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  • 1
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    Aeroheating Thermal Analysis Methods for Aerobraking Mars Missions (2002)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Mars missions often employ aerobraking upon arrival at Mars as a low-mass method to gradually reduce the orbit period from a high-altitude, highly elliptical insertion orbit to the final science orbit. Two recent missions that made use of aerobraking were Mars Global Surveyor (MGS) and Mars Odyssey. Both spacecraft had solar arrays as the main aerobraking surface area. Aerobraking produces a high heat load on the solar arrays, which have a large surface area exposed to the airflow and relatively low mass. To accurately model the complex behavior during aerobraking, the thermal analysis must be tightly coupled to the flight mechanics, aerodynamics, and atmospheric modeling efforts being performed during operations. To properly represent the temperatures prior to and during the drag pass, the model must include the orbital solar and planetary heat fluxes. The correlation of the thermal model to flight data allows a validation of the modeling process, as well as information on what processes dominate the thermal behavior. This paper describes the thermal modeling method that was developed for this purpose, as well as correlation for two flight missions, and a discussion of improvements to the methodology.
    Keywords: Spacecraft Design, Testing and Performance
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  • 2
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    Thermal Analysis and Correlation of the Mars Odyssey Spacecraft's Solar Array During Aerobraking Operations (2002)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Mars Odyssey spacecraft made use of multipass aerobraking to gradually reduce its orbit period from a highly elliptical insertion orbit to its final science orbit. Aerobraking operations provided an opportunity to apply advanced thermal analysis techniques to predict the temperature of the spacecraft's solar array for each drag pass. Odyssey telemetry data was used to correlate the thermal model. The thermal analysis was tightly coupled to the flight mechanics, aerodynamics, and atmospheric modeling efforts being performed during operations. Specifically, the thermal analysis predictions required a calculation of the spacecraft's velocity relative to the atmosphere, a prediction of the atmospheric density, and a prediction of the heat transfer coefficients due to aerodynamic heating. Temperature correlations were performed by comparing predicted temperatures of the thermocouples to the actual thermocouple readings from the spacecraft. Time histories of the spacecraft relative velocity, atmospheric density, and heat transfer coefficients, calculated using flight accelerometer and quaternion data, were used to calculate the aerodynamic heating. During aerobraking operations, the correlations were used to continually update the thermal model, thus increasing confidence in the predictions. This paper describes the thermal analysis that was performed and presents the correlations to the flight data.
    Keywords: Spacecraft Instrumentation and Astrionics
    Type: AIAA Paper 2002-4536 , AIAA/AAS Astrodynamics Conference; Aug 05, 2002 - Aug 08, 2002; Monterey, CA; United States
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  • 3
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    Thermal Analysis Methods For Earth Entry Vehicle (2000)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Thermal analysis of a vehicle designed to return samples from another planet, such as the Earth Entry vehicle for the Mars Sample Return mission, presents several unique challenges. The Earth Entry Vehicle (EEV) must contain Martian material samples after they have been collected and protect them from the high heating rates of entry into the Earth's atmosphere. This requirement necessitates inclusion of detailed thermal analysis early in the design of the vehicle. This paper will describe the challenges and solutions for a preliminary thermal analysis of an Earth Entry Vehicle. The aeroheating on the vehicle during entry would be the main driver for the thermal behavior, and is a complex function of time, spatial position on the vehicle, vehicle temperature, and trajectory parameters. Thus, the thermal analysis must be closely tied to the aeroheating analysis in order to make accurate predictions. Also, the thermal analysis must account for the material response of the ablative thermal protection system (TPS). For the exo-atmospheric portion of the mission, the thermal analysis must include the orbital radiation fluxes on the surfaces. The thermal behavior must also be used to predict the structural response of the vehicle (the thermal stress and strains) and whether they remain within the capability of the materials. Thus, the thermal analysis requires ties to the three-dimensional geometry, the aeroheating analysis, the material response analysis, the orbital analysis, and the structural analysis. The goal of this paper is to describe to what degree that has been achieved.
