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At least 667 records · Page 37

Overview of Ablation Modeling at NASA

The ambitious scientific payload and crew delivery goals of imminent and future NASA missions are associated with challenging and complex vehicle entries. Advanced ablative Thermal Protection System (TPS) materials will be re- quired for such missions, and, as such, a robust ablation modeling capability is critical to assessing performance by bridging the wide gap between ground testing and entry conditions. The traditional ablation modeling toolset - contin- uum thermal/materials response analysis - has grown recently to include high-fidelity, multi-scale, and multi-physics techniques that provide a more complete description of the rich physics and chemistry of ablation to better drive down risks related to extreme entries. The present talk provides a snapshot of ongoing NASA activities in ablation mod- eling, including a review of the current technical capabilities and tools at play within the Agency, the important role academia plays in supporting technical area advancements, and how such internal and external investments intersect with upcoming missions to drive down risks.

Justin Haskins↗

Three-Dimensional Numerical Modeling of Magnetohydrodynamic Augmented Propulsion Experiment

Over the past several years, NASA Marshall Space Flight Center has engaged in the design and development of an experimental research facility to investigate the use of diagonalized crossed-field magnetohydrodynamic (MHD) accelerators as a possible thrust augmentation device for thermal propulsion systems. In support of this effort, a three-dimensional numerical MHD model has been developed for the purpose of analyzing and optimizing accelerator performance and to aid in understanding critical underlying physical processes and nonideal effects. This Technical Memorandum fully summarizes model development efforts and presents the results of pretest performance optimization analyses. These results indicate that the MHD accelerator should utilize a 45deg diagonalization angle with the applied current evenly distributed over the first five inlet electrode pairs. When powered at 100 A, this configuration is expected to yield a 50% global efficiency with an 80% increase in axial velocity and a 50% increase in centerline total pressure.

Turner, M. W.↗

Thermo-Poro-Mechanical Modeling of RTV Intumescence

Room temperature vulcanizing (RTV) silicone is a high-temperature adhesive used as a gap-filler between heatshield tiles in numerous entry missions. Its propensity to intumesce, or swell upon exposure to heat, is a well-known effect that needs to be carefully quantified during design. At tile interfaces of charring ablators, intumescence, combined with differential recession, could cause the gap filler to protrude past the ablator outer mold line, forming a “fence”. Fencing can in turn cause transition to turbulence of the flow wetting the heat shield, leading to augmented surface heating. Recent experiments conducted at the Plasmatron X facility, the high enthalpy wind tunnel of the Center for Hypersonics and Entry Systems Studies, have shown prominent fencing of RTV gap fillers in PICA, under both nitrogen and air plasmas. Similar observations are well known in the arcjet literature. Further experiments under controlled environment, performed using in situ X-ray micro-computed tomography (micro-CT) at the Advanced Light Source (ALS), have shown heating rate-dependent swelling and shrinkage of RTV during pyrolysis. To simulate RTV intumescence, a novel model was introduced in the Porous Materials Analysis Toolbox based on OpenFOAM, PATO, to account for pore-pressure buildup within both closed- and open-pores. The governing equation for the thermo-poro-mechanical response were developed, assuming linear elasticity for the charring silicone. A new multi-pyrolysis model that tracks non-monotonic advancement of material properties with pyrolysis was proposed. This model addresses the limitations of state-of-the-art ablator models to capture the different stages of thermal degradation and coupled thermomechanics. Swelling of RTV was simulated using the new thermo-poro-mechanical model and compared against in situ micro-CT data. Results showed good agreement in intumescence height and temperature profiles at all heating rates, indicating that the key factor contributing to RTV swelling is the internal pressure build-up within closed- and open-pores. As RTV is cured into a soft (rubbery) compound with low-porosity and permeability, initial temperature increase and pyrolysis gas production cause a significant increase of internal pressure, causing a pronounced volume growth. As thermal degradation progresses, rigidization of the silicone occurs due to char hardening which counteract volume shrinkage after gas pressure relief. Overall, our model shows that accounting for changes in properties such porosity, permeability and key thermomechanical coefficients is crucial for capturing the RTV volume change during ablation and enable a predictive capability for heatshield tile interface response. A plan for future calibration of thermomechanical properties evolution during degradation is discussed, as a key next step to close the new model.

