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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 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 NTRS2011

Autonomous Aerobraking: Thermal Analysis and Response Surface Development

A high-fidelity thermal model of the Mars Reconnaissance Orbiter was developed for use in an autonomous aerobraking simulation study. Response surface equations were derived from the high-fidelity thermal model and integrated into the autonomous aerobraking simulation software. The high-fidelity thermal model was developed using the Thermal Desktop software and used in all phases of the analysis. The use of Thermal Desktop exclusively, represented a change from previously developed aerobraking thermal analysis methodologies. Comparisons were made between the Thermal Desktop solutions and those developed for the previous aerobraking thermal analyses performed on the Mars Reconnaissance Orbiter during aerobraking operations. A variable sensitivity screening study was performed to reduce the number of variables carried in the response surface equations. Thermal analysis and response surface equation development were performed for autonomous aerobraking missions at Mars and Venus.

Dec, John A.↗
NASA NTRSDate not supplied

MEDLI2: Ablator Models for Flight

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[1,2]. Discrepancies between MEDLI’s flight heating data and predicted thermal response and recession[3] led to a campaign to characterize and test Mars2020 and MEDLI2 (Mars Entry, Descent, and Landing Instrument 2)[4] flight-lot thermal protection materials (TPS) at relevant ground test conditions. This paper covers the development and validation of flight-lot material response models for the heatshield material Phenolic Impregnated Carbon Ablator (PICA) and the backshell material SLA-561V (Super Lightweight Ablator). Virgin and char properties such as thermal conductivity and emissivity were updated as a function of temperature and pressure in the MEDLI2 specific material response models. Other material properties were characterized and determined to be similar to prior TPS material measurements used for the heritage models. MEDLI2 also investigated the impact of the protective coating that was applied to MSL and Mars2020’s PICA heatshield to reduce the particulate matter shedding. Evidence from arc jet experiments on coated PICA showed that although there were no adverse effects on the performance of the TPS, there were differences in the in-depth temperatures and reduced recession compared to material response predictions of uncoated PICA.[5,6]. Fully Implicit Ablation and Thermal response (FIAT)[7] material response simulations were conducted with the heritage and MEDLI2 flight-lot material models. FIAT-predicted in-depth temperature responses and recession predictions were compared to flight-lot certification ground-test arc jet test data and MEDLI2 thermocouple data.

Materials Response↗
NASA NTRS1996

Layerwise Finite Elements for Smart Piezoceramic Composite Plates in Thermal Environments

Analytical formulations are presented which account for the coupled mechanical, electrical, and thermal response of piezoelectric composite laminates and plate structures. A layerwise theory is formulated with the inherent capability to explicitly model the active and sensory response of piezoelectric composite plates having arbitrary laminate configurations in thermal environments. Finite element equations are derived and implemented for a bilinear 4-noded plate element. Application cases demonstrate the capability to manage thermally induced bending and twisting deformations in symmetric and antisymmetric composite plates with piezoelectric actuators, and show the corresponding electrical response of distributed piezoelectric sensors. Finally, the resultant stresses in the thermal piezoelectric composite laminates are investigated.

Saravanos, Dimitris A.↗
NASA NTRS1985

Response time correlations for platinum resistance thermometers in flowing fluids

The thermal response of two types of Platinum Resistance Thermometers (PRT's), which are being considered for use in the National Transonic Wind Tunnel Facility, were studied. Response time correlations for each PRT, in flowing water, oil and air, were established separately. A universal correlation, tau WOA = 2.0 + 1264, 9/h, for a Hy-Cal Sensor (with a reference resistance of 100 ohm) within an error of 20% was established while the universal correlation for the Rosemount Sensor (with a reference resistance of 1000 ohm), tau OA = 0.122 + 1105.6/h, was found with a maximum percentage error of 30%. The correlation for the Rosemount Sensor was based on air and oil data only which is certainly not sufficient to make a correlation applicable to every condition. Therefore, the correlation needs more data to be gathered in different fluids. Also, it is necessary to state that the calculation of the parameter, h, was based on the available heat transfer correlations, whose accuracies are already reported in literature uncertain within 20-30%. Therefore, the universal response constant correlations established here for the Hy-Cal and Rosemount sensors are consistent with the uncertainty in the input data and are recommended for future use in flowing liquids and gases.

