided by system.

Table 9: Orbital Relay Mass Table

The communications system is further subdivided in Table 10 below.

Table 10: Orbital Relay Communications System Mass Table

Lander The mass of the lander consists of the body shells, phase-change thermal storage systems, total scientific instrument mass, descent control fins, legs, communication systems, and power systems. The total mass of 1,563 kg is higher than the 1,500 kg maximum; however, mass can be reduced by removing science instruments (leaving the 200 kg minimum) or reducing the amount of ammonia on board (at the cost of reduced mission time). As before, the mass growth allowance is assumed to be 3%. Table 11 below summarizes the masses of the lander, along with the cruise and EDL systems.

Table 11: Lander, Cruise, and EDL Mass Table

Overall The overall system mass consists of the lander, cruise stage, EDL system, orbital relay, and propellant; the total launch mass was found to be 8,542 kg. As before, the mass growth allowance of all systems is 3%, and Table 12 below summarizes the overall masses.

Table 12: Final Launch Masses

Power Draw

Orbital Relay The power draw of the orbital relay follows a similar standard as presented by the AIAA Mass Properties Control for Space Systems. Once again, all systems are preexisting, so they have a maximum percentage power growth allowance of 3%. As stated before, the communications system will require more power due to the denser atmosphere of Venus. Table 13 below details the power draw of the orbital relay, divided by system.

Table 13: Orbital Relay Input Power Table

Lander The lander power draw comes primarily from the scientific instruments, with additional power being drawn by the descent control fins during the EDL sequence, the phase change thermal storage system, and the communications system. As before, the power draw growth allowance is assumed to be 3%, and the power draws are shown in Table 14 below.

Table 14: Lander Power Draw Table

Concept of Operations The Concept of Operations (ConOps) was formulated using the launch window, launch vehicle, orbit parameters, and power draws found previously. The ConOps in its entirety can be found in Appendix C. The spacecraft conducts regular health checks with ground stations on Earth during its coast phase. Upon arrival at Venus, the vehicle directly inserts into the relay’s operational orbit. The relay separates and functions as an independent vehicle for the remainder of the mission. The relay orbit is such that its apocythe is over the landing site and the relay spends approximately 2.5 hours within 45° of the lander’s local vertical. This is possible due to Venus’s slow rate of rotation, which results in a ground track that does not deviate significantly during the mission. The lander and cruise stage enter a circular parking orbit to decrease velocity prior to entry. Following the deorbit burn, the cruise stage and relay adapter separate and destructively enter Venus’s atmosphere. The lander continues through the upper atmosphere within the heat shield and aeroshell. Once the hypersonic parachute deploys, the heat shield separates. The aeroshell and parachute separate next, and the lander is slowed by aerodynamic drag and controlled by its descent fins until touchdown. Following touchdown, the lander alternates between periods of recharging using the thermoelectric generator and collecting data using its scientific instruments. The lander’s interior temperature increases to 300°C during the 4 hours following heat shield separation, and the temperature is maintained at this value for 20 hours using the phase change thermal system. During this steady state period, each instrument is allocated 20 minutes of data

collection. The final data collection of the steady state period is an extended science period using the instruments deemed most relevant to studying Venus’s evolution. Once the ammonia is depleted, the lander’s interior temperature increases again. At this point, a different set of instruments powers up for continuous data collection until the end of the mission. The TEG output decreases due to the loss of the temperature differential, and the end of the mission occurs when the lander ceases transmissions due to heat damage or power depletion.

Impacts of Project on Engineering Solutions This project was designed to help NASA engineers with potential designs for a future Venusian lander. With data gained from Venus, questions surrounding planetary evolution and atmospheric development could be answered. Over several years of study, Mars has provided humanity with a wealth of information; Venus would provide a similar scientific reward once the challenges of the harsh environment are addressed. Atmospheric evolution is pertinent to Earth due to the increased greenhouse emissions from an increasingly developing population. Questions answered on Venus, such as how the accumulation of greenhouse gases started and continued over the course of the planet’s lifetime, could guide policies and practices taken on Earth regarding greenhouse gas emissions.