RTV↗

Orbiter Boundary Layer Transition Stability Modeling at Flight Entry Conditions

State of the art boundary layer stability modeling capabilities are increasingly seeing application to entry flight vehicles. With the advent of user friendly and robust implementations of two-dimensional chemical nonequilibrium stability modeling with the STABL/PSE-CHEM software, the need for flight data to calibrate such analyses capabilities becomes more critical. Recent efforts to perform entry flight testing with the Orbiter geometry related to entry aerothermodynamics and boundary layer transition is allowing for a heightened focus on the Orbiter configuration. A significant advancement in the state of the art can likely be achieved by establishing a basis of understanding for the occurrence of boundary layer transition on the Orbiter due to discrete protruding gap fillers and the nominal distributed roughness of the actual thermal protection system. Recent success in demonstrating centerline two-dimensional stability modeling on the centerline of the Orbiter at flight entry conditions provides a starting point for additional investigations. The more detailed paper will include smooth Orbiter configuration boundary layer stability results for several typical orbiter entry conditions. In addition, the numerical modeling approach for establishing the mean laminar flow will be reviewed and the method for determining boundary layer disturbance growth will be overviewed. In addition, if actual Orbiter TPS surface data obtained via digital surface scans become available, it may be possible to investigate the effects of an as-flown flight configuration on boundary layer transition compared to a smooth CAD reference.

Bartkowicz, Matt↗

Cryogenic Limb Array Etalon Spectrometer (CLAES) S/N001 cryostat test results

The results of a test program designed to verify the operational and thermal performance of the Cryogenic Limb Array Etalon Spectrometer (CLAES) S/N001 Ne/CO2 cryostat, as monitored and serviced by the Data Acquisition System and Ground Support Equipment, are presented. The S/N001 model test assembly consists of the S/N001 configuration coupled with an instrument thermal mass simulator (ITMS) for simulating the thermal behavior of the instrument package. The test data for the cryostat include the rate of temperature decrease of the various parts of the cryostat/ITMS and the final equilibrium temperatures of the shields, thermal links, and tank. The test results showed that system heat rates, operating temperatures, and expected end-of-life temperature predictions all agree with predictions within estimated margins.

Horsley, W. J.↗

A normative price for energy from an electricity generation system: An Owner-dependent Methodology for Energy Generation (system) Assessment (OMEGA). Volume 1: Summary

The utility owned solar electric system methodology is generalized and updated. The net present value of the system is determined by consideration of all financial benefits and costs (including a specified return on investment). Life cycle costs, life cycle revenues, and residual system values are obtained. Break even values of system parameters are estimated by setting the net present value to zero. While the model was designed for photovoltaic generators with a possible thermal energy byproduct, it applicability is not limited to such systems. The resulting owner-dependent methodology for energy generation system assessment consists of a few equations that can be evaluated without the aid of a high-speed computer.

Chamberlain, R. G.↗

Flutter Analysis of the Thermal Protection Layer on the NASA HIAD

A combination of classical plate theory and a supersonic aerodynamic model is used to study the aeroelastic flutter behavior of a proposed thermal protection system (TPS) for the NASA HIAD. The analysis pertains to the rectangular configurations currently being tested in a NASA wind-tunnel facility, and may explain why oscillations of the articles could be observed. An analysis using a linear flat plate model indicated that flutter was possible well within the supersonic flow regime of the wind tunnel tests. A more complex nonlinear analysis of the TPS, taking into account any material curvature present due to the restraint system or substructure, indicated that significantly greater aerodynamic forcing is required for the onset of flutter. Chaotic and periodic limit cycle oscillations (LCOs) of the TPS are possible depending on how the curvature is imposed. When the pressure from the base substructure on the bottom of the TPS is used as the source of curvature, the flutter boundary increases rapidly and chaotic behavior is eliminated.

Goldman, Benjamin D.↗

Airframe Icing Research Gaps: NASA Perspective

qCurrent Airframe Icing Technology Gaps: Development of a full 3D ice accretion simulation model. Development of an improved simulation model for SLD conditions. CFD modeling of stall behavior for ice-contaminated wings/tails. Computational methods for simulation of stability and control parameters. Analysis of thermal ice protection system performance. Quantification of 3D ice shape geometric characteristics Development of accurate ground-based simulation of SLD conditions. Development of scaling methods for SLD conditions. Development of advanced diagnostic techniques for assessment of tunnel cloud conditions. Identification of critical ice shapes for aerodynamic performance degradation. Aerodynamic scaling issues associated with testing scale model ice shape geometries. Development of altitude scaling methods for thermal ice protections systems. Development of accurate parameter identification methods. Measurement of stability and control parameters for an ice-contaminated swept wing aircraft. Creation of control law modifications to prevent loss of control during icing encounters. 3D ice shape geometries. Collection efficiency data for ice shape geometries. SLD ice shape data, in-flight and ground-based, for simulation verification. Aerodynamic performance data for 3D geometries and various icing conditions. Stability and control parameter data for iced aircraft configurations. Thermal ice protection system data for simulation validation.