Pandey, D. K.↗
NASA NTRS2016

Transient Thermal Analysis of the 12.5 KW HERMeS Hall Thruster

NASA’s Dawn spacecraft entered orbit around the dwarf planet Ceres on March 6, 2015 thereby becoming the first mission to orbit two celestial bodies outside of the Earth-moon system. This feat was made possible due to the use of electric propulsion (EP) to provide the 11 km/s of delta-V that was necessary. This type of propulsion is also enabling for many of NASA’s future missions, such as the proposed Asteroid Redirect Robotic Mission (ARRM). Specifically, the Hall Effect Rocket with Magnetic Shielding (HERMeS) is the propulsive design baseline due to its long operational lifetime and high specific impulse. Extensive modeling is required to verify the thruster's predicted thermal performance and survival during the bounding mission phases. Since the thruster would constrain plasma on the order of tens of thousands of hours, components in the thruster must be able to accommodate much higher sustained temperatures than typical spacecraft hardware. Furthermore, the high temperatures necessitate care in analyzing thermal gradients within the thruster. In this paper, we describe a thermal model of the HERMeS thruster and focus upon the thruster's transient performance under different initial firing conditions. The thruster's transient thermal response is critical to verify margin in the thermal stresses inherent in the thruster and to determine safe start-up procedures to protect the thruster's critical components. This model is used to perform transient thermal analysis of the thruster during startup and shutdown, provide temperature maps for structural analysis, and plan for thermal cycle testing conducted at JPL. The thermal cycle testing includes thruster starts at the cold temperature limits and operation at maximum expected external heat flux. This work will support laboratory environmental testing of the HERMeS thruster, which is being conducted at JPL in 2016.

Reilly, Sean↗
NASA NTRS2019

Probabilistic Risk Analysis and Margin Process for a Flexible Thermal Protection System

Atmospheric entry vehicle thermal protection systems are margined due to the uncertainties that exist in entry aeroheating environments and the thermal response of the materials and structures. Entry vehicle thermal protections systems are traditionally over-margined for the heat loads that are experienced along the entry trajectory by designing to survive stacked worst-case scenarios. Additionally, the conventional heat shield design and margin process offers very little insight into the risk of over-temperature during flight and the corresponding reliability of the heat shield performance. A probabilistic margin process can be used to appropriately margin the thermal protection system based on rigorously calculated risk of failure. This probabilistic margin process allows engineers to make informed aeroshell design, entry-trajectory design, and risk trades while preventing excessive margin from being applied. This study presents the methods of the probabilistic margin process and how the uncertainty analysis is used to determine the reliability of the entry vehicle thermal protection system and associated risks of failure.

Tobin, Steven A.↗
NASA NTRS1998

Thermal Effects Modeling Developed for Smart Structures

Applying smart materials in aeropropulsion systems may improve the performance of aircraft engines through a variety of vibration, noise, and shape-control applications. To facilitate the experimental characterization of these smart structures, researchers have been focusing on developing analytical models to account for the coupled mechanical, electrical, and thermal response of these materials. One focus of current research efforts has been directed toward incorporating a comprehensive thermal analysis modeling capability. Typically, temperature affects the behavior of smart materials by three distinct mechanisms: Induction of thermal strains because of coefficient of thermal expansion mismatch 1. Pyroelectric effects on the piezoelectric elements; 2. Temperature-dependent changes in material properties; and 3. Previous analytical models only investigated the first two thermal effects mechanisms. However, since the material properties of piezoelectric materials generally vary greatly with temperature (see the graph), incorporating temperature-dependent material properties will significantly affect the structural deflections, sensory voltages, and stresses. Thus, the current analytical model captures thermal effects arising from all three mechanisms through thermopiezoelectric constitutive equations. These constitutive equations were incorporated into a layerwise laminate theory with the inherent capability to model both the active and sensory response of smart structures in thermal environments. Corresponding finite element equations were formulated and implemented for both the beam and plate elements to provide a comprehensive thermal effects modeling capability.

Lee, Ho-Jun↗
NASA NTRS2001

Thermal/Fluid Analysis of a Composite Heat Exchanger for Use on the RLV Rocket Engine

As part of efforts to design a regeneratively cooled composite nozzle ramp for use on the reusable vehicle (RLV) rocket engine, a C-SiC composite heat exchanger concept was proposed for thermal performance evaluation. To test the feasibility of the concept, sample heat exchanger panels were made to fit the Glenn Research Center's cell 22 for testing. Operation of the heat exchanger was demonstrated in a combustion environment with high heat fluxes similar to the RLV Aerospike Ramp. Test measurements were reviewed and found to be valuable for the on-going fluid and thermal analysis of the actual RLV composite ramp. Since the cooling fluid for the heat exchanger is water while the RLV Ramp cooling fluid is LH2, fluid and therma models were constructed to correlate to the specific test set-up. The knowledge gained from this work will be helpful for analyzing the thermal response of the actual RLV Composite Ramp. The coolant thermal properties for the models are taken from test data. The heat exchanger's cooling performance was analyzed using the Generalized Fluid System Simulation Program (GFSSP). Temperatures of the heat exchanger's structure were predicted in finite element models using Patran and Sinda. Results from the analytical models and the tests show that RSC's heat exchanger satisfied the combustion environments in a series of 16 tests.

