Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Thermal performance”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Thermal fatigue performance of integrally cast automotive turbine wheels

Fluidized bed thermal fatigue testing was conducted on 16 integrally cast automotive turbine wheels for 1000-10,000 (600 sec total) thermal cycles at 935/50 C. The 16 wheels consisted of 14 IN-792 + 1% Hf and 2 gatorized AF2-1DA wheels; 6 of the IN-792 + Hf wheels contained crack arrest pockets inside the blade root flange. Temperature transients during the thermal cycling were measured in three calibration tests using either 18 or 30 thermocouples per wheel. Thermal cracking based on crack length versus accumulated cycles was greatest for unpocketed wheels developing cracks in 8-13 cycles compared to 75-250 cycles for unpocketed wheels. However, pocketed wheels survived up to 10,000 cycles with crack lengths less than 20 mm, whereas two unpocketed wheels developed 45 mm long cracks in 1000-2000 cycles.

Humphreys, V. E.↗
NASA NTRS2013

Atmospheric Entry Studies for Venus Missions: 45 Sphere-Cone Rigid Aeroshells and Ballistic Entries

The present study considers direct ballistic entries into the atmosphere of Venus using a 45deg sphere-cone rigid aeroshell, a legacy shape that has been used successfully in the past in the Pioneer Venus Multiprobe Mission. For a number of entry mass and heatshield diameter combinations (i.e., various ballistic coefficients) and entry velocities, the trajectory space in terms of entry flight path angles between skip out and -30deg is explored with a 3DoF trajectory code, TRAJ. From these trajectories, the viable entry flight path angle space is determined through the use of mechanical and thermal performance limits on the thermal protection material and science payload; the thermal protection material of choice is entry-grade carbon phenolic, for which a material thermal response model is available. For mechanical performance, a 200 g limit is placed on the peak deceleration load experienced by the science instruments, and 10 bar is assumed as the pressure limit for entry-grade carbon-phenolic material. For thermal performance, inflection points in the total heat load distribution are used as cut off criteria. Analysis of the results shows the existence of a range of critical ballistic coefficients beyond which the steepest possible entries are determined by the pressure limit of the material rather than the deceleration load limit.

Prabhu, Dinesh K.↗
NASA NTRS2016

Development and Implementation of a Generic Analysis Template for Structural-Thermal-Optical-Performance Modeling

Performance-related effects of system level temperature changes can be a key consideration in the design of many types of optical instruments. This is especially true for space-based imagers, which may require complex thermal control systems to maintain alignment of the optical components. Structural-Thermal-Optical-Performance (STOP) analysis is a multi-disciplinary process that can be used to assess the performance of these optical systems when subjected to the expected design environment. This type of analysis can be very time consuming, which makes it difficult to use as a trade study tool early in the project life cycle. In many cases, only one or two iterations can be performed over the course of a project. This limits the design space to best practices since it may be too difficult, or take too long, to test new concepts analytically. In order to overcome this challenge, automation, and a standard procedure for performing these studies is essential. A methodology was developed within the framework of the Comet software tool that captures the basic inputs, outputs, and processes used in most STOP analyses. This resulted in a generic, reusable analysis template that can be used for design trades for a variety of optical systems. The template captures much of the upfront setup such as meshing, boundary conditions, data transfer, naming conventions, and post-processing, and therefore saves time for each subsequent project. A description of the methodology and the analysis template is presented, and results are described for a simple telescope optical system.

Scola, Salvatore↗
NASA NTRS2003

Real-Time Inhibitor Recession Measurements in Two Space Shuttle Reusable Solid Rocket Motors

Real-time internal motor insulation char line recession measurements have been evaluated for two full-scale static tests of the Space Shuttle Reusable Solid Rocket Motor (RSRM). These char line recession measurements were recorded on the forward facing propellant grain inhibitors to better understand the thermal performance of these inhibitors. The RSRM propellant grain inhibitors are designed to erode away during motor operation, thus making it difficult to use post-fire observations to determine inhibitor thermal performance. Therefore, this new internal motor instrumentation is invaluable in establishing an accurate understanding of inhibitor recession versus motor operation time. The data for the first test was presented at the 37th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit (AIAA 2001-3280) in July 2001. Since that time, a second full scale static test has delivered additional real-time data on inhibitor thermal performance. The evaluation of this data is presented in this paper. The second static test, in contrast to the first test, used a slightly different arrangement of instrumentation in the inhibitors. This instrumentation has yielded a better understanding of the inhibitor time dependent inboard tip recession. Graphs of inhibitor recession profiles with time are presented. Inhibitor thermal ablation models have been created from theoretical principals. The model predictions compare favorably with data from both tests. This verified modeling effort is important to support new inhibitor designs for a five segment Space Shuttle solid rocket motor. The internal instrumentation project on RSRM static tests is providing unique opportunities for other real-time internal motor measurements that could not otherwise be directly quantified.

McWhorter, B. B.↗
NASA NTRS1981

Development and validation of cryogenic foam insulation for LH2 subsonic transports

Fourteen foam insulation specimens were tested. Some were plain foam while others contained flame retardants, chopped fiberglass reinforcement and/or vapor barriers. The thermal performance of the insulation was determined by measuring the rate at which LH2 boiled from an aluminum tank insulated with the test material. The test specimens were approximately 50 mm (2 in.) thick. They were structurally scaled so that the test cycle would duplicate the maximum thermal stresses predicted for the thicker insulation of an aircraft liquid hydrogen fuel tank during a typical subsonic flight. The simulated flight cycle of approximately 10 minutes duration heated the other insulation surface to 316 K (110 F) and cooled it to 226 K (20 F) while the inner insulation surface remained at liquid hydrogen temperature of 20 K (-423 F). Two urethane foam insulations exceeded the initial life goal of 2400 simulated flight cycles and sustained 4400 cycles with only minor damage. The addition of fiberglass reinforcement of flame retardant materials to an insulation degraded thermal performance and/or the life of the foam material. Installation of vapor barriers enhanced the structural integrity of the material but did not improve thermal performance. All of the foams tested were available materials; none were developed specifically for LH2 service.