Potapczuk, Mark↗

Space Shuttle Main Engine structural analysis and data reduction/evaluation. Volume 2: High pressure oxidizer turbo-pump turbine end bearing analysis

The high-pressure oxidizer turbopump (HPOTP) consists of two centrifugal pumps, on a common shaft, that are directly driven by a hot-gas turbine. Pump shaft axial thrust is balanced in that the double-entry main inducer/impeller is inherently balanced and the thrusts of the preburner pump and turbine are nearly equal but opposite. Residual shaft thrust is controlled by a self-compensating, non-rubbing, balance piston. Shaft hang-up must be avoided if the balance piston is to perform properly. One potential cause of shaft hang-up is contact between the Phase 2 bearing support and axial spring cartridge of the HPOTP main pump housing. The status of the bearing support/axial spring cartridge interface is investigated under current loading conditions. An ANSYS version 4.3, three-dimensional, finite element model was generated on Lockheed's VAX 11/785 computer. A nonlinear thermal analysis was then executed on the Marshall Space Flight Center Engineering Analysis Data System (EADS). These thermal results were then applied along with the interference fit and bolt preloads to the model as load conditions for a static analysis to determine the gap status of the bearing support/axial spring cartridge interface. For possible further analysis of the local regions of HPOTP main pump housing assembly, detailed ANSYS submodels were generated using I-DEAS Geomod and Supertab (Appendix A).

Sisk, Gregory A.↗

Mars Science Laboratory Rover System Thermal Test

On November 26, 2011, NASA launched a large (900 kg) rover as part of the Mars Science Laboratory (MSL) mission to Mars. The MSL rover is scheduled to land on Mars on August 5, 2012. Prior to launch, the Rover was successfully operated in simulated mission extreme environments during a 16-day long Rover System Thermal Test (STT). This paper describes the MSL Rover STT, test planning, test execution, test results, thermal model correlation and flight predictions. The rover was tested in the JPL 25-Foot Diameter Space Simulator Facility at the Jet Propulsion Laboratory (JPL). The Rover operated in simulated Cruise (vacuum) and Mars Surface environments (8 Torr nitrogen gas) with mission extreme hot and cold boundary conditions. A Xenon lamp solar simulator was used to impose simulated solar loads on the rover during a bounding hot case and during a simulated Mars diurnal test case. All thermal hardware was exercised and performed nominally. The Rover Heat Rejection System, a liquid-phase fluid loop used to transport heat in and out of the electronics boxes inside the rover chassis, performed better than predicted. Steady state and transient data were collected to allow correlation of analytical thermal models. These thermal models were subsequently used to predict rover thermal performance for the MSL Gale Crater landing site. Models predict that critical hardware temperatures will be maintained within allowable flight limits over the entire 669 Sol surface mission.

Novak, Keith S.↗

Mars Science Laboratory Rover System Thermal Test

On November 26, 2011, NASA launched a large (900 kg) rover as part of the Mars Science Laboratory (MSL) mission to Mars. The MSL rover is scheduled to land on Mars on August 5, 2012. Prior to launch, the Rover was successfully operated in simulated mission extreme environments during a 16-day long Rover System Thermal Test (STT). This paper describes the MSL Rover STT, test planning, test execution, test results, thermal model correlation and flight predictions. The rover was tested in the JPL 25-Foot Diameter Space Simulator Facility at the Jet Propulsion Laboratory (JPL). The Rover operated in simulated Cruise (vacuum) and Mars Surface environments (8 Torr nitrogen gas) with mission extreme hot and cold boundary conditions. A Xenon lamp solar simulator was used to impose simulated solar loads on the rover during a bounding hot case and during a simulated Mars diurnal test case. All thermal hardware was exercised and performed nominally. The Rover Heat Rejection System, a liquid-phase fluid loop used to transport heat in and out of the electronics boxes inside the rover chassis, performed better than predicted. Steady state and transient data were collected to allow correlation of analytical thermal models. These thermal models were subsequently used to predict rover thermal performance for the MSL Gale Crater landing site. Models predict that critical hardware temperatures will be maintained within allowable flight limits over the entire 669 Sol surface mission.

