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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2018

Creating a Voyager Thermal Model 39 Years Into the Flight Mission, Along With Model Correlation and Application

After 39 years of continuous operation in space, the output of the Voyager 1 & 2 spacecraft Radioisotope Thermoelectric Generator (RTG) power systems has decreased to the point where managing the power margin and maintaining thermal control has become increasingly difficult. As the total power dissipation in the bus has decreased, propellant line temperatures and margin above minimum Allowable Flight Temperature (AFT) have decreased, creating risk of the hydrazine freezing (at 1.6°C). This is further complicated by the lack of existing thermal models that can be used to assess propellant tank and line temperatures. In 2014, an effort was begun to create a Voyager spacecraft thermal model for that purpose. A steady-state Thermal Desktop model has been created from scratch over the past two years. Applied Sciences Laboratory (ASL) started the initial thermal model development under contract to Jet Propulsion Laboratory (JPL). The effort relied primarily on archived manufacturing drawings, limited documentation, interviews of senior engineers who worked on the Voyager design and implementation, and the experience of the Voyager Flight Operations team. Data from the Voyager System Thermal Vacuum (STV) tests is no longer available, making it necessary to correlate the model to more recent flight data and small in-flight tests. Correlation was achieved to within ±5°C for a hot case and a cold case (both data sets from 2014). However, the flight system has very few temperature sensors directly on propellant lines. Therefore, the task remains to determine how best to use the model, in conjunction with flight data, to make sure the Voyagers can continue to fly successfully.

Ledeboer, William C.↗

NASA NTRS2018

Thermal Assessment of OSIRIS-REx OVIRS Cryogenic Instrument During Flight System TVAC Test and in Flight

The Origins, Spectral Interpretation, Resource Identification, Security, Regolith Explorer (OSIRIS-REx) Visible and Infrared Spectrometer (OVIRS) is a cryogenic instrument. At the Outbound Cruise nominal spacecraft attitude, sunlight impinges on several multilayer insulation blankets on the forward deck. It is reflected or scattered to other components on the deck. This solar illumination adds heat load to the OVIRS, and causes its detector temperature to exceed the 105K maximum operating allowable flight temperature limit by 0.8K. During the flight system thermal vacuum test, the solar simulator beam reflected or scattered from the test fixtures to the OVIRS added non-flight heat load. The detector temperature was 9K warmer than that in flight. At those temperatures, the science data was acceptable, despite its quality was not as high as that of 105K or colder.

"OSIRIS-REx; OVIRS; cryogenic instrument; parasiti↗

NASA NTRS2007

Probabilistic Thermal Analysis During Mars Reconnaissance Orbiter Aerobraking

A method for performing a probabilistic thermal analysis during aerobraking has been developed. The analysis is performed on the Mars Reconnaissance Orbiter solar array during aerobraking. The methodology makes use of a response surface model derived from a more complex finite element thermal model of the solar array. The response surface is a quadratic equation which calculates the peak temperature for a given orbit drag pass at a specific location on the solar panel. Five different response surface equations are used, one of which predicts the overall maximum solar panel temperature, and the remaining four predict the temperatures of the solar panel thermal sensors. The variables used to define the response surface can be characterized as either environmental, material property, or modeling variables. Response surface variables are statistically varied in a Monte Carlo simulation. The Monte Carlo simulation produces mean temperatures and 3 sigma bounds as well as the probability of exceeding the designated flight allowable temperature for a given orbit. Response surface temperature predictions are compared with the Mars Reconnaissance Orbiter flight temperature data.

