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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 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Aerothermal analysis of a wing-elevon cove with variable leakage

A mathematical model of the heating associated with leakage within the space shuttle orbiter wing elevon cove was used to analyze the aerothermal response of the structure. The analysis was applied to wind tunnel results and to shuttle entry conditions. A parametric study described the mechanism by which the energy of the ingested mass is distributed as a function of the Reynolds number, wall conductivity, surface emissivity, and cove height and length. The rate at which energy was transferred into the cove interior was primarily determined by the combined effects of the convection, wall capacitance, and internal radiation terms. Correlation with wind tunnel results indicated that the cold wall convective heating rates, wall temperature, and gas temperature of the cove were predicted by the analytical model for short exposure times. Predicted transient thermal response of the elevon cove subjected to shuttle entry conditions indicated that the allowable leak area at the cove seal was 20 percent less than that previously indicated by experiment.

Hunt, R. L.↗
NASA NTRS2018

Significance of DSMC Computed Aerothermal Environments in the Rarefied Regime for Atmospheric Entry Material Response

During Mars atmospheric entry, the Mars Science Laboratory (MSL) was protected by a 4.5 meters diameter ablative heatshield assembled in 113 tiles. The heatshield was made of NASA's flagship ablative material, the Phenolic Impregnated Carbon Ablator (PICA). Prior work compared the traditional one-dimensional and three-dimensional material response models at different locations in the heatshield. It was observed that the flow was basically one-dimensional in the nose and flank regions, but three-dimensional flow effects were observed in the outer flank. The objective of this work is to study the effects of the aerothermal environment on the material response. We extend prior work by computing aerothermal environments using the direct simulation Monte Carlo (DSMC) code SPARTA and the CFD code Data Parallel Line Relaxation (DPLR). SPARTA is used to compute environment in the rarefied regime prior to 48.4s of entry where the Knudsen number is such that the Navier-Stokes equations can be inaccurate. Similarly to previous work, the DPLR software is used to compute the hypersonic environment for laminar then turbulent boundary layer assumptions from 48.4 s up to 100 s after Entry Interface (EI) along the MSL 08-TPS-02/01a trajectory. We observe that extending the aerothermal environments to times prior to 48.4 s modifies the thermal response of the heat shield at the surface and in-depth; however the effects on the recession are minimal. Additionally, using the assumption of a turbulent boundary layer versus a laminar one leads to higher surface and in-depth temperatures, larger recession, and a displacement of the peak heating and peak recession location.

Aerothermal↗
NASA NTRS2011

Thermal Protection System Aerothermal Screening Tests in HYMETS Facility

The Entry, Descent, and Landing (EDL) Technology Development Project has been tasked to develop Thermal Protection System (TPS) materials for insertion into future Mars Entry Systems. A screening arc jet test of seven rigid ablative TPS material candidates was performed in the Hypersonic Materials Environmental Test System (HYMETS) facility at NASA Langley Research Center, in both an air and carbon dioxide test environment. Recession, mass loss, surface temperature, and backface thermal response were measured for each test specimen. All material candidates survived the Mars aerocapture relevant heating condition, and some materials showed a clear increase in recession rate in the carbon dioxide test environment. These test results supported subsequent down-selection of the most promising material candidates for further development.

thermal protection systems (TPS)↗
NASA NTRS2000

Lunar Dust Contamination Effects on Lunar Base Thermal Control Systems

Many studies have been conducted to develop a thermal control system that can operate under the extreme thermal environments found on the lunar surface. While these proposed heat rejection systems use different methods to reject heat, each system contains a similar component, a thermal radiator system. These studies have always considered pristine thermal control system components and have overlooked the possible deleterious effects of lunar dust contamination. Since lunar dust has a high emissivity and absorptivity (greater than 0.9) and is opaque, dust accumulation on a surface should radically alter its optical properties and therefore alter its thermal response compared to ideal conditions. In addition, the non-specular nature of the dust particles will alter the performance of systems that employ specular surfaces to enhance heat rejection. To date, few studies have examined the effect of dust deposition on the normal control system components. These studies only focused on a single heat rejection or photovoltaic system. These studies did show that lunar dust accumulations alter the optical properties of any lunar base hardware, which in turn affects component temperatures, and heat rejection. Therefore, a new study was conducted to determine the effect of lunar dust contamination on heat rejection systems. For this study, a previously developed dust deposition model was incorporated into the Thermal Synthesizer System (TSS) model. This modeling scheme incorporates the original method of predicting dust accumulation due to vehicle landings by assuming that the thin dust layer can be treated as a semitransparent surface slightly above and in thermal contact with the pristine surface. The results of this study showed that even small amounts of dust deposits can radically alter the performance of the heat rejection systems. Furthermore. this study indicates that heat rejection systems be either located far from any landing sites or be protected from dust producing mechanisms.