Novak, Keith S.↗

Arcjet system integration development

Compatibility between an arcjet propulsion system and a communications satellite was verified by testing a Government-furnished, 1.4 kW hydrazine arcjet system with the FLTSATCOM qualification model satellite in a 9.1-meter (30-foot) diameter thermal-vacuum test chamber. Background pressure was maintained at 10(exp -5) torr during arcjet operation by cryopumping the thruster exhaust with an array of 5 K liquid helium cooled panels. Power for the arcjet system was obtained from the FLTSATCOM battery simulator. Spacecraft telemetry was monitored during each thruster firing period. No changes in telemetry data attributable to arcjet operation were detected in any of the tests. Electromagnetic compatibility data obtained included radiated emission measurements, conducted emission measurements, and cable coupling measurements. Significant noise was observed at lower frequencies. Above 500 MHz, radiated emissions were generally within limits, indicating that communication links at S-band and higher frequencies will not be affected. Other test data taken with a diagnostic array of calorimeters, radiometers, witness plates, and a residual gas analyzer evidenced compatible operation, and added to the data base for arcjet system integration. Two test series were conducted. The first series only included the arcjet and diagnostic array operating at approximately 0.1 torr background pressure. The second series added the qualification model spacecraft, a solar panel, and the helium cryopanels. Tests were conducted at 0.1 torr and 10(exp-5) torr. The arcjet thruster was canted 20 degrees relative to the solar panel axis, typical of the configuration used for stationkeeping thrusters on geosynchronous communications satellites.

Zafran, Sidney↗

Space Station multidisciplinary analysis capability - IDEAS2

A computer-aided engineering tool now in use at NASA for systems engineering and integration analysis of the Space Station is described. This tool, called IDEAS2, combines the spacecraft-specific analysis software from NASA IDEAS (truss synthesizer, rigid body controls, plume impingement, orbital lifetime, orbit heat loads, and life support systems) with the interactive graphics, geometric database, solids modeler, structural and thermal analysis capability of the SDRC I-DEAS to obtain an integrated package suitable for Space Station design. This integrated package was applied to a very recent study of growth scenarios of the Space Station dual keel configuration, in which a comparison was made between designs using a combination of photovoltaic solar arrays and solar dynamic collectors, and those using only solar dynamic collectors.

Baker, M.↗

Mechanisms of thermal barrier coating degradation and failure

The objectives and initial results of a Thermal Barrier Coating (TBC) Life Prediction Model Development Program are described. The goals of this program are to: identify and understand TBC failure modes; generate quantitative TBC life data; and develop and verify a TBC life prediction model. The coating being studied is a two layer thermal barrier system incorporating a nominal ten mil outer layer of seven percent yttria partially stabilized zirconia plasma deposited over an inner layer of highly oxidation resistant low pressure plasma sprayed NiCrAlY bond coating. This coating currently is in flight service on turbine vane platforms in the JT-9D and PW2037 engines and is bill-of- material on turbine vane airfoils in the advanced PW4000 and IAE V2500 engines. Effort currently is in progress on the first task, which involves the identification and understanding of TBC failures. Five modes of coating damage were considered: thermomechanical ceramic failure; oxidative bond coat failure; hot corrosion; foreign object damage (FOD); and erosion.

Demasi, J. T.↗

Variable gravity research facility

Spin and despin requirements; sequence of activities required to assemble the Variable Gravity Research Facility (VGRF); power systems technology; life support; thermal control systems; emergencies; communication systems; space station applications; experimental activities; computer modeling and simulation of tether vibration; cost analysis; configuration of the crew compartments; and tether lengths and rotation speeds are discussed.

Allan, Sean↗

Simplified micromechanical equations for thermal residual stress analysis of coated fiber composites

The fabrication of metal matrix composites poses unique problems to the materials engineer. The large thermal expansion coefficient (CTE) mismatch between the fiber and matrix leads to high tensile residual stresses at the fiber/matrix (F/M) interface which could lead to premature matrix cracking during cooldown. Fiber coatings could be used to reduce thermal residual stresses. A simple closed form analysis, based on a three phase composite cylinder model, was developed to calculate thermal residual stresses in a fiber/interphase/matrix system. Parametric studies showed that the tensile thermal residual stresses at the F/M interface were very sensitive to the CTE and thickness of the interphase layer. The modulus of the layer had only a moderate effect on tensile residual stresses. For a silicon carbide titanium aluminide composite, the tangential stresses were 20 to 30 pct. larger than the axial stresses, over a wide range of interphase layer properties, indicating a tendency to form radial matrix cracks during cooldown. Guidelines for the selection of appropriate material properties of the fiber coating were also derived in order to minimize thermal residual stresses in the matrix during fabrication.

Naik, R. A.↗