Dec, John A.↗

NASA NTRSDate not supplied

LOFTID Aeroshell Thermal Response Uncertainty Analysis Utilizing the End-to-End Monte Carlo Approach

A probabilistic thermal margin process has been performed to design the flexible thermal protection system (FTPS) and select entry trajectory constraints given an acceptable risk level for the Low-Earth Orbit Flight Test of and Inflatable Decelerator (LOFTID) project. Uncertainties exist in atmospheric entry aeroheating environments and the predicted thermal response of thermal protection system (TPS) material. Entry vehicle TPS is often over-sized to mitigate uncertainty by combining conservative bounding scenarios together. The probabilistic thermal margin process allows engineers to make informed aeroshell design, flight design, and FTPS performance risk trades while preventing excessive FTPS margin from being applied. This paper describes the uncertainty analysis methodology used to carry out a probabilistic thermal margin process used for LOFTID’s aeroshell and explains how the calculated probability of exceeding flight allowable temperatures is used to design the FTPS and establish nominal flight design constraints. This probabilistic thermal margin process had never been applied to entry, decent, and landing design for a flown entry vehicle and it is one of the LOFTID project’s goals to demonstrate its merits.

Steven Andrew Tobin↗

NASA NTRS2019

Follow-on studies using the Voyager spacecraft thermal model

The 42-year-old Voyager mission, now named the Voyager Interstellar Mission (VIM) is operating long beyond its design life. In 2012, Voyager 1 crossed the heliopause into interstellar space and Voyager 2 made the same transit in November 2018. Due to declining power output from the Radioisotope Thermoelectric Generators (RTGs) the Science and Flight Operations teams continue to make difficult choices in terms of managing both the power and thermal margins to preserve critical science observations and maintain the health of the two spacecraft. A previous paper, “Creating a Voyager Thermal Model 39 Years Into the Flight Mission, Along With Model Correlation and Application” described how a thermal model of these 1977 spacecraft was developed and correlated without many design artifacts and with limited temperature telemetry. This paper describes how the thermal model has been used to establish an Allowable Flight Temperature (AFT) limit for hydrazine propellant in the propulsion subsystem to minimize the risk of freezing. Voyager 2 temperatures have already descended to this limit in the vicinity of the Roll thruster propellant lines. The Flight Operations team has investigated several ways of detecting propellant freezing based on analysis and trending of thruster performance telemetry. In addition, the Voyager thermal model is being used to predict the spacecraft response to possible changes in power state. These changes could involve turning off outboard science instruments and/or their heaters to increase power margin and hence power dissipation inside the spacecraft bus (i.e. in Bay 7, where the power regulation electronics are located). Changes might also be made to turn on or off other loads inside the bus to more effectively heat the coldest propellant lines.

Medina, Enrique↗

NASA NTRS2019

Follow-on studies using the Voyager spacecraft thermal model

Medina, Enrique↗

NASA NTRS2016

Thermal Cycling for Development of Hybrid Fuel for a Notional Mars Ascent Vehicle

A Mars Ascent Vehicle (MAV) would be a critical part of a potential Mars sample return campaign. As currently envisioned, the MAV would launch samples from Mars into Mars orbit. Challenging requirements would be placed on a MAV as it would need to be able to survive in the harsh Martian environment which can get as cold as -120°C in the winter. At these cold temperatures, survival heating could be necessary to keep the MAV within the allowable flight temperature range. Estimates for survival heating power show that a fuel capable of operating at low temperature would significantly reduce the total required energy of the MAV. This has led to consideration of a hybrid propulsion system due to its anticipated low temperature capability. A new wax-based fuel, SP7, has been developed for a notional MAV mission. Thermal modeling of the fuel in the Mars environment was completed in order to estimate the temperature response of the fuel during Mars diurnal temperature cycling. Preliminary thermal testing was conducted to determine the max temperature range, radial gradient, and ramp rate the fuel could tolerate without cracking. The modeling and preliminary testing was used to develop a comprehensive, 100-day thermal cycling campaign. A total of 200 cycles are planned that are representative of 200 Martian diurnal cycles, roughly one-third of a Martian year. After 150 cycles, the fuel samples show no signs of cracking that would jeopardize the rocket’s performance. Though not yet complete, this testing is showing that SP7 has the potential to successfully be used on a MAV mission.

