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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 91 records · Page 5

Galileo Heat Shield Response Experiment Data Reduction and Analyses

The Galileo heat shield response experiment consists of 10 Analog Resistance Ablation Detectors (ARADs) embedded in the carbon phenolic heat shield. As the vehicle descends into the atmosphere of Jupiter, these ARADs will in effect measure the recession as a function of time at various locations on the heat shield. This recession data will be used to reconstruct the time-dependent shape of the vehicle. The shape data is of critical importance to the Atmospheric Reconstruction Experiment which must know mass and drag coefficient of the vehicle in order to determine the structure of the Jovian atmosphere. The data is also intrinsically useful as a database for evaluating the accuracy of our numerical codes and methodology for predicting highly radiating reentry flowfields and coupled heat shield material response. This paper will document the reduction of the data from the ARAD sensors, reconstruction of the heat shield shape history, and comparison of the actual heat shield response with previous calculations and experiments. If possible, we will also include results from new calculations which couple the shock layer flow with the transient material response.

Milos, Frank S.↗

Reentry Thermal Analysis of a Generic Crew Exploration Vehicle Structure

Comparative studies were performed on the heat-shielding characteristics of honeycomb-core sandwich panels fabricated with different materials for possible use as wall panels for the proposed crew exploration vehicle. Graphite/epoxy sandwich panel was found to outperform aluminum sandwich panel under the same geometry due to superior heat-shielding qualities and lower material density. Also, representative reentry heat-transfer analysis was performed on the windward wall structures of a generic crew exploration vehicle. The Apollo low Earth orbit reentry trajectory was used to calculate the reentry heating rates. The generic crew exploration vehicle has a graphite/epoxy composite honeycomb sandwich exterior wall and an aluminum honeycomb sandwich interior wall, and is protected with the Apollo thermal protection system ablative material. In the thermal analysis computer program used, the TPS ablation effect was not yet included; however, the results from the nonablation heat-transfer analyses were used to develop a "virtual ablation" method to estimate the ablation heat loads and the thermal protection system recession thicknesses. Depending on the severity of the heating-rate time history, the virtual ablation period was found to last for 87 to 107 seconds and the ablation heat load was estimated to be in the range of 86 to 88 percent of the total heat load for the ablation time period. The thermal protection system recession thickness was estimated to be in the range of 0.08 to 0.11 inches. For the crew exploration vehicle zero-tilt and 18-degree-tilt stagnation points, thermal protection system thicknesses of h = {0.717, 0.733} inches were found to be adequate to keep the substructural composite sandwich temperature below the limit of 300 F.

Ko, William L.↗

Three-Dimensional (3D) Additive Construction: Printing with Regolith

Three dimensional (3D) printing is a new and booming topic in many realms of research and engineering technology. When it comes to space science and aerospace engineering, it can be useful in numerous ways. As humans travel deeper into space and farther from Earth, sending large quantities of needed supplies from Earth for a mission becomes astronomically expensive and less plausible. In order to reach further to new places, In Situ Resource Utilization (ISRU), a project that pushes for technologies to use materials already present in the destination's environment, is necessary. By using materials already available in space such as regolith from the Moon, Mars, or an asteroid's surface, fewer materials need to be brought into space on a launched vehicle. This allows a vehicle to be filled with more necessary supplies for a deep space mission that may not be found in space, like food and fuel. This project's main objective was to develop a 3D printer that uses regolith to "print" large structures, such as a dome, to be used as a heat shield upon a vehicle's reentry into the atmosphere or even a habitat. 3D printing is a growing technology that uses many different methods to mix, heat, and mold a material into a specific shape. In order to heat the regolith enough to stick together into a solid shape, it must be sintered at each layer of material that is laid. Sintering is a process that heats and compresses a powdered material until it fuses into a solid, which requires a lot of energy input. As an alternative, a polymer can be mixed with the regolith before or as it is sent to the 3D printer head to be placed in the specific shape. The addition of the polymer, which melts and binds at much lower temperatures than sintering temperatures, greatly decreases the required heating temperature and energy input. The main task of the project was to identify a functional material for the printer. The first step was to find a miscible. polymer/solvent solution. This solution was added to the regolith and the solvent was evaporated essentially leaving polymer-coated regolith particles. This material would be sent through the printer head and heated layer by layer to melt the polymer and bind the regolith. This method was one of many in a large goal to utilize materials in space with a custom-made 3D printer that builds dome-shaped habitats and other essential equipment for future deep space missions.