Nguyen, Dalton↗
NASA NTRS2002

Thermal/Fluid Analysis of a Composite Heat Exchanger for Use on the RLV Rocket Engine

As part of efforts to design a regeneratively cooled composite nozzle ramp for use on the reusable vehicle (RLV) rocket engine, an C-SiC composites heat exchanger concept was proposed for thermal performance evaluation. To test the feasibility of the concept, sample heat exchanger panels were made to fit the Glenn Research Center's cell 22 for testing. Operation of the heat exchanger was demonstrated in a combustion environment with high heat fluxes similar to the RLV Aerospike Ramp. Test measurements were reviewed and found to be valuable for the on going fluid and thermal analysis of the actual RLV composite ramp. Since the cooling fluid for the heat exchanger is water while the RLV Ramp cooling fluid is LH2, fluid and thermal models were constructed to correlate to the specific test set-up. The knowledge gained from this work will be helpful for analyzing the thermal response of the actual RLV Composite Ramp. The coolant thermal properties for the models are taken from test data. The heat exchanger's cooling performance was analyzed using the Generalized Fluid System Simulation Program (GFSSP). Temperatures of the heat exchanger's structure were predicted in finite element models using Patran and Sinda. Results from the analytical models and the tests show that RSC's heat exchanger satisfied the combustion environments in a series of 16 tests.

Nguyen, Dalton↗
NASA NTRS1991

Development of advanced modal methods for calculating transient thermal and structural response

Higher-order modal methods for predicting thermal and structural response are evaluated. More accurate methods or ones which can significantly reduce the size of complex, transient thermal and structural problems are desirable for analysis and are required for synthesis of real structures subjected to thermal and mechanical loading. A unified method is presented for deriving successively higher-order modal solutions related to previously-developed, lower-order methods such as the mode displacement and mode-acceleration methods. A new method, called the force-derivative method, is used to obtain higher-order modal solutions for both uncoupled (proportionally-damped) structural problems as well as thermal problems and coupled (non-proportionally damped) structural problems. The new method is called the force-derivative method because, analogous to the mode-acceleration method, it produces a term that depends on the forcing function and additional terms that depend on the time derivatives of the forcing function.

Camarda, Charles J.↗
NASA NTRS1991

Thermally-induced structural dynamic response of flexural configurations influenced by linear/non-linear thermal effects

The thermally-induced strucural dynamic response of flexural configurations influenced by linear/nonlinear thermal effects is presented in conjunction with 'unified' transient approaches for effectively tackling this class of interdisciplinary problems. For illustrative purposes, the flexural structural models are assumed to be of the Euler-Bernoulli type. The purpose of the present paper is to not only provide an understanding of the influence of general linear/nonlinear thermal effects on flexural configurations, but also to provide to the analyst effective computational tools which help preserve a unified technology for the interdisciplinary areas encompassing structural mechanics/dynamics and thermal sciences. Several numerical test models illustrate the representative thermally-induced structural dynamic response of flexural configurations subjected to general linear/nonlinear temperature effects.

Namburu, Raju R.↗
NASA NTRS1982

Space Shuttle orbiter entry heating and TPS response: STS-1 predictions and flight data

Aerothermodynamic development flight test data from the first orbital flight test of the Space Transportation System (STS) transmitted after entry blackout is given. Engineering predictions of boundary layer transition and numerical simulations of the orbiter flow field were confirmed. The data tended to substantiate preflight predictions of surface catalysis phenomena. The thermal response of the thermal protection system was as expected. The only exception is that internal free convection was found to be significant in limiting the peak temperature of the structure in areas which do not have internal insulation.

Ried, R. C.↗
NASA NTRS2000

Curved Thermopiezoelectric Shell Structures Modeled by Finite Element Analysis

"Smart" structures composed of piezoelectric materials may significantly improve the performance of aeropropulsion systems through a variety of vibration, noise, and shape-control applications. The development of analytical models for piezoelectric smart structures is an ongoing, in-house activity at the NASA Glenn Research Center at Lewis Field focused toward the experimental characterization of these materials. Research efforts have been directed toward developing analytical models that account for the coupled mechanical, electrical, and thermal response of piezoelectric composite materials. Current work revolves around implementing thermal effects into a curvilinear-shell finite element code. This enhances capabilities to analyze curved structures and to account for coupling effects arising from thermal effects and the curved geometry. The current analytical model implements a unique mixed multi-field laminate theory to improve computational efficiency without sacrificing accuracy. The mechanics can model both the sensory and active behavior of piezoelectric composite shell structures. Finite element equations are being implemented for an eight-node curvilinear shell element, and numerical studies are being conducted to demonstrate capabilities to model the response of curved piezoelectric composite structures (see the figure).

