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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 NTRS1996

PICA Forebody Heatshield Qualification for the Stardust Discovery Class Mission

This paper presents the qualification of the light weight Phenolic Impregnated Carbon Ablators (PICA) as the forebody heatshield for the Stardust Discovery Class Mission. The Stardust spacecraft will be launched in early 1999 and fly by Comet Wild-2 to collect cometary and interstellar dust and return them back to earth in the Sample Return Capsule (SRC). This earth re-entry will be the fastest to date, at 12.6 km/s, and therefore requires a heatshield that can withstand very high heating rates and stagnation pressures, as well as provide the necessary insulation to the vehicle structure. The PICA material was developed as part of the Lightweight Ceramic Ablators program at NASA Ames Research Center, and was baselined as the forebody heatshield because of its low density and superior ablation and thermal performance at severe aerothermodynamic conditions. Under a Small Business Innovative Research (SBIR) program with NASA Ames, Fiber Materials, Inc. developed a process to manufacture a single-piece PICA heatshield for the forebody of the SRC, along with witness material for the fabrication of the test models. The test models were fabricated and instrumented by the staff of Lockheed Martin Astronautics in Denver, Colorado. Full body preliminary aerothermal CFD calculations were performed at NASA Ames to determine the heating and stagnation pressure conditions. The Heat shield sizing was also performed at NASA Ames by using a new material response code that accounts for the highly porous characteristics of the PICA material. The ablation and thermal performance of PICA was qualified in the NASA Ames Interaction Heating Arc Jet Facility. A total of 24 models and four test conditions were used to qualify PICA at the predicted peak heat flux, heat load, shear, and stagnation pressure conditions. Surface and in-depth temperatures were measured using optical pyrometers and thermocouples. Surface recession was measured by using a template and a height gage. Several models were tested to evaluate repair procedures, and two models were cold soaked in liquid nitrogen, prior th testing, to investigate the effect of the cold space environment on the performance of the material. In addition, material cored from a demonstration single-piece heatshield was tested to verify that the PICA process can be successfully completed on a large, complex heatshield shape.

Tran, Huy K.↗
NASA NTRSDate not supplied

MEDLI2 Material Response Model Development and Validation

During the entry of Mars Science Laboratory (MSL) the heatshield was equipped with the instrumentation suite Mars Entry, Descent, and Landing Instruments (MEDLI). In-depth thermocouple (TC) data was used to reconstruct the surface heating and temperatures. Discrepancies between MEDLI’s recorded in-depth heating data and the predicted thermal response and recession led to a campaign to characterize and test Mars Entry, Descent, and Landing Instrument 2 (MEDLI2) flight lot thermal protection materials (TPS) at relevant temperatures and pressures for Martian entry conditions. This paper covers the development and validation of MEDLI2 flight lot material response models for the heatshield material Phenolic Impregnated Carbon Ablator (PICA) and the backshell material Super-Lightweight Ablator (SLA-561V). Virgin and char thermal conductivities were updated as a function of temperature and pressure in the MEDLI2 specific material response models. Other material properties such as virgin and char density, emissivity, absorptivity, and specific heat capacitance was also characterized and compared to the “Heritage” models used during the TPS design phase of Mars 2020. Fully Implicit Ablation and Thermal response (FIAT) simulations were completed using the Heritage and MEDLI2 material response models to provide evidence of the increased accuracy of the MEDLI2 model. FIAT predicted in-depth temperatures were compared to flight lot certification ground-test arc jet test data and MEDLI2 flight thermocouple data.

Material response↗
NASA NTRS1973

Testing and evaluation of the LES-6 pulsed plasma thruster by means of a torsion pendulum system

Performance characteristics of the LES-6 pulsed plasma thruster over a range of input conditions were investigated by means of a torsion pendulum system. Parameters of particular interest included the impulse bit and time average thrust (and their repeatability), specific impulse, mass ablated per discharge, specific thrust, energy per unit area, efficiency, and variation of performance with ignition command rate. Intermittency of the thruster as affected by input energy and igniter resistance were also investigated. Comparative experimental data correlation with the data presented. The results of these tests indicate that the LES-6 thruster, with some identifiable design improvements, represents an attractive reaction control thruster for attitude contol applications on long-life spacecraft requiring small metered impulse bits for precise pointing control of science instruments.