Keller, John R.↗
NASA NTRS1995

Low-Noise Wide Bandwith, Hot Electron Bolometer Mixers for Submillimeter Wavelengths

Recently a novel superconductive hot-electron micro-bolometer has been proposed which is both fast and sensitive (D. E. Prober, Appl. Phys. Lett. 62, 2119, 1993). This device has several important properties which make it useful as a heterodyne sensor for radioastronomy applications at frequencies above 1 THz. The thermal response time of the device is fast enough, several 10's of picoseconds, to allow for IF's of several GHz. This bolometer mixer should operate well up to at least 10 THz. There is no energy gap limitation as in an SIS mixer, since the mixing process relies on heating of the electron gas. In fact, rf power is absorbed more uniformly above the gap frequency. The mixer noise should be near quantum-limited, and the local oscillator (LO) power requirement is very low: / 10 nW for a Nb device. One of the unique features of this device is that it employs rapid electron diffusion into a normal metal, rather than phonon emission, as the thermal conductance that cools the heated electrons. In order for diffusion to dominate over phonon emission, the device must be short, less than 0.5.

Bolometer Mixers↗
NASA NTRSDate not supplied

Analysis of PICA-NuSil at the Hypersonic Materials Environmental Test System (HyMETS)

Phenolic Impregnated Carbon Ablator (PICA) gained heritage during the stardust mission and is the baseline material for missions to Mars. The phenolic phase of PICA is friable, and therefore the surface is treated with a siloxane resin (NuSil) to mitigate the shedding of particulate matter. Historically, material response models have not accounted for the presence of this coating. Therefore, a test campaign was executed at the Hypersonic Materials Environmental Test System (HyMETS) to investigate the performance of NuSil under multiple heating rates and gas composition. Sphere-cone models were instrumented with thermocouples to measure the in-depth temperature response, and surface temperatures were monitored with a pyrometer. Furthermore, emission spectrometers were used to collect spectra from species emitting at the surface of the test article and the post-shock stagnation region. Post-test analysis suggests that the coating decomposes according to multiple stages. Inspection of silicon emission data reveals a rapid rise and decline in intensity within the first couple of seconds of testing, which is attributed to pyrolysis of NuSil. Pyrolysis is followed by a stagnation in silicon emission and surface temperature which are attributed to the formation of a thin layer of oxidation-resistant silicon oxycarbide. Moreover, the presence of the silicon oxycarbide layer significantly reduces the in-depth thermal response when compared to the baseline material. Eventually, the silicon oxycarbide layer decomposes via high-temperature reactions, which expose the underlying char layer to a highly reactive flow.

PICA-NuSil↗
NASA NTRS2016

System Level Aerothermal Testing for the Adaptive Deployable Entry and Placement Technology (ADEPT)

The Adaptive Deployable Entry and Placement Technology (ADEPT), a mechanically deployable entry vehicle technology, has been under development at NASA since 2011. As part of the technical maturation of ADEPT, designs capable of delivering small payloads (10 kg) are being considered to rapidly mature sub 1 m deployed diameter designs. The unique capability of ADEPT for small payloads comes from its ability to stow within a slender volume and deploy to achieve a mass efficient drag surface with a high heat rate capability. The low ballistic coefficient results in entry heating and mechanical loads that can be met by a revolutionary three-dimensionally woven carbon fabric supported by a deployable skeleton structure. This carbon fabric has test proven capability as both primary structure and payload thermal protection system. In order to rapidly advance ADEPTs technical maturation, the project is developing test methods that enable thermostructural design requirement verification of ADEPT designs at the system level using ground test facilities. Results from these tests are also relevant to larger class missions and help us define areas of focused component level testing in order to mature material and thermal response design codes. The ability to ground test sub 1 m diameter ADEPT configurations at or near full-scale provides significant value to the rapid maturation of this class of deployable entry vehicles. This paper will summarize arc jet test results, highlight design challenges, provide a summary of lessons learned and discuss future test approaches based upon this methodology.

Hypersonic↗
NASA NTRS1973

Analysis of sublimation-cooled coated mirrors in convective and radiative environments.

