Thermal Architecture of A Conceptual Mars Sample Return

ICES-2020-46

Copyright © 2020 Jet Propulsion Laboratory, California Institute of Technology

Thermal Architecture of A Conceptual Mars Sample Return

Lander during Cruise and on Mars

Pradeep Bhandari1, Martin Greco2, Brant Cook3, Alexander Eremenko4, Carl Guernsey5, Ashley Karp6, Austin

Nicholas7

Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA 91109

A Pre-Project team is currently studying a conceptual Mars Sample Return (MSR)

architecture that would return samples collected by Mars 2020 to Earth1-2. The basic

architecture comprises of acquisition of these samples using a Sample Retrieval Lander (SRL),

which would also house the Mars Ascent Vehicle (MAV) and Sample Fetch Rover (SFR). The

MAV would put the orbiting sample container (OS) in a Martian orbit, which would then

rendezvous with an Earth return orbiter (ERO) and be sent to Earth. This paper focuses on

the SRL portion of the potential MSR campaign. The thermal architecture of this mission

during cruise to Mars presents several challenges that pertain to the thermal control of the

spacecraft and the lander/MAV/rover throughout cruise under varying thermal environments

& operating conditions. Additionally, the control of these systems within their allowable

operating temperature limits on the Martian surface is very challenging because of the large

fluctuations in the environment, operating conditions and limited electrical power and energy

availability. This paper will describe the thermal architecture for a potential SRL mission,

the key thermal requirements and interfaces.

Nomenclature

MSR = Mars Sample Return

SRL = Sample Retrieval Lander

SFR = Sample Fetch Rover

MAV = Mars Ascent Vehicle

ERO = Earth Return Orbiter

NEPA = National Environmental Policy Act

Ls = Solar Longitude

AFT = Allowable Flight Temperature (limit)

QUAL = Qualification (temperature limit)

Op = Operational

Non-Op = Non-Operational

I. Introduction

A Mars Sample Return (MSR) campaign to return samples back to Earth for analysis would be an exciting and

challenging effort both from the scientific as well as engineering points of view. A Pre-Project team is currently

studying a conceptual MSR architecture that would return samples collected by Mars 2020 to Earth. The basic

architecture comprises of acquisition of these samples using a Sample Retrieval Lander (SRL), which would also

house the Mars Ascent Vehicle (MAV) and Sample Fetch Rover (SFR). The MAV would put the orbiting sample

1 Principal Thermal Engineer, 4800 Oak Grove Dr. M/S 125-123, Pasadena, CA 91109 2 Flight System Engineer, 4800 Oak Grove Dr. M/S 79-24, Pasadena, CA 91109 3 Mechanical Engineer, 4800 Oak Grove Dr. M/S 125-123, Pasadena, CA 91109 4 Systems Engineer, 4800 Oak Grove Dr. M/S 157-316, Pasadena, CA 91109 5 Principal Systems Engineer, 4800 Oak Grove Dr. M/S 157-316, Pasadena, CA 91109 6 Technologist, 4800 Oak Grove Dr. M/S 157-316, Pasadena, CA 91109 7 Systems Engineer, 4800 Oak Grove Dr. M/S 157-316, Pasadena, CA 91109

International Conference on Environmental Systems

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container (OS) in a Martian orbit, which would then rendezvous with an Earth return orbiter (ERO) and be sent to

Earth. A schematic for this campaign is show in Figure 1. The three key elements of this mission are shown in

Figure 2.

Figure 1. Notional Mars Sample Return Mission.

Figure 2. SRL and ERO would collect and return the sample of Mars to be gathered by Mars 2020.

The SRL portion of the potential MSR campaign presents a challenging set of thermal design constraints because

the lander has to survive the harsh environments on Mars for a surface mission duration of 430 Sols. The thermal

control of the lander components within their allowable operating temperature limits on the Martian surface is very

challenging because of the large fluctuations in the environment, operating conditions and limited electrical power

and energy availability. Key current schematics of the SRL are shown in Figure 3.


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Figure 3. Sample Return Lander Concept (On Mars, Top; Stowed During Cruise in Aero-Shell, Bottom).

II. Thermal Environment on Mars

The landing site of the SRL is close to where the Mars 2020 rover would land (Jezero crater, 18.5 N) to be in close

proximity to where M2020 would deposit the Martian samples. The Sample Fetch Rover (SFR) would fetch the

samples from the Martian surface and bring them back to the SRL. The Martian thermal environment at Jezero during

the SRL’s landed mission (spring, Ls = 0° through summer, Ls = 90°) is shown in Figure 4 with Martian ground

temperatures ranging between -90°C and 14°C. Luckily, due to the opposition of the orientation of the tilt of Mars

and its orbit eccentricity, the Jezero latitude has a relatively flat thermal environmental profile during a Martian year,

when compared to other latitudes. Due to the lander being powered by solar arrays only, the deposition of dust on the

solar arrays obscuring the sun severely constraints the timing of the mission on Mars with the key driver being the

avoidance of global dust storms that start after Ls = 180°.