Hamidian, J. P.↗
NASA NTRS2014

Development of Naphthalene PLIF for Visualizing Ablation Products From a Space Capsule Heat Shield

The Orion Multi-Purpose Crew Vehicle (MPCV) will use an ablative heat shield. To better design this heat shield and others that will undergo planetary entry, an improved understanding of the ablation process would be beneficial. Here, a technique developed at The University of Texas at Austin that uses planar laser-induced fluorescence (PLIF) of a low-temperature sublimating ablator (naphthalene) to enable visualization of the ablation products in a hypersonic flow is applied. Although high-temperature ablation is difficult and expensive to recreate in a laboratory environment, low-temperature sublimation creates a limited physics problem that can be used to explore ablation-product transport in a hypersonic flow-field. In the current work, a subscale capsule reentry vehicle model with a solid naphthalene heat shield has been tested in a Mach 5 wind tunnel. The PLIF technique provides images of the spatial distribution of sublimated naphthalene in the heat-shield boundary layer, separated shear layer, and backshell recirculation region. Visualizations of the capsule shear layer using both naphthalene PLIF and Schlieren imaging compared favorably. PLIF images have shown high concentrations of naphthalene in the capsule separated flow region, intermittent turbulent structures on the heat shield surface, and interesting details of the capsule shear layer structure. It was shown that, in general, the capsule shear layer appears to be more unsteady at lower angels of attack. The PLIF images demonstrated that during a wind tunnel run, as the model heated up, the rate of naphthalene ablation increased, since the PLIF signal increased steadily over the course of a run. Additionally, the shear layer became increasingly unsteady over the course of a wind tunnel run, likely because of increased surface roughness but also possibly because of the increased blowing. Regions with a relatively low concentration of naphthalene were also identified in the capsule backshell recirculation region and are most likely the result of cross-flow-induced vortices on the capsule afterbody.

Combs, C. S.↗
NASA NTRS1994

Light Weight Ceramic Ablators for Mars Follow-on Mission Vehicle Thermal Protection System

New Light Weight Ceramic Ablators (LCA) were produced by using ceramic and carbon fibrous substrates, impregnated with silicone and phenolic resins. The special infiltration techniques (patent pending) were developed to control the amount of organic resins in the highly porous fiber matrices so that the final densities of LCA's range from 0.22 to 0.24 g/cc. This paper presents the thermal and ablative performance of the Silicone Impregnated Reusable Ceramic Ablators (SIRCA) in simulated entry conditions for Mars-Pathfinder in the Ames 60 MW Interaction Heating Facility (I HF). Arc jet test results yielded no evidence of char erosion and mass loss at high stagnation pressures to 0.25 atm. Minimal silica melt was detected on surface char at a stagnation pressure of 0.31 atm. Four ceramic substrates were used in the production of SIRCA's to obtain the effective of boron oxide present in substrate so the thermal performance of SIRCA's. A sample of SIRCA was also exposed to the same heating condition for five cycles and no significant mass loss or recession was observed. Tensile testing established that the SIRCA tensile strength is about a factor of two higher than that of the virgin substrates. Thermogravimetric Analysis (TGA) of the char in nitrogen and air showed no evidence of free carbon in the char. Scanning Electron Microscopy of the post test sample showed that the char surface consists of a fibrous structure that was sealed with a thin layer of silicon oxide melt.

