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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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At least 667 records · Page 37
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Mars Sample Retrieval Lander Thermal Protection System Design

The Mars Sample Return (MSR) Sample Retrieval Lander (SRL) was slated to launch in 2031 and enter Mars’ atmosphere after a 15-month cruise. After several iterations of vehicle design and trajectories, the latest mission architecture included a 4.72-meter diameter entry capsule entering the Martian atmosphere at 8 km/s, marking a 3.5x increase in kinetic energy when entering the atmosphere in comparison to Mars 2020. Designed to be the heaviest and fastest vehicle to enter Mars’ atmosphere, SRL required slim mass margins, and as such, the thermal protection system (TPS) design methodology needed to be tailored appropriately. The entry capsule aeroshell is composed of a heatshield, backshell, parachute support cone, and parachute lid, as shown in Fig. 1.

Hannah Alpert↗
NASA NTRS1971

Computational techniques for design optimization of thermal protection systems for the space shuttle vehicle. Volume 1: Final report

Computational techniques were developed and assimilated for the design optimization. The resulting computer program was then used to perform initial optimization and sensitivity studies on a typical thermal protection system (TPS) to demonstrate its application to the space shuttle TPS design. The program was developed in Fortran IV for the CDC 6400 but was subsequently converted to the Fortran V language to be used on the Univac 1108. The program allows for improvement and update of the performance prediction techniques. The program logic involves subroutines which handle the following basic functions: (1) a driver which calls for input, output, and communication between program and user and between the subroutines themselves; (2) thermodynamic analysis; (3) thermal stress analysis; (4) acoustic fatigue analysis; and (5) weights/cost analysis. In addition, a system total cost is predicted based on system weight and historical cost data of similar systems. Two basic types of input are provided, both of which are based on trajectory data. These are vehicle attitude (altitude, velocity, and angles of attack and sideslip), for external heat and pressure loads calculation, and heating rates and pressure loads as a function of time.

Source record↗
NASA NTRS1993

Thermal Protection System design studies for lunar crew module

The results of a study to predict aeroheating and Thermal Protection System (TPS) requirements for manned entry vehicles returning to Earth from the moon are presented. The effects of vehicle size and lunar-return strategies on the aerothermodynamic environment and TPS design were assessed. Study guidelines were based on an Apollo Command Module (CM) configuration and lunar return strategies included direct entry and aerocapture followed by Low Earth Orbit entry (LEO). Convective heating was obtained by the boundary layer integral matrix procedure (BLIMP) code, and radiative heating was computed with the QRAD program. The AESOP-STAB code and the AESOP-THERM code were used for TPS analysis for ablating materials and nonablating materials respectively. Results indicated that there was an optimum size for minimum heating and that direct entry had higher heating rates than aerocapture. Aerocapture resulted in higher heat loads and TPS weight. The TPS weight factor was 6-8 percent for all lunar return strategies, with the TPS weight being about 50 percent less than that of the original Apollo CM vehicle.

Williams, S. D.↗
NASA NTRS1973

Engineering design, stress and thermal analysis, and documentation for SATS program

An in-depth analysis and mechanical design of the solar array stowage and deployment arrangements for use in Small Applications Technology Satellite spacecraft is presented. Alternate approaches for the major elements of work are developed and evaluated. Elements include array stowage and deployment arrangements, the spacecraft and array behavior in the spacecraft despin mode, and the design of the main hinge and segment hinge assemblies. Feasibility calculations are performed and the preferred approach is identified.

Source record↗
NASA NTRS1993

Cryogenic propellant thermal control system design considerations, analyses, and concepts applied to a Mars human exploration mission

This paper analyzes, defines, and sizes cryogenic storage thermal control systems that meet the requirements of future NASA Mars human exploration missions. The design issues of this system include the projection of the existing Multilayer Insulation data base for cryogenic storage to much thicker (10 cm or more) insulation systems, the unknown heat leak from mechanical interfaces, and the thermal and structural performance effects of the large tank sizes required for a Mars mission. Acknowledging these unknown effects, heat loss projections are made based on extrapolation of the existing data base. The results indicate that hydrogen, methane, and oxygen are feasible propellants, and that the best suited thermal control sytems are 'thick' MLI, thermodynamic vent sytems, cryocoolers, and vacuum jackets.