Lee, Ho-Jun↗
NASA NTRSDate not supplied

Material Response Analysis of PICA-NuSil at the Hypersonic Materials Environmental Test System

Phenolic impregnated carbon ablator (PICA) is a thermal protection system (TPS) that gained heritage during the Stardust mission and is the baseline TPS material for expeditions to Mars (e.g., Mars Science Laboratory and Mars 2020).1 PICA is friable; therefore, a polysiloxane resin (NuSil CV-1144-0) is deposited onto the surface of flight hardware before assembly, test, and launch operations (ATLO) to mitigate particle shedding. Furthermore, the Mars Science Laboratory (MSL) and Mars 2020 heatshields were instrumented with an array of thermocouple plugs (MEDLI Integrated Sensor Plugs or MISP) designed to measure the in-depth thermal response of PICA during flight. Post-flight analysis of MISP data suggests that recession was lower than predicted, and NuSil is suspected as a reasonable cause. NuSil primarily consists of a copolymer blend of dimethyl and diphenyl-substituted polysiloxane resin. Four decades of research in the field of polymer-derived ceramics demonstrate that polysiloxane resins form oxidation-resistant silicon oxycarbide materials at pyrolysis temperatures as low as 800 °C. -2 Furthermore, silicon oxycarbide materials exhibit excellent thermal stability at relatively high-temperatures (< 1200 °C) but eventually decompose via carbothermal reduction. -3 Therefore, it is reasonable to suggest that NuSil forms an oxidation-resistant thermal barrier coating on the TPS surface and influences the material response during atmospheric entry. Evidence for this notion was observed during prior arc-jet testing at both the Aerodynamic Heating Facility (AHF) and the Panel Test Facility (PTF) at the NASA Ames Research Center. Articles of PICA were subjected to testing in airflow, and a temperature discontinuity was recorded at the surface of coupons coated with NuSil. A glassy coating was observed to form at the surface in shear flow configurations (wedge and panel) under moderate heat fluxes. Furthermore, state-of-the-art material response models have yet to account for the underlying physics and chemistry of the temperature discontinuity. To this end, a pathfinder test campaign was executed at the Hypersonic Materials Environmental Test System (HyMETS) to enhance the fundamental understanding of PICA-NuSil ablation phenomena and provide a preliminary data set to aid in the development of material response models.

PICA↗
NASA NTRSDate not supplied

Acquisition and Processing of Passive Thermography Fatigue Test Data

Passive thermography is a large area inspection technique and is commonly used during fatigue loading of composite structures. Passive thermography allows for real time tracking of damage. This provides an accurate assessment of the damage progression as a function of load cycles and determines when the loading is stopped for ultrasonic or X-ray CT evaluations for detail damage assessment through-the-thickness. In addition, if new areas of damage growth are detected, with passive thermography, then those areas are additionally inspected with ultrasound or X-ray CT. A custom developed passive thermography inspection system for fatigue loading is presented. Passive thermography of damage progression is compared to ultrasound inspections as a function of the fatigue cycles. Discussion is made on processing of the thermal data for periodic and non-periodic thermal responses.

Thermal nondestructive evaluation↗
NASA NTRS1988

Computer codes for thermal analysis of a solid rocket motor nozzle

A number of computer codes are available for performing thermal analysis of solid rocket motor nozzles. Aerotherm Chemical Equilibrium (ACE) computer program can be used to perform one-dimensional gas expansion to determine the state of the gas at each location of a nozzle. The ACE outputs can be used as input to a computer program called Momentum/Energy Integral Technique (MEIT) for predicting boundary layer development development, shear, and heating on the surface of the nozzle. The output from MEIT can be used as input to another computer program called Aerotherm Charring Material Thermal Response and Ablation Program (CMA). This program is used to calculate oblation or decomposition response of the nozzle material. A code called Failure Analysis Nonlinear Thermal and Structural Integrated Code (FANTASTIC) is also likely to be used for performing thermal analysis of solid rocket motor nozzles after the program is duly verified. A part of the verification work on FANTASTIC was done by using one and two dimension heat transfer examples with known answers. An attempt was made to prepare input for performing thermal analysis of the CCT nozzle using the FANTASTIC computer code. The CCT nozzle problem will first be solved by using ACE, MEIT, and CMA. The same problem will then be solved using FANTASTIC. These results will then be compared for verification of FANTASTIC.

Chauhan, Rajinder Singh↗