Tran, Huy K.↗
NASA NTRS2005

In Situ Resource Utilization (ISRU) Experiments for Mars Exploration

In situ resource utilization can best be described as living off the land. In our case the land is the planet Mars. ISRU is based on the idea that some fraction of the consumables, life support and propellant materials do not have to be flown from earth. Rather, they can be manufactured or extracted from resources already present on Mars. The primary resources on Mars are the atmosphere, polar caps and regolith. The atmosphere of Mars is mostly carbon dioxide as shown in the table below. The proportion of oxygen on the other hand is quite small. Still, there is quite a bit of oxygen in the Martian atmosphere, but it is unfortunately tied up with carbon. Thus, one of the goals of ISRU is the separation of breathable oxygen from the carbon dioxide. Several means of separation have been proposed. We have begun experiments on another approach for production of oxygen with carbon monoxide as a useful by product. Our work on a CO2 separator is described later in this report. Regolith melting is another means of obtaining materials. Two materials of interest are iron and silicon. Iron oxide is plentiful on Mars and is of obvious importance for structural components. Silicon is the foundation of solid state devices. Power generation on Mars may be accomplished using silicon solar cells. There is discussion of the feasibility of in situ production of solar cells. This would require a means of extracting silicon from the regolith. We have conducted several experiments concerning melting and glassification of the Mars soil simulant. Other summer faculty fellows have tried various means of processing the stimulant material. These include furnace melting, microwave melting and laser ablation. We have conducted several furnace melting experiments in both air and carbon dioxide environments. We have also carried out experiments to test spark melting in a carbon dioxide atmosphere. These experiments suggest the possibility of using arc melting in a reducing atmosphere. It is important to keep in mind that we are working with a soil stimulant. Any simulant, no matter how chemically similar it is to Martian regolith, may differ in mineralogy. The underlying assumption in this work is that once a glass is formed, any differences between simulant and regolith are unimportant. The exact means of forming the glass do, however, depend on the mineralogy of the regolith. A sample return mission is required to help answer these questions.

Marone, Matt↗
NASA NTRS2009

Validation of PICA Ablation and Thermal-Response Model at Low Heat Flux

Phenolic Impregnated Carbon Ablator (PICA) was the forebody heatshield material on the Stardust sample-return capsule and is also a primary candidate material for the Mars Science Lander (MSL), the Orion Crew Module, and the SpaceX Dragon vehicle. As part of the heatshield qualification for Orion, physical and thermal properties of virgin and charred PICA were measured, and an ablation and thermal response model was developed. We validated the model by comparing it with recession and temperature data from stagnation arcjet tests conducted over a wide range of stagnation heat flux of 107 to 1102 W/sq cm. The effect of orthotropic thermal conductivity was evident in the thermal response of the arcjet models. In general, model predictions compared well with the data; however, the uncertainty of the recession prediction was greatest for heat fluxes below 200 W/sq cm. More recent MSL testing focused on the low heat flux regime of 45 to 250 W/sq cm. The new results confirm the recession uncertainty, especially for pressures below 6 kPa. In this work we focus on improving the model predictions for MSL and Orion tests below 250 W/sq cm.

Milos, Frank S.↗
NASA NTRS1972

Torsion pendulum measurements on viscoelastic materials during vacuum exposure

A torsional pendulum apparatus designed for testing in situ in vacuum, the dynamic mechanical properties of materials is described. The application of this apparatus to an experimental program to measure the effects of vacuum on the mechanical properties of two ablator materials (a foamed material and a filled elastomer) and a solid rocket propellant (a filled elastomer) is presented. Results from the program are discussed as to the effects of vacuum on the mechanical properties of these three materials. In addition, time-temperature-superposition, as a technique for accelerating vacuum induced changes in mechanical properties, is discussed with reference to the three materials tested in the subject program.

Ward, T. C.↗
NASA NTRS1959

Space: Past, Present, and Future: for Presentation to the American Ordnance Association, Washington Post

It is obvious that the last decade has produced dramatic changes in our national viewpoints on the question of the exploration of space. I would like to refer back, as a starting point, to a series of tests which were performed in 1949 under the sponsorship of the Army Ordnance Department under the code name, "Bumper." This project was carried out jointly by the Jet Propulsion Laboratory, the General Electric Company, and the Douglas Aircraft Company. The object of the project was to launch a Wac Corporal upper-stage vehicle from a V-2, and demonstrate a prototype high-performance, multi-stage rocket vehicle system. As you may recall, the project was entirely successful in demonstrating the feasibility of several unproved concepts: the staging of rocket vehicles at high altitudes, the utility of ablative materials for protection from aerodynamic heating, and a method of stabilizing rocket propelled vehicles operating in the vacuum of space. One of the most significant interpretations of this test to the engineers familiar with the work was the fact that the test showed that there were no longer any purely scientific barriers to the construction of vehicles capable of space flight. The barrier was engineering.