Tsoras, Alexandra↗

Testing Lunar Return Thermal Protection Systems using Sub-Scale Flight Test Vehicles

A key objective of NASA's Vision for Space Exploration is to revisit the lunar surface. Such an ambitious goal requires the development of a new human-rated spacecraft, the Orion Crew Exploration Vehicle (CEV), to ferry crews to low earth orbit and to the moon. The successful conclusion of both types of missions will require a thermal protection system (TPS) capable of protecting the vehicle and crew from the extreme heat of atmospheric reentry. As a part of the TPS development, various materials are being tested in arcjet tunnels; however, the combined lunar return aerothermal environment of high heat flux, shear stress, and surface pressure cannot be duplicated using only existing ground test facilities. To ensure full TPS qualification, a flight test program using sub-scale Orion capsules has been proposed to test TPS materials and heat shield construction techniques under the most stressing combination of lunar return aerothermal environments. Originally called Testing Of Reentry Capsule Heat Shield, or TORCH, but later renamed LEX, for Lunar Reentry Experiment, the proposed flight test program is presented along with the driving requirements and descriptions of the vehicle and the TPS instrumentation suite slated to conduct in-flight measurements.

heatshield↗

Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) Mission Overview, Science Return, and Future Applications of This Technology

The Low-Earth Orbit (LEO) Flight Test of an Inflatable Decelerator (LOFTID) mission was the culmination of two decades of research and development for Hyper-sonic Inflatable Aerodynamic Decelerator (HIAD) technology. LOFTID was a project overseen by the Technology Demonstration Mission (TDM) program within NASA’s Space Technology Mission Directorate. The success of the LOFTID mission could enable new NASA missions to Mars, Venus, and most solar-system destinations with atmospheres, as well as cost-effective payload returns to Earth, including in-space manufactured materials and launch vehicle asset recovery. LOFTID, with its unique inflatable heat shield, was the first-of-a-kind orbital reentry flight, and the largest blunt body atmospheric entry of any kind. On November 10, 2022, just over 10 years since the previous flight test of the smaller sub-orbital Inflatable Reentry Vehicle Experiment-3 (IRVE-3) [1], NASA Langley Research Center, with partner United Launch Alliance (ULA), successfully launched and achieved reentry and recovery of the LOFTID Reentry Vehicle (RV), further demonstrating the viability of the HIAD technology for large-diameter, inflatable heat shields to safely and accurately deliver large payloads through an atmosphere via a controlled descent and landing. Launching as a secondary payload with the Joint Polar Satellite System 2 (JPSS-2) from Vandenberg Space Force Base, California, stowed inside an ex-tended payload adapter of the Atlas V 401 launch vehicle, the LOFTID mission officially began after the JPSS-2 payload was delivered to its orbit by the Centaur second stage. The LOFTID RV was flying solo on its spin-stabilized ballistic reentry trajectory about one hour after launch, and the flight ended approximately one hour later with a gentle splashdown under parachute in the Pacific Ocean off the east coast of Hawaii, where the RV was recovered and later shipped back to NASA Langley. LOFTID endured the harsh environments of atmospheric reentry while demonstrating stable aerodynamics through the entire spectrum of hypersonic, supersonic, transonic, and subsonic flight. The LOFTID RV was exposed to an aeroheating environment representative of many Mars and LEO HIAD applications, while successfully demonstrating the ability of the heat-affected inflatable structure to with-stand aerodynamic forces that exceeded those expected at Mars. This flight demonstration of a 6m diameter HIAD confirmed the technology structural and thermal performance as it protected the 1100kg RV entering Earth’s atmosphere at 8 km/s, reaching Mach 30, and experiencing 9g deceleration before deploying parachutes and splashing down in the Pacific Ocean. HIAD technology involves an aeroshell that can be hard packed into a small volume for launch and then deployed prior to atmospheric entry for a controlled deceleration through the atmosphere. Large deployable heat shields enable spacecraft to carry bigger, heavier payloads, including scientific instruments and human support systems for planetary landing and exploration. Much larger than traditional fixed diameter aeroshells that are constrained by the size of launch vehicle shrouds, inflatable decelerators create more drag and start the deceleration process in the upper reaches of the atmosphere with greater efficiency and stability. LOFTID’s successful demonstration of the HIAD technology has greatly expanded NASA’s options for future planetary missions and opened new commercial opportunities for lower-cost mission pay-load recovery including recovery of launch vehicle assets, or as a delivery system to return cargo from the International Space Station or cislunar space. The developments for some of these applications are al-ready underway. While LOFTID was indeed a first-of-a-kind flight for an inflatable heatshield, the largest blunt-body entry ever, its remarkable performance assured that it will not be the last of its kind.