Figure 4. Jezero Thermal Environment (Coldest, Left; Hottest, Right)


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III. Key Thermal Subsystems of the SRL

There are two key elements of the SRL that require special thermal attention: The Lander per se and the Mars

Ascent Vehicle (MAV) and these are depicted in Figure 5. Since the SRL configuration and design is in its early

phase and is evolving, the key emphasis in this paper is on the thermal architecture, rather on the detailed thermal

design. The lander thermal architecture adopts to some extent the basic architecture of the Mars InSight mission that

is currently on the surface of Mars. It thermally isolates the key components of the lander - principally the avionics

and the battery - from the Martian environment. During cruise, heat pipes transfer the heat from the lander to radiators

viewing the cold heat shield. The heat pipes stop working on Mars due to them being oriented in the adverse gravity

direction allowing the lander conserve its heat. The MAV is ensconced in the lander inside a cocoon that protects it

from the Martian environment as shown in Figure 5. The MAV is thermally isolated from its cocoon by a gap of ≥10

cm that is naturally filled with CO2 at 8 Torr from the Martian environment. This CO2 insulation gap design has been

successfully flown on Mars in the Mars Science Laboratory (MSL) Curiosity rover. Additionally, low emissivity

surfaces are employed on the surface of the MAV and the inside of the cocoon to provide radiative isolation between

the two.

Figure 5. SRL/MAV Configuration (Isometric of SRL, left; Avionics/Battery @ Bottom of SRL, right)

IV. Key and Driving Thermal Requirements

The current temperature limits of key components of the SRL, including the MAV are shown in Table 1. The

avionics have the typical temperature limits for hardware procured by JPL for Mars missions, with a 15 C margin

between the minimum allowable and qualification limits and 20 C for the maximum. The battery also has the standard

limits for a lithium ion unit. The MAV is currently baselined to provide propulsion based on solid fuels. The solid

fuel has a storage or non-operational minimum limit of -40 C, with a maximum of +40 C. During operation (firing)

it needs to be warmed up to -20 C. Additionally, there is a placeholder thermal gradient requirement of ≤10 C in any

direction to ensure uniform burn for minimizing the spread in the imparted impulse and minimum fuel mass.

Non-Op Qual Non-Op AFT Op Qual Op AFT

MAV (Fuel) -50 to +50 C -40 to +40 C -35 to -5 C -25 to -15 C

Avionics -55 to +70 C -40 to +50 C -55 to +70 C -40 to +50 C

Battery -10 to +40 C -10 to +40 C -10 to +40 C 0 to +30 C

Table 1. Key Thermal Requirements of SRL.

V. Overall Drivers & Constraints for Lander Thermal Architecture

The overall approach to arrive at the thermal architecture for the SRL mission was to maintain heritage as much

as possible from previous successful missions, the InSight and Phoenix Mars Landers. The minimization of

power/energy & mass of thermal system as well as ramifications on rest of lander was an important goal, particularly

since this is a non-nuclear powered mission relying entirely on the power from solar arrays. Large solar arrays have


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to be stowed and accommodated in the aero shell to carry the SRL to Mars, which dictates the maximum size that can

be accommodated. In order to simplify the architecture a passive thermal system was baselined. Ease of

implementation and testing was also an important driver, and since this architecture has already been implemented

and tested before (InSight/Phoenix), it makes it attractive for use in the SRL. Cruise is ~2.5 years long and the mission

duration on Mars is about 6 Earth months. The heaters used for supplying power to the lander and the MAV are

primarily flight software controlled with some being controlled by mechanical thermostats. Software control allows

for set point changes in flight, whereas mechanical ones do not.

VI. Lander Thermal Architecture

The basic thermal architecture of the SRL is depicted in Figures 6 & 7. It consists of the avionics mounted inside

an enveloping chassis that in turn is isolated from the primary lander structure by struts. The lander battery is further

isolated thermally from the avionics to allow for battery to be held at a warmer temperature to meet its higher minimum

operating temperature limit. The MAV is also supported by and thermally isolated from the lander’s primary structure.

The entire lander is thermally isolated from the Mars environment by insulation around it.

Figure 6. Lander Thermal Schematic (I).

The insulation options are simple geometric gaps between the thermally controlled components and enveloping

chassis (exposed to the environment). These gaps are prescribed to be at least 5 cm thick and naturally get filled by

the CO2 on Mars. 5 cm gaps were chosen to be based on research1 done by JPL for the MSL Curiosity rover in which

it was shown by analysis and testing that this is the optimal thickness to minimize heat loss by conduction through

them - gaps larger than these do not yield to any higher thermal isolation due to convective cells that form in these

gaps. For gaps equal to or smaller than ~5 cm heat transfer is by gas conduction only. Another option for thermal

isolation is the employment of batt insulation, although it may not be as effective as CO2 gaps and obviously costs

some mass. Batt insulation is not as effective in space as MLI blankets would be. Hence, to allow for effective

thermal isolation, MLI blankets would supplement the batt insulation. Batt insulation would work on Mars, and MLI

during cruise in a vacuum environment. MLI would not be effective on Mars due to thermal shorting by the CO2 gas.