Stewart, Homer J.↗
NASA NTRS2013

Woven TPS Enabling Missions Beyond Heritage Carbon Phenolic

WTPS is a new approach to producing TPS architectures that uses precisely engineered 3D weaving techniques to customize material characteristics needed to meet specific missions requirements for protecting space vehicles from the intense heating generated during atmospheric entry. Using WTPS, sustainable, scalable, mission-optimized TPS solutions can be achieved with relatively low life cycle costs compared with the high costs and long development schedules currently associated with material development and certification. WTPS leverages the mature state-of-the-art weaving technology that has evolved from the textile industry to design TPS materials with tailorable performance. Currently, missions anticipated encountering heat fluxes in the range of 1500 4000 Wcm2 and pressures greater than 1.5 atm are limited to using fully dense Carbon Phenolic. However, fully dense carbon phenolic is only mass efficient at higher heat fluxes greater than 4000 Wcm2), and current mission designs suffer this mass inefficiency for lack of an alternative mid-density TPS. WTPS not only bridges this mid-density TPS gap but also offers a replacement for carbon phenolic, which itself requires a significant and costly redevelopment effort to re-establish its capability for use in the high heat flux missions recently prioritized in the NRC Decadal survey, including probe missions to Venus, Saturn and Neptune. This presentation will overview the WTPS concept and present some results from initial testing completed comparing WTPS architectures to heritage carbon phenolic.

ablators↗
NASA NTRS2012

Performance of Conformable Phenolic Impregnated Carbon Ablator in Aerothermal Environments

Conformable Phenolic Impregnated Carbon Ablator, a cousin of Phenolic Impregnated Carbon Ablator (PICA), was developed at NASA Ames Research Center as a lightweight thermal protection system under the Fundamental Aeronautics Program. PICA is made using a brittle carbon substrate, which has a very low strain to failure. Conformable PICA is made using a flexible carbon substrate, a felt in this case. The flexible felt significantly increases the strain to failure of the ablator. PICA is limited by its thermal mechanical properties. Future NASA missions will require heatshields that are more fracture resistant than PICA and, as a result, NASA Ames is working to improve PICA's performance by developing conformable PICA to meet these needs. Research efforts include tailoring the chemistry of conformable PICA with varying amounts of additives to enhance mechanical properties and testing them in aerothermal environments. This poster shows the performance of conformable PICA variants in arc jets tests. Some mechanical and thermal properties will also be presented.

ablative materials↗
NASA NTRS2021

Pyrolysis Rate and Yield Strength Reduction Incarbon Fiber and Glass Fiber Composites Underreentry Heating Conditions

The behavior of composite materials, specifically carbon fiber reinforced plastic (CFRP) and glass fiber reinforced plastic (GFRP), under reentry conditions poses a problem for space debris reentry risk modeling. Since these materials pyrolyze rather than melt and their different components demise at different rates, modeling their destruction to determine ground impact risk is complex. Modern spacecraft are using these materials in ever-greater quantities owing to their superior strength-to-weight characteristics, and this has required that the orbital debris community improve its understanding of how these materials demise on reentry. The NASA Orbital Debris Program Office undertook an extensive test campaign to better understand the rate at which several types of GFRP and CFRP materials pyrolyze under reentry heating conditions and how that pyrolysis affects the ultimate strength of the material. GFRP with a polyester resin (G10/FR-4) and CFRP with epoxy, cyanate ester, vinyl ester, and phenolic resins were tested. The test campaign was carried out at the Inductively Coupled Plasma (ICP) Torch Facility at the University of Texas at Austin. Because the ICP facility operates in a shirt-sleeve environment, test samples can be changed within seconds or minutes, allowing many samples to be tested in a short period. Two heat flux rates, 20 W/m2 and 30W/m2, and two oxygen concentration conditions, 0% and 2% of atmospheric, were applied to all five types of material. To measure both the char rate and the effect of pyrolysis on the ultimate material strength, two types of test were carried out for each material: a char rate test on a ~10 mm thick sample of material and an in-situ bending stress test of a ~2 mm thick sample of material. Measurements of the char rate showed very similar average pyrolysis front velocity in epoxy resin CFRP as in G10 at 3.6 mm/min and 3.4 mm/min, respectively. However, the total mass loss rate in theG10 was nearly double that of the CFRP at 3.8 g/min and 2.2 g/min, respectively. This result represented a slow ablation rate of carbon fibers in the CFRP at the temperatures encountered in low Earth orbit reentry and a comparatively rapid removal of the glass fibers in G10 due to melting and spallation. Pyrolysis front velocity was more significantly affected by the type of plastic than the type of fiber, with the cyanate ester CFRP samples displaying an average pyrolysis front velocity of only 1.9 mm/min. Similarly, the effect of thermal exposure on the ultimate strength of the material depended heavily on the type of plastic and very little on the type of fiber in the material. Epoxy, vinyl ester, and polyester resins all behaved very similarly, with complete structural failure at between 400 J/g and 600 J/g of specific absorbed heat. Phenolic and cyanate ester resins, on the other hand, displayed a change in structural properties that was only barely measurable with the current apparatus even after the maximum exposure time tested. These data are being incorporated into a numerical model of the ablation and demise of composite materials that will be used to more accurately calculate the ground casualty risk of future spacecraft.