Analytical solutions were obtained for the thermal response of a transpiration- or sublimation-cooled spherical mirror coating exposed to convective and radiative heating. The solutions allow unlimited spectral detail to be accounted for. Results indicate that transpiration-cooled thick coatings (1 cm) may withstand up to 10 kW/sq cm on a steady basis without excessive temperature rise for quartzlike materials with an internal absorption coefficient of 0.01 per cm. On a transient basis, fluxes up to 20 kW/sq cm can be accommodated for a second (cW laser exposure time), 4 kW/sq cm for 5 sec (planetary entry heating time), and of the order of MW/sq cm for millisecond times (short-duration laser bursts) without transpiration cooling for a material with an absorption coefficient of 0.1 per cm. Proportionately higher fluxes can be accommodated with lower absorption coefficients. Thermal stresses produced by the heat pulse are found to be high but within the strength of the materials. The regime in which meaningful solutions may be obtained is mapped in detail.

Howe, J. T.↗
NASA NTRS2016

Investigation of Performance Envelope for Phenolic Impregnated Carbon Ablator (PICA)

The present work provides the results of a short exploratory study on the performance of Phenolic Impregnated Carbon Ablator, or PICA, at high heat flux and pressure in an arcjet facility at NASA Ames Research Center. The primary objective of the study was to explore the thermal response of PICA at cold-wall heat fluxes well in excess of 1500 W/cm (exp 2). Based on the results of a series of flow simulations, multiple PICA samples were tested at an estimated cold wall heat flux and stagnation pressure of 1800 W/cm (exp 2) and 130 kPa, respectively. All samples survived the test, and no failure was observed either during or after the exposure. The results indicate that PICA has a potential to perform well at environments with significantly higher heat flux and pressure than it has currently been flown.

Ablative Material↗
NASA NTRS2016

Heatshield for Extreme Entry Environment Technology (HEEET) Development Status

The Heat shield for Extreme Entry Environment Technology (HEEET) Project is a NASA STMD and SMD co-funded effort. The goal is to develop and mission infuse a new ablative Thermal Protection System that can withstand extreme entry. It is targeted to support NASAs high priority missions, as defined in the latest decadal survey, to destinations such as Venus and Saturn in-situ robotic science missions. Entry into these planetary atmospheres results in extreme heating. The entry peak heat-flux and associated pressure are estimated to be between one and two orders of magnitude higher than those experienced by Mars Science Laboratory or Lunar return missions. In the recent New Frontiers community announcement NASA has indicated that it is considering providing an increase to the PI managed mission cost (PIMMC) for investigations utilizing the Heat Shield for Extreme Entry Environment Technology (HEEET) and in addition, NASA is considering limiting the risk assessment to only their accommodation on the spacecraft and the mission environment. The HEEET ablative TPS utilizes 3D weaving technology to manufacture a dual layer material architecture. The 3-D weaving allows for flat panels to be woven. The dual layer consists of a top layer designed to withstand the extreme external environment while the inner or insulating layer by design, is designed to achieve low thermal conductivity, and it keeps the heat from conducting towards the structure underneath. Both arc jet testing combined with material properties have been used to develop thermal response models that allows for comparison of performance with heritage carbon phenolic. A 50 mass efficiency is achieved by the dual layer construct compared to carbon phenolic for a broad range of missions both to Saturn and Venus. The 3-D woven flat preforms are molded to achieve the shape as they are compliant and then resin infusion with curing forms a rigid panels. These panels are then bonded on to the aeroshell structure. Gaps exist between the panels and these gaps have to be filled with seams. The seam material then has to be bonded on to adjacent panels and also to the structure. The heat-shield assembly is shown in Figure 1. One of the significant challenges we have overcome recently is the design, development and testing of the seam. HEEET material development and the seam concept development have utilized some of the unique test capabilities available in the US. The various test facilities utilized in thermal testing along with the entry environment for Saturn and Venus missions are shown in Figure 2. The HEEET project is currently in its 3rd year of a four-year development. Figure 3 illustrates the key accomplishments to date and the challenges yet to be overcome before the technology is ready for mission infusion. This proposed presentation will cover both progress that has been made in the HEEET project and also the challenges to be overcome that is highlighted in Figure 3. Objective of the HEEET project is to mature the system in time to support the next New Frontiers opportunity and we believe we are well along the way to mission infuse HEEET.