Farias, Edgardo↗

NASA NTRS2013

Dreaming on Mars: How Curiosity Performs Actuator Warm-Up While Sleeping

Before the Curiosity rover can perform its science activities for the day, such as driving, moving its robotic arm, or drilling, it first has to ensure that its actuators are within their allowable flight temperatures (AFTs). When the rover is awake, flight software uses heaters to warm up and maintain thermal zones at operational temperatures. However, Curiosity spends about 70% of its time sleeping, with the flight computer off, in order to conserve energy. Dream Mode is a special behavior that allows the rover to execute warm-up activities while sleeping. Using Dream Mode, actuators can be warmed up to their AFTs before the flight computer wakes up and uses them - saving power and improving operational efficiency. This paper describes the motivation behind Dream Mode, how it was implemented and tested on Curiosity, and the challenges and lessons learned along the way.

fault protection↗

NASA NTRS2003

A Thermal Analysis Approach for the Mars Odyssey Spacecraft's Solar Array

There are numerous challenges associated with placing a spacecraft in orbit around Mars. Often. trades must be made such as the mass of the payload and the amount of fuel that can be carried. One technique employed to more efficiently place a spacecraft in orbit while maximizing payload mass (minimizing fuel use) is aerobraking. The Mars Odyssey Spacecraft made use of aerobraking to gradually reduce its orbit period from a highly elliptical insertion orbit to its final science orbit. Aerobraking introduces its own unique challenges, in particular, predicting the thermal response of the spacecraft and its components during each aerobraking drag pass. This paper describes the methods used to perform aerobraking thermal analysis using finite element thermal models of the Mars Odyssey Spacecraft's solar array. To accurately model the complex behavior during aerobraking, the thermal analysis must be tightly coupled to the spatially varying, time dependent aerodynamic heating analysis. Also, to properly represent the temperatures prior to the start of the drag pass. the model must include the orbital solar and planetary heat fluxes. It is critical that the thermal behavior be predicted accurately to maintain the solar array below its structural flight allowable temperature limit. The goal of this paper is to describe a thermal modeling method that was developed for this purpose.

Dec, John A.↗

NASA NTRS2017

Thermal Control System of the ECOsystem Spaceborne Thermal Radiometer Experiment on Space Station (ECOSTRESS)

The ECOSystem Spaceborne Thermal Radiometer Experiment on Space Station (ECOSTRESS) is a thermal infrared imaging multispectral scanner science mission. ECOSTRESS is designed and built by the NASA Jet Propulsion Laboratory and will be deployed on the International Space Station (ISS). ECOSTRESS will measure the water loss from growing leaves and the evaporation from the soil by measuring the temperature of plants and will gather data on the health of the agriculture system on Earth. The ISS orbit will allow ECOSTRESS to take observations at different times during each day over the seasons to provide coverage over the contiguous United States. The ECOSTRESS Thermal Control System (TCS) consists of a combination of active and passive components to maintain the ECOSTRESS components within the allowable flight temperature (AFT) limits. The active thermal control systems include mechanical cryocoolers, heaters and a single-phase pumped fluid loop for Instrument and Payload. The focal plane detector is cooled to 65K by a pair of mechanical cryocoolers and a third mechanical cryocooler cools an intermediate cold shield to 130K. The pumped fluid loop transfers the instrument waste heat to JAXA’s Japanese Exposed Module External Facility (JEM-EF) provided external fluid loop before rejection to space from the ISS radiators. The passive TCS includes multi-layer and single layer insulations, flexible thermal links and coatings on the radiometer. This paper describes the ECOSTRESS instrument TCS architecture, instrument thermal requirements and key design drivers, the top level thermal design and analysis approach, and reports preliminary test results.