Conclusions and Recommendations The Endurance lander is designed to survive on Venus for 48 hours, which is approximately 24 times longer than the Venera 13 lander survived on the surface (NSSDCA, n.d.e). In addition, the lander brings over 200 kg of science instruments to the surface, including several instruments that have never been used on Venus. However, additional design refinement is required to reduce the mass below the 1,500 kg threshold. The Endurance lander design relies on a hypothetical thermoelectric material with a figure of merit of 20; thus, additional research is required to develop such a material. Other recommended research areas are batteries with higher energy densities that will survive high temperatures, low-power refrigeration systems, and phase-change thermal storage materials with high specific heats and high densities. To further develop the system, analyses of component system operations are recommended as well as additional work on system designs that are outside the expertise of the original developers. These recommendations include, but are not limited to: • Thermal analysis of the lander system with ammonia expulsion • Lander and relay communication system analysis and refinement • Wiring schematics and solutions to internal-external pressure/temperature changes regarding wiring • Researching more efficient and longer-lasting active cooling methods • Develop an interface through which science instruments can be exposed to the Venusian atmosphere • Simulation model of the EDL control system • Heat transfer and load analysis of heat shield and aeroshell accounting for ablative • Design of hypersonic parachute specific to the Venusian atmosphere • Research lighter materials to withstand Venusian pressure and temperature • Use generative design to optimize mass without sacrificing structural performance • Creating an internal layout for the scientific instruments • Detailed propulsion system design and analysis

Acknowledgements

The team would like to express their gratitude to their faculty advisor, Dr. Kristina Lemmer, for her guidance and support over the course of this project. In addition, the team would like to acknowledge Geoffrey Landis, Jeff Woytach, Roger Storm, Vanessa Webbs, Carol Tol- bert, David Kankam, and Coy Bush from NASA’s Glenn Research Center, who provided feedback and facilitated presentations as part of the University Student Design Challenge. Finally, the team would like to thank Western Michigan University’s College of Engineering and Applied Sciences for providing engineering software and the coursework needed to complete this project.

References

Aerogel Reinforced Composites (2020). In NASA Technology Transfer Program. https://technology.nasa.gov/patent/LEW-TOPS-141

Afework, B., Hanania, J., Heffernan, B., Jenden, J., Stenhouse, K., & Donev, J. (2018, July 21). Rankine cycle. Energy Education. https://energyeducation.ca/encyclopedia/Rankine_cycle

Afework, B., Hanania, J., Stenhouse, K., & Donev, J. (2018, June 25). Stirling engine. Energy Education. https://energyeducation.ca/encyclopedia/Stirling_engine

ArianeGroup. (2020). 1N, 20N, 400N and Heritage Thruster. In ArianeGroup. Retrieved from https://www.space-propulsion.com/brochures/hydrazine-thrusters/hydrazine- thrusters.pdf

Borghino, D. (2014, May 23). MIT finds new way to harvest energy from heat. New Atlas. https://newatlas.com/thermogalvanic-harvesting-waste-heat/32212/

Chung, D.Y., Hogan, T., Schindler, J., Iordanidas, L., Brazis, P., Kannewurf, C.R., Chen, B., Uher, C., & Kanatzidas, M.G. (1997, August). Complex bismuth chalcogenides as thermoelectrics. In Proceedings ICT’97. https://doi.org/10.1109/ICT.1997.667185

Clean Energy Institute. (n.d.) What is a lithium-ion battery and how does it work? University of Washington. https://www.cei.washington.edu/education/science-of-solar/battery- technology/

Curtis, H.D. (2014). Orbital mechanics for engineering students (3rd edition). Elsevier Ltd.