Benton R. Greene↗
NASA NTRS1973

Flexible fire retardant polyisocyanate modified neoprene foam

Lightweight, fire resistant foams have been developed through the modification of conventional neoprene-isocyanate foams by the addition of an alkyl halide polymer. Extensive tests have shown that the modified/neoprene-isocyanate foams are much superior in heat protection properties than the foams heretofore employed both for ballistic and ablative purposes.

Parker, J. A.↗
NASA NTRS1992

Conductor-backed coplanar waveguide resonators of YBa2Cu3O(7-delta) on LaAlO3

Conductor-backed coplanar waveguide (CBCPW) resonators operating at 10.8 GHz have been fabricated from laser ablated and off-axis magnetron sputtered YBa2Cu3O(7-delta) (YBCO) high-temperature superconducting (HTS) thin films on LaAlO3. These resonators were tested in the temperature range from 14 to 92 K. The unloaded quality factor at 77 K of the HTS CBCPW resonators was 3 to 4 times that of a similar gold resonator. To the authors' knowledge, these results represent the first reported measurements of HTS-based CBCPW resonators.

Miranda, Felix A.↗
NASA NTRS2009

In-Situ Measurement of Hall Thruster Erosion Using a Fiber Optic Regression Probe

One potential life-limiting mechanism in a Hall thruster is the erosion of the ceramic material comprising the discharge channel. This is especially true for missions that require long thrusting periods and can be problematic for lifetime qualification, especially when attempting to qualify a thruster by analysis rather than a test lasting the full duration of the mission. In addition to lifetime, several analytical and numerical models include electrode erosion as a mechanism contributing to enhanced transport properties. However, there is still a great deal of dispute over the importance of erosion to transport in Hall thrusters. The capability to perform an in-situ measurement of discharge channel erosion is useful in addressing both the lifetime and transport concerns. An in-situ measurement would allow for real-time data regarding the erosion rates at different operating points, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over a thruster s operating envelope would also be useful in the modeling of the detailed physics inside the discharge chamber. There are many different sensors and techniques that have been employed to quantify discharge channel erosion in Hall thrusters. Snapshots of the wear pattern can be obtained at regular shutdown intervals using laser profilometry. Many non-intrusive techniques of varying complexity and sensitivity have been employed to detect the time-varying presence of erosion products in the thruster plume. These include the use quartz crystal microbalances, emission spectroscopy, laser induced flourescence, and cavity ring-down spectroscopy. While these techniques can provide a very accurate picture of the level of eroded material in the thruster plume, it is more difficult to use them to determine the location from which the material was eroded. Furthermore, none of the methods cited provide a true in-situ measure of erosion at the channel surface while the thruster is in operation (i.e. none yield a continuous channel erosion measurement). A recent fundamental sensor development effort has led to a novel regression, erosion, and ablation sensor technology (REAST). The REAST sensor allows for measurement of real-time surface erosion rates at a discrete surface location. The sensor was tested using a linear Hall thruster geometry (see Fig. 1), which served as a means of producing plasma erosion of a ceramic discharge chamber. The mass flow rate, discharge voltage, and applied magnetic field strength could be varied, allowing for erosion measurements over a broad thruster operating envelope. Results are presented demonstrating the ability of the REAST sensor to capture not only the insulator erosion rates but also changes in these rates as a function of the discharge parameters.