Carroll, Brian↗

NASA NTRS2021

Thermal Systems Modeling of Chemical Integrated Power Source (CHIPS) to Survive Lunar Night Environments

This paper presents the results of the systems level thermal modeling for a conceptual Chemical Heat Integrated Power Source (CHIPS). This proposed system offers a combined thermal and electrical power source to support survival of spacecraft operating in extreme low temperature lunar environments, without the use of radioisotope-based sources. A conceptual design study has been completed for this system, that uses heat generated by an exothermic chemical reaction in place of radioisotope or electrical heat sources. The goal of the study was to evaluate the feasibility of such as system through thermodynamic and chemical analysis and thermal modeling, and to identify technology gaps to inform a technology development and maturation plan. The specific technical objectives focused on delivery of 90-100 Wth thermal power and 30-40 We electrical power for 336 hours to a representative Commercial Lunar Payload Services (CLPS) lander, to support lunar surface survival and limited operations through a lunar night. The total system mass was targeted at ≤50 kg. A highly exothermic chemical reaction system is used to generate on-board electrical power for spacecraft systems, and thermal power to maintain critical spacecraft/lander systems within their allowable flight temperature (AFTs). Based on the very high energy content of the chemical reaction system, a much higher amount of heat per unit mass can be delivered to the spacecraft, relative to a rechargeable lithium-ion battery and electrical heater(s). Since radioisotope heaters or generators are not used, the system will be orders of magnitude lower in cost than a radioisotope heating/power unit, without the attendant regulatory complexities. As part of the CHIPS concept, a fraction of the thermal power generated is converted to electrical power via an appropriate thermal-to-electric converter technology (such as a free piston Stirling converter or a thermoelectric generator module), to provide power to critical loads. This approach is ideally suited to support operation of commercial landers (e.g., via the CLPS program). In most cases, these landers are only designed to operate during a portion of the lunar day, with no provision for survival through the lunar night. By supporting lunar night survival, a mission can be extended through the lunar night and at least into another lunar day, thus turning a nominal eight-day mission into a 36-day mission. Therefore, the current system demonstration will focus on scalability to support at least 336 hours (one lunar night) of continuous thermal and electrical power generation. Although initially targeted to support lunar equatorial landings, the technology is extensible to missions at the lunar poles, other extreme environments in the Solar System or even air-independent applications on Earth (e.g., ocean exploration).

West, William↗

NASA NTRS2021

Thermal Systems Modeling of Chemical Heat Integrated Power Source (CHIPS) to Survive Lunar Night Environments

This paper presents the results of the systems level thermal modeling for a conceptual Chemical Heat Integrated Power Source (CHIPS). This proposed system offers a combined thermal and electrical power source to support survival of spacecraft operating in extreme low temperature lunar environments, without the use of radioisotope-based sources. A conceptual design study has been completed for this system, that uses heat generated by an exothermic chemical reaction in place of radioisotope or electrical heat sources. The goal of the study was to evaluate the feasibility of such a system through thermodynamic and chemical analysis and thermal modeling, and to identify technology gaps to inform a technology development and maturation plan. The specific technical objectives focused on delivery of 90 to 100 Wth thermal power and 30 to 40 We electrical power for 336 hours to a representative Commercial Lunar Payload Services (CLPS) lander, to support lunar surface survival and limited operations through a lunar night. The total system mass was targeted at ≤50 kg. A highly exothermic chemical reaction system is used to generate on-board electrical power for spacecraft systems, and thermal power to maintain critical spacecraft/lander systems within their allowable flight temperature (AFTs). Based on the very high energy content of the chemical reaction system, a much higher amount of heat per unit mass can be delivered to the spacecraft, relative to a rechargeable lithium-ion battery and electrical heater(s). Since radioisotope heaters or generators are not used, the system will be orders of magnitude lower in cost than a radioisotope heating/power unit, without the attendant regulatory complexities. As part of the CHIPS concept, a fraction of the thermal power generated is converted to electrical power via an appropriate thermal-to-electric converter technology (such as a free piston Stirling converter or a thermoelectric generator module), to provide power to critical loads. This approach is ideally suited to support operation of commercial landers (e.g., via the CLPS program). In most cases, these landers are only designed to operate during a portion of the lunar day, with no provision for survival through the lunar night. By supporting lunar night survival, a mission can be extended through the lunar night and at least into another lunar day, thus turning a nominal eight-day mission into a 36-day mission. Therefore, the current system demonstration will focus on scalability to support at least 336 hours (one lunar night) of continuous thermal and electrical power generation. Although initially targeted to support lunar equatorial landings, the technology is extensible to missions at the lunar poles, other extreme environments in the Solar System or even air-independent applications on Earth (e.g., ocean exploration).