Emilio, M.D. (2019, December 11). New thermoelectric material has huge IoT potential. EE Times. https://www.eetimes.com/new-thermoelectric-material-has-huge-iot-potential/

Enescu, D. (2019, January 21). Thermoelectric harvesting: Basic principles and applications. In D. Enescu (Ed.), Green energy advances. IntechOpen. https://doi.org/10.5772/intechopen.83495

Fimmel, R.O., Colin, L., & Burgess, E. (1983). Pioneer Venus. NASA. https://atmos.nmsu.edu/data_and_services/atmospheres_data/PIONEER_ Venus/logs/Pioneer.pdf

Henry, C. (2019, August 21). ULA’s Delta IV Heavy down to final five missions. SpaceNews. https://spacenews.com/ulas-delta-4-heavy-down-to-final-five-missions/

Hinterleitner, B., Knapp, I., Poneder, M., Shi, Y., Müller, H., Eguchi, G., Eisenmenger-Sittner, C., Stöger-Pollach, M., Kakefuda, Y., Kawamoto, N., Guo, Q., Baba, T., Mori, T., Ullah, S., Chen, X., & Bauer, E. (2019, November 13). Thermoelectric performance of a metastable thin-film Heusler alloy. Nature, 576, 85–90. https://doi.org/10.1038/s41586- 019-1751-9

Hodge, Andrew & Kaul, Raj & McMahon, William. (2000). Sandwich composite, syntactic foam core based application for space structures. Proceedings of the 45th International SAMPE Symposium, 45, 2293-2304.

Hsu, K.F., Loo, S., Guo, F., Chen, W., Dyck, J.S., Uher, C., Hogan, T., Polychroniadis, E.K.,

& Kanatzidas, M.G. (2004, February 6). Cubic AgPbmSbTe2+m: Bulk thermoelectric materials with high figure of merit. Science, 303(5659), 818– 821. https://doi.org/10.1126/science.1092963

Jaziri, N., Boughamoura, A., Müller, J., Mezghani, B., Tounsi, F., & Ismail, M. (2020, December). A comprehensive review of thermoelectric generators: Technologies and common applications. Energy Reports, 6(7), 264– 287. https://doi.org/10.1016/j.egyr.2019.12.011

Kisi, E., Sugo, H., Cuskelly, D., Fiedler, T., Rawson, A., Post, A. Bradley, J., Copus, M., Reed, S. (2018, January 30), Miscibility Gap Alloys: A New Thermal Storage Solution. https://link.springer.com/chapter/10.1007/978-3-319-69844-1_48

Kudo, K. (2008, January). Overseas trends in the development of human occupied deep submersibles and a proposal for Japan’s way to take. Quarterly Review, 26, 104- 123. http://citeseerx.ist.psu.edu/viewdoc/download?doi=10.1.1.728.8907&rep=rep1&typ e=pdf

Krebs, G. (2021, February 12). Interplanetary probes. Gunther’s Space Page. https://space.skyrocket.de/doc_sat/ip_probe.htm

Li, Z., Kim, T., Han, S., Yun, Y., Jeong, S., Jo, B., ... Parl, H. (2019, December 19). Wide- Bandgap Perovskite/Gallium Arsenide Tandem Solar Cells. Advanced Energy Materials.

Lopes, R.M.C., & Gregg, T.K.P. (2004) Volcanic worlds: Exploring the solar system’s volcanoes. Springer Publishing.Parl, H. (2019, December 19). Wide- Bandgap Perovskite/Gallium Arsenide Tandem Solar Cells. Advanced Energy Materials.

Lukco, D., Spry, D. J., Neudeck, P. G., Nakley, L. M., Phillips, K. G., Okojie, R. S., & Hunter, G. W. (2020, November). Experimental Study of Structural Materials for Prolonged Venus Surface Exploration Missions. Journal of Spacecraft and Rockets, 57(6).

Lumen Learning (n.d.) The massive atmosphere of Venus. https://courses.lumenlearning.com/astronomy/chapter/the-massive-atmosphere-of- venus/

Mitchell, Don (2004). Inventing the interplanetary probe. Don P. Mitchell Homepage. http://mentallandscape.com/V_OKB1.htm

NASA Atmospheric Experiments Laboratory. (n.d.). Introduction to the Mass Spectrometer. https://attic.gsfc.nasa.gov/huygensgcms/Mass_Spec_Intro.htm

NASA Jet Propulsion Laboratory. (1996, February 5). PIA00149: Venus - Maxwell Montes and Cleopatra Crater. In Photojournal. Retrieved from https://photojournal.jpl.nasa.gov/catalog/PIA00149

NASA Jet Propulsion Laboratory [NASA JPL]. (n.d.a). Power systems. https://rps.nasa.gov/power-and-thermal-systems/power-systems/current/