Basic Thermal Architecture of SRL Lander- Cruise & Mars Surface

Cruise Martian Surface

• Cruise: Heat Pipe works to transfer heat from Lander avionics to Radiator → 0-g, No concerns with adverse direction

• Mars Surface: Heat Pipe stops working → Condenser below evaporator pools liquid & cannot replenish to evaporator in 0.4 g

3D view


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Figure 7. Lander Thermal Schematic (II).

During cruise the heat pipes attached to avionics and battery chassis connect it to radiators that view the cold heat

shield. Since heat pipes work in any orientation in zero gravity, they effectively remove heat from these components

during cruise. However, on the Martian surface, since the heat pipes have the heat sources (avionics attached to their

evaporators and their condensers attached to the heat sink (radiators), they stop operating because the heat pipe

working fluid is unable to return to the evaporators due to gravity overpowering the wicking forces. Hence the heat

pipes essentially function like passive heat switches and allow for their complete thermal isolation from Martian

environment except for the parasitic heat flow from their thermal insulation. During powered conditions when the

avionics dissipate power, the thermal mass of the lander avionics and battery in conjunction with cold Martian

atmosphere significantly dampens the temperatures of the avionics and prevents them from overheating – their

temperatures rise during heat dissipation and falls when the dissipation stops with the thermal mass damping the

amplitude of changes.

VII. MAV Thermal Architecture

The MAV is surrounded by a protective cocoon to protect it from the cold Martian environment to minimize heater

power. This cocoon is separated from the MAV exterior by a 10 cm thick CO2 gas gap. The cocoon would be

thermally isolated from the MAV by isolating supports to minimize parasitic heat loss through them. Additionally,

the surface of the MAV and the inner surface of the cocoon is covered with low emissivity tape to minimize heat loss

from the MAV. Since a 10 cm gap is too large to minimize heat loss via a CO2 gas gap (due to convection cells

forming in large gaps), an intermediate layer of low emissivity aluminized Mylar/Kapton is utilized to break this gap

into two 5 cm gaps. Film heaters would be installed on the MAV’s exterior surface to provide the necessary survival

or operational heat for maintaining the required temperatures. The entire MAV is about 3 m long and 0.6 m in

diameter. Figure 8 depicts the MAV configuration.


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Figure 8. MAV Configuration (Clockwise from top left: MAV inside thermal enclosure with open

enclosure; Blow up of MAV, enclosure & lander; MAV outside of enclosure; Cross section of enclosure)

VIII. Preliminary Thermal Assessments

To help size the solar arrays simple thermal models have been created to approximately estimate the heater powers

required for maintaining survival and operational temperatures. Since the configuration is still in its evolving phase,

these thermal models are in their early stages, hence very approximate in nature. For the MAV ~30-40 W survival

power would be required to maintain it at its -40 C survival temperature and ~60-70 W for operation at -20 C. The

Lander avionics would need ~10 W in addition to the battery needing ~20 W to maintain their respective minimum

survival temperatures. For the avionics, the minimum dissipated power in their survival mode should be close

sufficient to overcome the heat loss during survival, hence would not need to be augmented by heaters.

IX. Conclusions

The thermal architecture of the Mars sample return lander and the Mars Ascent vehicle concepts presented in this

paper combines maximum utilization of heritage from previous mission (InSight/Phoenix), is passive and relatively

simple and easy to implement. The key goal of the thermal architecture is to minimize thermal power due to the severe

limitations imposed by the sizing of the solar arrays. The lander utilizes thermal mass, thermal isolation and heat

pipes to arrive at this architecture. The MAV utilized a CO2 gas gap cocoon to maintain its temperatures by

minimizing heat loss from it to the cold Martian environment.

Acknowledgments

The work described in this paper was carried out at the Jet Propulsion Laboratory, California Institute of

Technology, under a contract with the National Aeronautics and Space Administration. Copyright 2020 California

Institute of Technology. Government sponsorship acknowledged. It should be noted that the decision to implement.

MSR will not be finalized until NASA’s completion of the National Environmental Policy Act (NEPA) process. This

document is being made available for information purposes only.


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References 1NASA, “Mission to Mars: Mars Sample Return,” accessed June 30, 2020, https://www.jpl.nasa.gov/missions/mars-sample-

return-msr/ 2NASA, “Mission to Mars: Mars Sample Return,” accessed June 30, 2020, https://mars.nasa.gov/mars-

exploration/missions/mars-sample-return/

Documents

Thermal Architecture of A Conceptual Mars Sample Return