West, William↗

NASA NTRS2012

The Mars Science Laboratory (MSL) MMRTG In-Flight: A Power Update

The MSL Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) was fueled on October 28, 2008 by the Department Of Energy (DOE) in preparation for a late 2009 launch. Shortly after, the MSL launch was delayed approximately 2 years until 2011. The fueled MMRTG was placed in storage to await the new date for liftoff. Occasional measurements of the MMRTG's power output were taken and compared with power predictions that pre-dated fueling. An error in the predictive models was quickly recognized and remedied. The predictions, while improved, carried significant uncertainty. This uncertainty did not deter the launch of MSL, but did alter the planned mission on the surface of Mars. Once launched, the MSL spacecraft provided a hi-fidelity telemetry stream measuring the generator's electrical and thermal performance. These data were used to update the predictive models and a new prediction of the performance of the MMRTG on the surface of Mars was run just before Entry, Descent, and Landing (EDL) at Mars. The MSL MMRTG is working extremely well, providing power above predictions and operating within its flight allowable temperature limits. The generator was producing approximately 114 W at the beginning of the surface mission. This paper will elaborate on power modeling for the MSL MMRTG along with a review of some of the data recorded from the MSL cruise to Mars, EDL, and the early days of the surface mission.

thermoelectric generator↗

NASA NTRS2013

The Mars Science Laboratory (MSL) MMRTG In-Flight: A Power Update

The MSL Multi-Mission Radioisotope Thermoelectric Generator (MMRTG) was fueled on October 28, 2008 by the Department Of Energy (DOE) in preparation for a late 2009 launch. Shortly after, the MSL launch was delayed approximately 2 years until 2011. The fueled MMRTG was placed in storage to await the new date for liftoff. Occasional measurements of the MMRTG's power output were taken and compared with power predictions that pre-dated fueling. An error in the predictive models was quickly recognized and remedied. The resultant predictions, while improved, carried significant uncertainty. This uncertainty did not deter the launch of MSL, but did alter the planned mission on the surface of Mars. Once launched, the MSL spacecraft provided a hi-fidelity telemetry stream measuring the generator's electrical and thermal performance. These data were used to update the predictive models and a new prediction of the performance of the MMRTG on the surface of Mars was run just before Entry, Descent, and Landing (EDL) at Mars. The MSL MMRTG is working extremely well, providing power above predictions and operating within its flight allowable temperature limits. The generator was producing approximately 114 W at the beginning of the surface mission. This paper will elaborate on power modeling for the MSL MMRTG along with a review of some of the data recorded from the MSL cruise to Mars, EDL, and the early days of the surface mission.

thermoelectric↗

NASA NTRS2017

Design and Development of the Heat Redistribution System for the Mission to Europa Spacecraft:

This paper details the conceptual design of the Heat Redistribution System (HRS) used for thermal control of the Europa Clipper spacecraft. The solar powered spacecraft will launch in the early 2020s and will study Jupiter’s Icy moon, Europa, where previous investigations have indicated the possibility of a liquid water ocean underneath its ice-encrusted surface. The spacecraft will execute multiple flybys of Europa while relying on the HRS for thermal control of the spacecraft bus. The HRS utilizes a Mechanically Pumped Fluid Loop (MPFL) to reclaim waste heat from an avionics and payload electronics compartment and redistribute it to other parts of the spacecraft, such as the propulsion module structure and the louver-radiator assembly. A “replacement heater block” provides supplemental heat when needed to maintain the spacecraft hardware above its allowable flight temperatures. Additionally, a set of thermal control valves autonomously actuate to reduce the fluid flow through the radiator during cold environment cases. The HRS was used in previous flight projects at JPL such as Mars Science Laboratory, but the uniqueness of the Europa Clipper mission has mandated changes to previous HRS designs; the most significant design drivers and design changes for this iteration of the HRS are discussed.