John DiNonno↗

Numerical Analysis of the Transient Response of Advanced Thermal Protection Systems for Atmospheric Entry

Equations for the transfer of heat through thermal protection shields are derived in finite difference form. These equations are applicable to charring ablators, impregnated ceramics, subliming ablators, heat sinks, and insulating materials and have been programed for solution on a high-speed digital computer. In the program, thermal properties can be functions of temperature. Provision is made for analysis of heat shields subjected to simultaneous convective and radiative heat inputs. Some typical results are presented. Limited comparisons with experimental results are made.

THERMAL PROTECTION↗

The disposal of nuclear waste in space

The important problem of disposal of nuclear waste in space is addressed. A prior study proposed carrying only actinide wastes to space, but the present study assumes that all actinides and all fission products are to be carried to space. It is shown that nuclear waste in the calcine (oxide) form can be packaged in a container designed to provide thermal control, radiation shielding, mechanical containment, and an abort reentry thermal protection system. This package can be transported to orbit via the Space Shuttle. A second Space Shuttle delivers an oxygen-hydrogen orbit transfer vehicle to a rendezvous compatible orbit and the mated OTV and waste package are sent to the preferred destination. Preferred locations are either a lunar crater or a solar orbit. Shuttle traffic densities (which vary in time) are given and the safety of space disposal of wastes discussed.

Burns, R. E.↗

Induction Heating of Hypervelocity Impact Samples to 2500 Degrees Centigrade

The Remote Hypervelocity Test Laboratory (RHTL) at White Sands Test Facility (WSTF) was asked to heat samples up to 2500 degrees Centigrade (4532 degrees Fahrenheit) to simulate reentry scenarios of crafts where heated shields are impacted with single small particles ranging from 0.2 to 1.0 millimeters (.008 to.039 inches) of various materials. The team decided an electromagnetic induction (induction heater) was the best method to achieve and control the temperatures in a rapid manner. The samples consisted of three-dimensional carbon-carbon and two-dimensional carbon-phenolic, which are both electrically conductive. After several attempts the team was able to achieve over 2500 degrees Centigrade (4532 degrees Fahrenheit) in ambient atmosphere. When the system was moved to the target chamber and the vacuum system evacuated down to 250 millitorr, arcing occurred between the bus bars and tank, the feedthrough fittings that carried the coolant and current, and between the target sample and coil. To overcome this arcing, conformal coatings, room temperature vulcanization (RTV) silicone, and other non-conductive materials were used to isolate the electromagnetic fields.

Simmons, Joshua↗

Adhesion of Silicone Elastomer Seals for NASA's Crew Exploration Vehicle

Silicone rubber seals are being considered for a number of interfaces on NASA's Crew Exploration Vehicle (CEV). Some of these joints include the docking system, hatches, and heat shield-to-back shell interface. A large diameter molded silicone seal is being developed for the Low Impact Docking System (LIDS) that forms an effective seal between the CEV and International Space Station (ISS) and other future Constellation Program spacecraft. Seals between the heat shield and back shell prevent high temperature reentry gases from leaking into the interface. Silicone rubber seals being considered for these locations have inherent adhesive tendencies that would result in excessive forces required to separate the joints if left unchecked. This paper summarizes adhesion assessments for both as-received and adhesion-mitigated seals for the docking system and the heat shield interface location. Three silicone elastomers were examined: Parker Hannifin S0899-50 and S0383-70 compounds, and Esterline ELA-SA-401 compound. For the docking system application various levels of exposure to atomic oxygen (AO) were evaluated. Moderate AO treatments did not lower the adhesive properties of S0899-50 sufficiently. However, AO pretreatments of approximately 10(exp 20) atoms/sq cm did lower the adhesion of S0383-70 and ELA-SA-401 to acceptable levels. For the heat shield-to-back shell interface application, a fabric covering was also considered. Molding Nomex fabric into the heat shield pressure seal appreciably reduced seal adhesion for the heat shield-to-back shell interface application.

deGroh, Henry C., III↗