    Keywords: Spacecraft Design, Testing and Performance
    Type: Thermal and Fluids Analysis; Aug 21, 2000 - Aug 25, 2000; Cleveland, OH; United States
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  • 4
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    Thermal Analysis Methods for an Earth Entry Vehicle (2000)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Thermal analysis of a vehicle designed to return samples from another planet, such as the Earth Entry vehicle for the Mars Sample Return mission, presents several unique challenges. The Earth Entry Vehicle (EEV) must contain Martian material samples after they have been collected and protect them from the high heating rates of entry into the Earth's atmosphere. This requirement necessitates inclusion of detailed thermal analysis early in the design of the vehicle. This paper will describe the challenges and solutions for a preliminary thermal analysis of an Earth Entry Vehicle. The aeroheatina on the vehicle during entry would be the main driver for the thermal behavior. and is a complex function of time, spatial position on the vehicle, vehicle temperature, and trajectory parameters. Thus. the thermal analysis must be closely tied to the aeroheating analysis in order to make accurate predictions. Also, the thermal analysis must account for the material response of the ablative thermal protection system TPS. For the exo-atmospheric portion of the mission, the thermal analysis must include the orbital radiation fluxes on the surfaces. The thermal behavior must also be used to predict the structural response of the vehicle (the thermal stress and strains) and whether they remain within the capability of the materials. Thus, the thermal analysis requires ties to the three-dimensional geometry, the aeroheating analysis, the material response analysis, the orbital analysis. and the structural analysis. The goal of this paper is to describe to what degree that has been achieved.
    Keywords: Spacecraft Design, Testing and Performance
    Type: Thermal and Fluids Analysis; Aug 21, 2000 - Aug 25, 2000; Cleveland, OH; United States
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  • 5
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    Preliminary Thermal Analysis of a Mars Sample Return Earth Entry Vehicle (2000)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Thermal analysis of a vehicle designed to return samples from another planet, such as the Earth Entry vehicle for the Mars Sample Return mission, presents several unique challenges. The scientific purpose of a sample return mission is to return samples to Earth for detailed investigation. The Earth Entry Vehicle (EEV) must contain the samples after they have been collected and protect them from the high heating rates of entry into the Earth's atmosphere. This requirement necessitates inclusion of detailed thermal analysis early in the design of the vehicle. This paper will describe the challenges and solutions for a preliminary thermal analysis of an Earth Entry Vehicle. The primary challenges included accurate updates of model geometry, applying heat fluxes that change with position and time during exo-atmospheric cruise and entry, and incorporating orthotropic material properties. Many different scenarios were evaluated for the exoatmospheric cruise to attain the desired thermal condition. The severity of the heat pulse during entry and the material response led to some unique modeling solutions. Overall, advanced modeling techniques and mathematical solutions were successfully used in predicting the thermal behavior of this complex system.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA Paper 2000-2584 , Thermophysics; Jun 19, 2000 - Jun 22, 2000; Denver, CO; United States
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  • 6
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    Three-Dimensional Finite Element Ablative Thermal Response and Thermostructural Design of Thermal Protection Systems (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A finite element ablation and thermal response program is presented for simulation of three-dimensional transient thermostructural analysis. The three-dimensional governing differential equations and finite element formulation are summarized. A novel probabilistic design methodology for thermal protection systems is presented. The design methodology is an eight step process beginning with a parameter sensitivity study and is followed by a deterministic analysis whereby an optimum design can determined. The design process concludes with a Monte Carlo simulation where the probabilities of exceeding design specifications are estimated. The design methodology is demonstrated by applying the methodology to the carbon phenolic compression pads of the Crew Exploration Vehicle. The maximum allowed values of bondline temperature and tensile stress are used as the design specifications in this study.