Plachta, David W.↗
NASA NTRS1999

Thermal Analysis and Design Considerations of NASA's Passively Cooled 35K Next Generation Space Telescope (NGST)

The configuration and thermal analyses of NASA's Next Generation Space Telescope (NGST) Yardstick concept utilizing a novel sunshield approach for passive cooling is described. The NGST mission concept of a large aperture optical telescope passively cooled to less than 40 K and instrument detectors passively cooled to below 30 K is unique from any other mission flown to date. Achieving such a low operational temperature requires reducing by a factor of several thousand the internal heat dissipation and environmental heating of the telescope. The techniques for achieving these requirements, i.e. orbit selection, configuration, etc., along with the supporting thermal analyses are described.

Parrish, Keith↗
NASA NTRSDate not supplied

Design of a Thermal Testbed for Metrology of Active Antennas

NASA mission requirements have driven an increased interest in active phased array antennas (APAA) for space-based user communication terminals. Recent advancements in 5G technology have driven down the cost of APAA development and manufacturing all while providing a technology solution that covers many existing Ka-Band satellite communication spectrum bands. While these developments have provided ample opportunities to leverage new chips and arrays for use in space, there is also a need to evaluate these antennas in a relevant environment. Active arrays, as designed for use in 5G, require thermal management to avoid damage to the device as well as to maintain performance. Measured performance under various thermal conditions is essential both for understanding overall APAA performance and ensuring operation is within required tolerances. To address these measurement needs, the SmallSat Ka-band Operations User Terminal (SKOUT) project at the NASA Glenn Research Center (GRC) developed a test environment that combines traditional antenna and communication system metrology with a conduction cooling/heating thermal control system to simulate the space thermal environment. This paper will address the metrologysystem design and performance specifications as well as test article setup and operation. Tests that are typically performed at ambient temperature can now be performed over the typical temperature range of a Low Earth Orbit (LEO) mission. These tests include error vector magnitude (EVM), gain to noise temperature (G/T), antenna patterns, and non-linear characterization. The paper will cover the configuration for each test and provide results from a recent test campaign. Results illustrate the importance of higher fidelity environmental testing when evaluating the performance of an APAA.

active phased array↗
NASA NTRS1982

Starprobe thermal shield system design concepts

The mission goals, flight trajectory, and material durability requirements for the NASA Starprobe spacecraft are reviewed. The spacecraft will use a Jovian gravity assist to pass within four solar radii of the sun to study fields and particles near the sun, perform experiments dealing with relativity and gravity, and observe the structure of the solar atmosphere from the photosphere to the corona. Constraints on the system size and mass design are given, and the system is noted to be required to withstand 2500 K at perihelion, thermally insulate the instrument payload, have a tube for optical measurements, and provide protection from meteorite damage. A secondary shield is also required to dispense thermal radiation that passes the primary shield and could endanger the payload. Design options are discussed, along with temperature control requirements and a conical carbon-carbon primary shield with mass-loss rate characteristics sufficient to meet a 2.5 mg/sec criterion.

Maag, C. R.↗
NASA NTRS2010

Design and Preliminary Thermal Performance of the Mars Science Laboratory Rover Heat Exchangers

The challenging range of proposed landing sites for the Mars Science Laboratory Rover requires a rover thermal management system that is capable of keeping temperatures controlled across a wide variety of environmental conditions. On the Martian surface where temperatures can be as cold as -123 degrees Centigrade and as warm as 38 degrees Centigrade, the Rover relies upon a Mechanically Pumped Fluid Loop (MPFL) and external radiators to maintain the temperature of sensitive electronics and science instruments within a -40 degrees Centigrade to 50 degrees Centigrade range. The MPFL also manages significant waste heat generated from the Rover power source, known as the Multi Mission Radioisotope Thermoelectric Generator (MMRTG). The MMRTG produces 110 Watts of electrical power while generating waste heat equivalent to approximately 2000 Watts. Two similar Heat Exchanger (HX) assemblies were designed to both acquire the heat from the MMRTG and radiate waste heat from the onboard electronics to the surrounding Martian environment. Heat acquisition is accomplished on the interior surface of each HX while heat rejection is accomplished on the exterior surface of each HX. Since these two surfaces need to be at very different temperatures in order for the MPFL to perform efficiently, they need to be thermally isolated from one another. The HXs were therefore designed for high in-plane thermal conductivity and extremely low through-thickness thermal conductivity by using aerogel as an insulator inside composite honeycomb sandwich panels. A complex assembly of hand welded and uniquely bent aluminum tubes are bonded onto the HX panels and were specifically designed to be easily mated and demated to the rest of the Rover Heat Recovery and Rejection System (RHRS) in order to ease the integration effort. During the cruise phase to Mars, the HX assemblies serve the additional function of transferring heat from the Rover MPFL to the separate Cruise Stage MPFL so that heat generated deep inside the Rover can be dissipated via the Cruise Stage radiators. Significant fabrication challenges had to be overcome in order to make the HX design a reality. The cruise phase thermal performance of the Rover HXs was verified in the cruise phase system level thermal vacuum test that was performed at JPL in January of 2009. The Rover HXs were modeled in I-DEAS TMG and predictions are compared to actual data from the test.