NASA Jet Propulsion Laboratory [NASA JPL]. (n.d.b). Tunable laser spectrometer on NASA’s Curiosity Mars Rover. https://www.nasa.gov/jpl/msl/pia19086

NASA Mars Exploration Program [NASA MEP]. (n.d.a). APXS. https://mars.nasa.gov/msl/spacecraft/instruments/apxs/

NASA Mars Exploration Program [NASA MEP]. (n.d.b). ChemCam. https://mars.nasa.gov/msl/spacecraft/instruments/chemcam/

NASA Mars Exploration Program [NASA MEP]. (n.d.c). GRS. NASA Mars Odyssey. https://mars.nasa.gov/odyssey/mission/instruments/grs/

NASA Mars Exploration Program [NASA MEP]. (n.d.d). Mastcam-Z. NASA Mars 2020 Mission. https://mars.nasa.gov/mars2020/spacecraft/instruments/mastcam-z/

NASA Mars Exploration Program [NASA MEP]. (n.d.e). MEDA. NASA Mars 2020 Mission. https://mars.nasa.gov/mars2020/spacecraft/instruments/meda/

NASA Mars Exploration Program [NASA MEP]. (n.d.f). Mössbauer spectrometer (MB). NASA Mars Exploration Rovers. https://mars.nasa.gov/mer/mission/instruments/mb/

NASA Mars Exploration Program [NASA MEP]. (n.d.g). PIXL. NASA Mars 2020 Mission. https://mars.nasa.gov/mars2020/spacecraft/instruments/pixl/

NASA Mars Exploration Program [NASA MEP]. (n.d.h). RIMFAX. NASA Mars 2020 Mission. https://mars.nasa.gov/mars2020/spacecraft/instruments/rimfax/

NASA Mars Exploration Program [NASA MEP]. (n.d.i). SAM. https://mars.nasa.gov/msl/spacecraft/instruments/sam/

NASA Mars Exploration Program [NASA MEP]. (n.d.j). SHERLOC. NASA Mars 2020 Mission. https://mars.nasa.gov/mars2020/spacecraft/instruments/sherloc/

NASA Patents (n.d.). Aerogel Reinforced Composites. NASA technology Transfer Program. https://technology.nasa.gov/patent/LEW-TOPS-141

NASA Space Science Data Coordinated Archive [NSSDCA]. (n.d.a). Nephelometer (SN). https://nssdc.gsfc.nasa.gov/nmc/experiment/display.action?id=1978-078G-02

NASA Space Science Data Coordinated Archive [NSSDCA]. (n.d.b). Venera 5. https://nssdc.gsfc.nasa.gov/nmc/spacecraft/display.action?id=1969-001A

NASA Space Science Data Coordinated Archive [NSSDCA]. (n.d.c). Venera 9 descent craft. https://nssdc.gsfc.nasa.gov/nmc/spacecraft/display.action?id=1975-050D

NASA Space Science Data Coordinated Archive [NSSDCA]. (n.d.d). Venera 12 descent craft. https://nssdc.gsfc.nasa.gov/nmc/spacecraft/display.action?id=1978-086C

NASA Space Science Data Coordinated Archive [NSSDCA]. (n.d.e). Venera 13 descent craft. https://nssdc.gsfc.nasa.gov/nmc/spacecraft/display.action?id=1981-106D

NASA Solar System Exploration [NASA SSE]. (n.d.a). Ion and neutral mass spectrometer (INMS). https://solarsystem.nasa.gov/missions/cassini/mission/spacecraft/cassini- orbiter/ion-and-neutral-mass-spectrometer/

NASA Solar System Exploration [NASA SSE]. (n.d.b). INMS engineering technical write- up. https://solarsystem.nasa.gov/missions/cassini/mission/spacecraft/cassini-orbiter/ion- and-neutral-mass-spectrometer//inms-technical-write-up/

Petsagkourakis, I., Tybrant, K., Crispin, X., Ohkubo, I., Satoh, N., & Mori, T. (2018, November 14). Thermoelectric materials and applications for energy harvesting power generation. Science and Technology of Advanced Materials, 19(1), 836– 862. https://doi.org/10.1080/14686996.2018.1530938