Polzin, Kurt↗
NASA NTRS1972

Continuing development of the short-pulsed ablative space propulsion system.

Major advancements have been made in solid propellant pulsed plasma propulsion technology. A LES-6 microthruster has operated for 8900 hours in orbit. Millipound thrustors have been developed and tested at performance levels of as low as 112 watts/mlb. Efficiencies of up to 50% and specific impulse values up to 5000 sec have been measured. An improved microthrustor has operated at 207 watts/mlb continuously for 1440 hours. Correlations of design and performance data exist to allow thrustors to be built to meet a particular level of performance.

Palumbo, D. J.↗
NASA NTRSDate not supplied

Predictive Modeling of Carbon Ablators Using Micro and Macro-Scale Modeling

Efforts to build a Predictive Material Modeling (PMM) framework from the micro-scale to the macro-scale are presented in this abstract. To reduce the need for extensive testing, accelerate the design cycle process, and reduce uncertainty margins applied to final designs, NASA is developing simulation and modeling tools that enable characterization of material properties and response to high-enthalpy environments. The Porous Microstructure Analysis (PuMA) code has been developed for computing macroscale (volume averaged) properties of porous materials using microscale images from micro-computed tomography (micro-CT). Microscale modeling requires a realistic representation of a material microstructure; these are obtained either synthetically during the design of the material or through X-ray micro-CT. Volume averaged properties are then used to inform macroscale material response models, such as those implemented in the Porous-material Analysis Toolbox based on OpenFOAM (PATO) software, also actively developed by NASA. The computational model in PATO is a generic heat and mass transfer model for porous reactive materials containing several solid phases and a single gas phase. The detailed chemical interactions occurring between the solid phases and the gas phase are modeled at the pore scale assuming local thermal equilibrium. These tools were developed to efficiently interface with other pre-existing codes such as SPARTA (direct simulation Monte Carlo), DPLR (hypersonic CFD), NEQAIR (radiative transport) and DAKOTA (uncertainty quantification and optimization). Detailed flight data (Mars Science Laboratory [MSL] Entry Descent and Landing Instrument [MEDLI]) is critical for validating these computational tools for NASA applications. Examples of modeling ablative material response using these codes will be presented including 3D simulations of the full-scale heatshield of the MSL capsule. The simulations demonstrate the ability of the modern material response code, PATO, to handle the material response of geometrically complex and large domains, through the use of massively parallel computations.

Thermal Protection Systems↗
NASA NTRSDate not supplied

Carbon Nanotube – Polysiloxane Nanocomposites for Ablative Thermal Protection Systems

Ultra-high temperature resins (UHTR) such as polysiloxane possess excellent characteristics for ablation including low thermal conductivity, lightweight composition, high thermal resistance, and the ability to form a durable char layer. To enhance its mechanical strength, the aerospace industry utilizes fiber reinforcement, such as carbon fiber reinforced polysiloxane (CF-polysiloxane), due to its high strength-to-weight ratio. Carbon nanotubes (CNTs) have advantages over microfibers including higher specific strength, lower density, and customizable thermal conductivity. Their nano-scale dimensions allow for a more homogenous material compared to traditional microfibers. Conventional methods of incorporating CNTs into resins often have structural defects such as voids, CNT agglomerations, and randomly dispersed CNTs, resulting in marginal improvements in properties. A novel approach known as bulk nanocomposite laminating (BNL) has been introduced involving CNT densification, resin polymer infiltration, and pressurized curing to achieve a high packing density exceeding 35% by volume. Characterization and testing to simulate re-entry conditions is performed with a sample of CNT-polysiloxane nanocomposite cured onto a CF-polysiloxane backing. These samples are subjected to ablation by an oxyacetylene flame while collecting temperature data to evaluate thermal protection properties.

Palak B Patel↗