Heat Rejection System↗
NASA NTRSDate not supplied

Pterodactyl: Thermal Protection System Design Methodology for a Flap Control System

As interest in non-traditional entry vehicles continues to grow, the need for a Thermal Protection System (TPS) analysis approach that can account for entry solutions with changing geometry and complicated flow dynamics becomes invaluable. The NASA Space Technology Mission Directorate (STMD) Pterodactyl project aims to accomplish this through examination of a flap controlled Adaptable, Deployable Entry Placement Technology (ADEPT)-style Deployable Entry Vehicle (DEV). This paper details an improved methodology for modeling the aerothermodynamic environment and initial TPS design for a flap control system integrated with a symmetric DEV. Improvements include i) the addition of an anchoring process that uses an increased fidelity aerodynamic solution to anchor the aerothermal environment predictions and ii) increased surface resolution, for the 1D heat transfer analysis, to isolate the hottest area on the flap that will require the thickest TPS. It was found that the anchoring process provided an improved aerothermal environment prediction. Additionally, this analysis demonstrated that there is a need for increased surface resolution in the 1D thermal analysis since the predicted location of the hottest point was significantly different than the location identified using very coarse surface resolution.

Zane B Hays↗
NASA NTRS1981

Design study for Thermal Infrared Multispectral Scanner (TIMS)

The feasibility of dividing the 8-12 micrometer thermal infrared wavelength region into six spectral bands by an airborne line scanner system was investigated. By combining an existing scanner design with a 6 band spectrometer, a system for the remote sensing of Earth resources was developed. The elements in the spectrometer include an off axis reflective collimator, a reflective diffraction grating, a triplet germanium imaging lens, a photoconductive mercury cadmium telluride sensor array, and the mechanical assembly to hold these parts and maintain their optical alignment across a broad temperature range. The existing scanner design was modified to accept the new spectrometer and two field filling thermal reference sources.

Stanich, C. G.↗
NASA NTRS1995

A design perspective on thermal barrier coatings

This technical paper addresses the challenges for maximizing the benefit of thermal barrier coatings for turbine engine applications. The perspective is from a customer's viewpoint, a turbine airfoil designer, who is continuously challenged to increase the turbine inlet temperature capability for new products while maintaining cooling flow levels or even reducing them. This is a fundamental requirement to achieve increased engine thrust levels. Developing advanced material systems for the turbine flowpath airfoils is one approach to solve this challenge, for example, high temperature nickel based superalloys or thermal barrier coatings to insulate the metal airfoil from the hot flowpath environment. The second approach is to increase the cooling performance of the turbine airfoil, which enables increased flowpath temperatures and reduced cooling flow levels. Thermal barrier coatings have been employed in jet engine applications for almost 30 years. The initial application was on augmenter lines to provide thermal protection during afterburner operation. However, the production use of thermal barrier coating in the turbine section has only occurred in the past 15 years. The application was limited to stationary parts, and only recently incorporated on the rotating turbine blades. This lack of endorsement of thermal barrier coatings resulted from the poor initial durability of these coatings in high heat flux environments. Significant improvements have been made to enhance spallation resistance and erosion resistance which has resulted in increased reliability of these coatings in turbine applications.