    Keywords: Aerodynamics
    Type: AIAA Paper 2011-3319 , NF1676L-11612 , 42nd AIAA Thermophysics Conference; Jun 27, 2011 - Jun 30, 2011; Honolulu, HI; United States
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  • 7
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    Two-Dimensional Finite Element Ablative Thermal Response Analysis of an Arcjet Stagnation Test (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The finite element ablation and thermal response (FEAtR, hence forth called FEAR) design and analysis program simulates the one, two, or three-dimensional ablation, internal heat conduction, thermal decomposition, and pyrolysis gas flow of thermal protection system materials. As part of a code validation study, two-dimensional axisymmetric results from FEAR are compared to thermal response data obtained from an arc-jet stagnation test in this paper. The results from FEAR are also compared to the two-dimensional axisymmetric computations from the two-dimensional implicit thermal response and ablation program under the same arcjet conditions. The ablating material being used in this arcjet test is phenolic impregnated carbon ablator with an LI-2200 insulator as backup material. The test is performed at the NASA, Ames Research Center Interaction Heating Facility. Spatially distributed computational fluid dynamics solutions for the flow field around the test article are used for the surface boundary conditions.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA Paper 2011-3617 , NF1676L-11639 , 42nd AIAA Thermophysics Conference; Jun 27, 2011 - Jun 30, 2011; Honolulu, HI; United States
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  • 8
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    Autonomous Aerobraking: Thermal Analysis and Response Surface Development (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A high-fidelity thermal model of the Mars Reconnaissance Orbiter was developed for use in an autonomous aerobraking simulation study. Response surface equations were derived from the high-fidelity thermal model and integrated into the autonomous aerobraking simulation software. The high-fidelity thermal model was developed using the Thermal Desktop software and used in all phases of the analysis. The use of Thermal Desktop exclusively, represented a change from previously developed aerobraking thermal analysis methodologies. Comparisons were made between the Thermal Desktop solutions and those developed for the previous aerobraking thermal analyses performed on the Mars Reconnaissance Orbiter during aerobraking operations. A variable sensitivity screening study was performed to reduce the number of variables carried in the response surface equations. Thermal analysis and response surface equation development were performed for autonomous aerobraking missions at Mars and Venus.
    Keywords: Spacecraft Design, Testing and Performance
    Type: NF1676L-12504 , 2011 AAS/AIAA Astrodynamics Specialist Conference; Jul 31, 2011 - Aug 04, 2011; Girdwood, AK; United States
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  • 9
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    Flexible Thermal Protection System Development for Hypersonic Inflatable Aerodynamic Decelerators (2012)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Hypersonic Inflatable Aerodynamic Decelerators (HIAD) project has invested in development of multiple thermal protection system (TPS) candidates to be used in inflatable, high downmass, technology flight projects. Flexible TPS is one element of the HIAD project which is tasked with the research and development of the technology ranging from direct ground tests, modelling and simulation, characterization of TPS systems, manufacturing and handling, and standards and policy definition. The intent of flexible TPS is to enable large deployable aeroshell technologies, which increase the drag performance while significantly reducing the ballistic coefficient of high-mass entry vehicles. A HIAD requires a flexible TPS capable of surviving aerothermal loads, and durable enough to survive the rigors of construction, handling, high density packing, long duration exposure to extrinsic, in-situ environments, and deployment. This paper provides a comprehensive overview of key work being performed within the Flexible TPS element of the HIAD project. Included in this paper is an overview of, and results from, each Flexible TPS research and development activity, which includes ground testing, physics-based thermal modelling, age testing, margins policy, catalysis, materials characterization, and recent developments with new TPS materials.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NF1676L-14329 , International Planetary Probe Workshop 9; Jun 18, 2012 - Jun 22, 2012; Toulouse; France
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  • 10
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    A Thermal Analysis Approach for the Mars Odyssey Spacecraft's Solar Array (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: 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.
    Keywords: Spacecraft Design, Testing and Performance
    Type: AIAA Paper 2003-3764 , 36th AIAA Thermophysics Conference; Jun 23, 2003 - Jun 26, 2003; Orlando, FL; United States
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