Bertagne, Christopher↗

NASA NTRS2017

Mid Infrared Instrument Cooler Subsystem Test Facility Overview

The Cryocooler for the Mid Infrared Instrument (MIRI) on the James Webb Space Telescope (JWST) provides cooling at 6.2K on the instrument interface. The cooler system design has been incrementally documented in previous publications [1][2][3][4][5]. It has components that traverse three primary thermal regions on JWST: Region 1, approximated by 40K; Region 2, approximated by 100K; and Region 3, which is at the allowable flight temperatures for the spacecraft bus. However, there are several sub-regions that exist in the transition between primary regions and at the heat reject interfaces of the Cooler Compressor Assembly (CCA) and Cooler Control Electronics Assembly (CCEA). The design and performance of the test facility to provide a flight representative thermal environment for acceptance testing and characterization of the complete MIRI cooler subsystem are presented.

Weilert, M.↗

NASA NTRS2018

Assessment of the Mars Helicopter Thermal Design Sensitivities Using the Veritrek Software

The Mars Helicopter will be a technology demonstration conducted during the Mars 2020 mission. The primary mission objective is to achieve several 90-second flights and capture visible light images via forward and nadir mounted cameras. These flights could possibly provide reconnaissance data for sampling site selection for other Mars surface missions. The helicopter is powered by a solar array, which stores energy in secondary batteries for flight operations, imaging, communications, and survival heating. The helicopter thermal design is driven by minimizing survival heater energy while maintaining compliance with allowable flight temperatures in a variable thermal environment. Due to the small size of the helicopter and its complex geometries, along with the fact that it operates with very low power and small margins, additional care had to be paid while planning thermal tests and designing the thermal system. A Thermal Desktop® model has been developed to predict the thermal system’s performance. A reduced-order model (ROM) created with the Veritrek software has been utilized to explore the sensitivities of the thermal system’s drivers, such as electronics dissipations, gas gaps, heat transfer coefficients, etc., as well as to assess and verify the final thermal design. This paper presents the performance of the Veritrek software products and the details of the ROM creation process. The results produced by Veritrek were utilized to study the effect of the major thermal design drivers and Mars environment on the Mars Helicopter in as little as 10 days, an effort that would have taken over 4 months using traditional thermal analysis techniques.

Hengeveld, Derek W.↗

NASA NTRS2019

Thermal design and on-orbit performance of the ECOSTRESS instrument

The ECOSystem Spaceborne Thermal Radiometer Experiment on Space Station (ECOSTRESS) instrument, launched on June 29, 2018, has completed nine months of operation in space. ECOSTRESS is a multispectral thermal infrared imaging radiometer and its primary mission is to investigate and understand how climate change affects water and carbon usage on Earth. The thermal control subsystem consists of a combination of active and passive components to maintain instrument components within the allowable flight temperature limits. Its focal plane detector is cooled to 65K by a pair of mechanical cryocoolers and a third mechanical cryocooler cools an intermediate cold shield to 130K. The heat dissipation generated by cryocoolers and electronics, is removed through non-planar cold plates and tube-on plate heat exchanger, which are cooled by a circulating fluid -EF module. The heat collected by the JEM-EF fluid is exchanged with a fluid loop that circulates through radiators located on the exterior of the Space Station. This paper provides an overview of the thermal and cryogenic system design, thermal analysis results and reviews the on-orbit thermal performance.

Romero, M.↗