Soechting, F. O.↗
NASA NTRS1995

A design perspective on thermal barrier coatings

This technical paper addresses the challenges for maximizing the benefit of thermal barrier coatings for turbine engine applications. The perspective is from a customer's viewpoint, a turbine airfoil designer, who is continuously challenged to increase the turbine inlet temperature capability for new products while maintaining cooling flow levels or even reducing them. This is a fundamental requirement to achieve increased engine thrust levels. Developing advanced material systems for the turbine flowpath airfoils is one approach to solve this challenge; such as high temperature nickel based superalloys or thermal barrier coatings to insulate the metal airfoils from the hot flowpath environment. The second approach is to increase the cooling performance of the turbine airfoil, which enables increased flowpath temperatures and reduced cooling flow levels. Thermal barrier coatings have been employed in jet engine applications for almost 30 years. The initial application was on augmentor liners to provide thermal protection during afterburner operation. However, the production use of thermal barrier coating in the turbine section has only occurred in the past 15 years. The application was limited to stationary parts, and only recently incorporated on the rotating turbine blades. This lack of endorsement of thermal barrier coatings resulted from the poor initial durability of these coatings in high heat flux environments. Significant improvements have been made to enhance spallation resistance and erosion resistance which has resulted in increased reliability of these coatings in turbine applications.

Soechting, Friedrich O.↗
NASA NTRS2001

Thermal Analysis and Design of Multi-layer Insulation for Re-entry Aerodynamic Heating

The combined radiation/conduction heat transfer in high-temperature multi-layer insulations was modeled using a finite volume numerical model. The numerical model was validated by comparison with steady-state effective thermal conductivity measurements, and by transient thermal tests simulating re-entry aerodynamic heating conditions. A design of experiments technique was used to investigate optimum design of multi-layer insulations for re-entry aerodynamic heating. It was found that use of 2 mm foil spacing and locating the foils near the hot boundary with the top foil 2 mm away from the hot boundary resulted in the most effective insulation design. A 76.2 mm thick multi-layer insulation using 1, 4, or 16 foils resulted in 2.9, 7.2, or 22.2 percent mass per unit area savings compared to a fibrous insulation sample at the same thickness, respectively.

Daryabeigi, Kamran↗
NASA NTRS1983

The Shuttle Orbiter thermal protection system materials, designs, and flight performance overview

The design requirements for the Orbiter thermal protection system (TPS), the various TPS materials that are used, the different design approaches associated with each of the materials, and the performance experienced during the flight test program are described. The first five flights of the Orbiter Columbia have provided the necessary data to verify the TPS thermal performance, structural integrity, and reusability. The flight performance characteristics of each TPS material are discussed. This discussion is based on postflight inspections and postflight interpretation of the flight instrumentation data. The flights to date indicate that the thermal and structural design requirements for the Orbiter TPS have been met and that the overall performance has been outstanding.

Dotts, R. L.↗
NASA NTRS1973

Thermal control subsystem design of a Saturn/Uranus atmospheric entry probe for descent missions to 20 bars.

A study of scientific probes to explore the atmospheres of Saturn and Uranus to depths of between 2 and 30 bars is summarized. The study assessed configuration tradeoffs of recommended exploratory payloads, determined the types and quantities of the thermal control devices required, and evaluated thermal structural details and viable mission alternatives. These results indicated that a vented probe design with an environmental pressurization system for the first 2.5 bars of descent would satisfy the requirements for missions to either Saturn or Uranus to depths of approximately 20 bars. Parametric atmospheric descent thermal analyses are discussed and the effect of worst-case atmospheric models and navigational arrival uncertainties on mission performance are presented.

Webb, C. M.↗
NASA NTRS2012

Efficient Thermally Conductive Strap Design for Cryogenic Propellant Tank Supports and Plumbing

After evalu1ating NASA space architecture goals, the Office of Chief Technologist identified the need for developing enabling technology for long term loiters in space with cryogenic fluids. One such technology is structural heat interception. In this prototype, heat interception at the tank support strut was accomplished using a thermally conductive link to the broad area cooled shield. The design methodology for both locating the heat intercept and predicting the reduction in boil-off heat leak is discussed in detail. Results from the chosen design are presented. It was found that contact resistance resulting from different mechanical attachment techniques played a significant role in the form and functionality of a successful design.

Elchert, J. P.↗