Piggott, A. (n.d.). How thermoelectric generators work. Applied Thermoelectric Solutions. https://thermoelectricsolutions.com/how-thermoelectric-generators-work/

Quick, D. (2012, September 20). World’s most efficient thermoelectric material developed. New Atlas. https://newatlas.com/most-efficient-thermoelectric-material/24210/

Rioux, N. (2006, March). ANSI/AIAA guide for estimating spacecraft systems contingencies applied to the NASA GLAST mission. In 2006 IEEE Aerospace Conference. https://doi.org/10.1109/AERO.2006.1656180

Ross, F.E. (1928, July). Photographs of Venus. Astrophysical Journal, 68, 57- 92. https://doi.org/10.1086/143130

Siddiqi, Asif (2018). Beyond Earth: A Chronicle of deep space exploration. https://www.nasa.gov/sites/default/files/atoms/files/beyond-earth- tagged.pdf

Snyder, G.J., & Snyder, A.H. (2017, September 22). Figure of merit ZT of a thermoelectric defined from materials properties. Energy and Environmental Science, 10(11), 2280– 2283. https://doi.org/10.1039/C7EE02007D

Spaceflight101. (n.d.). Falcon Heavy countdown timeline. https://spaceflight101.com/falcon- heavy-countdown-timeline/

Thermoelectric generators. (2021, January 20). In Wikipedia. https://en.wikipedia.org/w/index.php?title=Thermoelectric_generator&oldi d=1001622987

Thermoelectric materials. (2021, February 3). In Wikipedia. https://en.wikipedia.org/w/index.php?title=Thermoelectric_materials&oldi d=1004664629

Titanium Alloys Ti6Al4V Grade 5. (2002, July 30). In U.S. Titanium Industry Inc.

Venera-D Joint Science Definition Team [Venera-D JSDT]. (2017). Venera-D: Expanding our horizon of terrestrial planet climate and geology through the comprehensive exploration of Venus. https://www.lpi.usra.edu/vexag/reports/Venera-D-STDT013117.pdf

WebElements. (n.d.). Lead telluride. https://www.webelements.com/compounds/lead/lead_telluride.html

Webster, C.R., & Mahaffy, P.R. (2009, November). Measuring methane and its isotopic ratios 13C/12C and D/H with the Tunable Laser Spectrometer (TLS) on SAM for the Mars Science Laboratory (MSL) mission [Workshop session]. Joint ESA-ASI Workshop on Methane on Mars. Frascati, Italy. https://sci.esa.int/documents/33745/35957/1567259793745- Methane2009_Session8_2_TLS_Webster.pdf

Wiens, R.C., Maurice, S., Barraclough, B., Saccoccio, M., Barkley, W.C., Bell, J.F., III, Bender, S., Bernardin, J., Blaney, D., Blank, J., Bouyé, M., Bridges, N., Bultman, N., Caïs, P., Clanton, R.C., Clark, B., Clegg, S., Cousin, A., Cremers, D., … Wong-Swanson, B. (2012, June 21). The ChemCam instrument suite on the Mars Science Laboratory (MSL) rover: Body unit and combined system tests. Space Science Reviews, 170, 167– 227. https://doi.org/10.1007/s11214-012-9902-4

Williams, David R. (15 April 2005). Venus fact sheet. NASA. https://nssdc.gsfc.nasa.gov/planetary/factsheet/venusfact.html

Zeynali, O., Masti, D., & Gandomkar, S. (2012). Shielding protection of electronic circuits against radiation effects of space high energy particles. Pelagia Research Library. Retrieved from https://www.imedpub.com/articles/shielding-protection-of-electronic- circuits-against-radiation-effects-of-spacehigh-energy-particles.pdf

Zhao, L., Lo, S., Zhang, Y., Sun, H., Tan, G., Uher, C., Wolverton, C., Dravid, V.P., & Kanatzidis, G. (2014, April 16). Ultralow thermal conductivity and high thermoelectric figure of merit in SnSe crystals. Nature, 508, 373– 377. https://doi.org/10.1038/nature13184