Thermal Protection System↗
NASA NTRS2016

Heat-shield for Extreme Entry Environment Technology (HEEET) Development Status

The Heat shield for Extreme Entry Environment Technology (HEEET) Project is a NASA STMD and SMD co-funded effort. The goal is to develop and mission infuse a new ablative Thermal Protection System that can withstand extreme entry. It is targeted to support NASA's high priority missions, as defined in the latest decadal survey, to destinations such as Venus and Saturn in-situ robotic science missions. Entry into these planetary atmospheres results in extreme heating. The entry peak heat-flux and associated pressure are estimated to be between one and two orders of magnitude higher than those experienced by Mars Science Laboratory or Lunar return missions. In the recent New Frontiers community announcement NASA has indicated that it is considering providing an increase to the PI managed mission cost (PIMMC) for investigations utilizing the Heat Shield for Extreme Entry Environment Technology (HEEET) and in addition, NASA is considering limiting the risk assessment to only their accommodation on the spacecraft and the mission environment. The HEEET ablative TPS utilizes 3D weaving technology to manufacture a dual layer material architecture. The 3-D weaving allows for flat panels to be woven. The dual layer consists of a top layer designed to withstand the extreme external environment while the inner or insulating layer by design, is designed to achieve low thermal conductivity, and it keeps the heat from conducting towards the structure underneath. Both arc jet testing combined with material properties have been used to develop thermal response models that allows for comparison of performance with heritage carbon phenolic. A 50% mass efficiency is achieved by the dual layer construct compared to carbon phenolic for a broad range of missions both to Saturn and Venus. The 3-D woven flat preforms are molded to achieve the shape as they are compliant and then resin infusion with curing forms a rigid panels. These panels are then bonded on to the aeroshell structure. Gaps exist between the panels and these gaps have to be filled with seams. The seam material then has to be bonded on to adjacent panels and also to the structure. The heat-shield assembly is shown in Figure 1. One of the significant challenges we have overcome recently is the design, development and testing of the seam. HEEET material development and the seam concept development have utilized some of the unique test capabilities available in the US. The various test facilities utilized in thermal testing along with the entry environment for Saturn and Venus missions are shown in Figure 2. The HEEET project is currently in it's 3rd year of a four-year development. Figure 3 illustrates the key accomplishments to-date and the challenges yet to be overcome before the technology is ready for mission infusion. This proposed presentation will cover both progress that has been made in the HEEET project and also the challenges to be overcome that is highlighted in Figure 3. Objective of the HEEET project is to mature the system in time to support the next New Frontiers opportunity and we believe we are well along the way to mission infuse HEEET.

TPS↗
NASA NTRS1972

An evaluation of flight data for the Apollo thermal protection system

A study was conducted to correlate Apollo ablation and thermal response flight data using advanced state-of-the-art analytical procedures. The agreement between flight data and predictions is consistently excellent for in-depth temperature distributions, char density profiles, and surface ablation, thus validating the analytical procedures.

Bartlett, E. P.↗
NASA NTRS1971

A note on laminar shear flow over impulsively started bodies.

Prediction of the shear flow around bodies impulsively set into motion at a uniform velocity. Information is presented on the local wall shear stress, velocity distribution, steady flow times, and thermal response for wedge flows where local flow acceleration occurs. The essential features of the flow field are found to be describable by the approximate series expansion method of Goldstein and Rosenhead (1936). This method would appear to be useful in rapidly calculating the viscous drag on the forward face of various shaped bodies where local flow acceleration occurs.

Back, L. H.↗
NASA NTRS1972

Transient, compressible heat and mass transfer in porous media using the strongly implicit iteration procedure.

Coupled nonlinear partial differential equations describing heat and mass transfer in a porous matrix are solved in finite difference form with the aid of a new iterative technique (the strongly implicit procedure). Example numerical results demonstrate the characteristics of heat and mass transport in a porous matrix such as a charring ablator. It is emphasized that multidimensional flow must be considered when predicting the thermal response of a porous material subjected to nonuniform boundary conditions.

Curry, D. M.↗
NASA NTRS1971

Computer simulation of the thermal environment of large-scale integrated circuits - Computer time-saving techniques.

This paper is concerned with the computer costs for both the steady-state and transient thermal responses of large-scale integrated circuits (LSI) when metal is present within the substrate. For the more cost-sensitive transient case, an extrapolation technique for computer time savings is compared with the accuracy loss in this study. This approach could be useful for design-cost planning.

Thompson, R. R.↗