Anthony, F. M.↗
NASA NTRS1993

Calibration of a calorimeter for measuring the performance of thermal control surfaces

The calibration to evaluate the heat leak in terms of the deviation from a perfect calorimeter is described. A thermal vacuum test was carried out to characterize the performance of the calorimeter. The calorimeter was equipped with a heater to control the sample disc temperature and with specific instrumentation to measure the heat leak. The radiation sink temperature of the black cavity target was varied from -192 to +31C, while the heater power was varied from 0 to 311 milliwatts. A steady state thermal model was developed to correlate the results. The calorimeter performance was characterized in terms of the heat leak as a percentage of the ideal heat flow of the calorimeter disc. Large deviations from ideal performance occur at low sink temperatures. The effect of the use of the heater is discussed. The effects of transient conditions during low Earth orbit are discussed. It is concluded that heat leak calibrations are necessary for a wide range of conditions.

Houseman, John↗
NASA NTRS1976

Degradation of a Multilayer Insulation Due to a Seam and a Penetration

The degradation of the thermal performance of a multilayer insulation due to the presence of a seam and a penetration was studied. The multilayer insulation had 30 aluminized Mylar radiation shields with silk net spacers. The seam, an offset butt joint, caused a heat input of 0.169 watt per meter in addition to the basic insulation thermal performance of 0.388 watt per square meter obtained before the installation of the butt joint. The penetration, a fiberglass tank support strut, provided a heat input (including the degradation of the insulation) of 0.543 watt in addition to the basic insulation thermal performance of 0.452 watt per square meter obtained before the penetration.

Sumner, I. E.↗
NASA NTRS2001

Evaluation of GLAS Demonstration Model Loop Heat Pipe Thermal Vacuum Performance with Various Fluid Charges

Two loop heat pipes (LHPs) are to be used for tight thermal control of the Geoscience Laser Altimeter System (GLAS) instrument, planned for flight in late 2001. The LHPs are charged with Propylene as a working fluid. One LHP will be used to transport 110 W from a laser a radiator, the other will transport 190 W from electronic boxes to a separate radiator. The application includes a large amount of thermal mass in each LHP system and low initial startup powers. This along with some non-ideal flight design compromises, such as a less than ideal charge level for this design concept with a symmetrical secondary wick, lead to inadequate performance of the flight LHPs during the flight thermal vacuum test in October of 2000. This presentation focuses on identifying; the sources of the flight test difficulties by modifying the charge and test setup of the successfully tested Development Model Loop Heat Pipe (DM LHP). While very similar to the flight design, the DM L14P did have several significant difference in design and method of testing. These differences were evaluated for affect on performance by conforming the DM LHP to look more like the flight units. The major difference that was evaluated was the relative fill level of the working fluid within the concentrically design LHP compensation chamber. Other differences were also assessed through performance testing including starter heater size and "hot biasing" of major interior components. Performance was assessed with respect to startup, low power operation, conductance, and control heater power. The results of the testing showed that performance improves as initial charge increases, and when the starter heater is made smaller. The "hot biasing" of the major components did not appear to have a detrimental effect. As a result of test results of the DM LHP, modifications are being made to the flight units to increase the fluid charge and increase the watt-density of the starter heater.

Baker, Charles↗
NASA NTRS1993

Performance of thermal adhesives in forced convection

Cooling is critical for the life and performance of electronic equipment. In most cases cooling may be achieved by natural convection but forced convection may be necessary for high wattage applications. Use of conventional type heat sinks may not be feasible from the viewpoint of specific applications and the costs involved. In a heat sink, fins can be attached to the well by ultrasonic welding, by soldering, or with a number of industrially available thermal adhesives. In this paper, the author investigates the heat transfer characteristics of several adhesives and compares them with ultrasonic welding and theoretically calculated values. This experiment was conducted in an air flow chamber. Heat was generated by using heaters mounted on the well. Thermstrate foil, Uniset A401, and Aremco 571 adhesives were tested along with an ultrasonically welded sample. Ultrasonic welding performed far better than the adhesives and Thermstrate foil. This type of experiment can be adapted for a laboratory exercise in an upper level heat transfer course. It gives students an exposure to industrial applications that help them appreciate the importance of the course material.

Kundu, Nikhil K.↗
NASA NTRS1978

Performance evaluations of a liquid solar collector

Report presents procedures and results of thermal performance tests on single-covered liquid solar collector under simulated comditions. Test conditions and data are given in tables and graphs for stagnation tests and thermal performance test. In addition, time-constant test and incident-angle modifier test were conducted to determine transient effect and incident-angle effect. Results of collector load tests are also given.

Shih, K.↗
NASA NTRS1994

Cargo systems manual: Heat Pipe Performance (HPP) STS-66

The purpose of the cargo systems manual (CSM) is to provide a payload reference document for payload and shuttle flight operations personnel during shuttle mission planning, training, and flight operations. It includes orbiter-to-payload interface information and payload system information (including operationally pertinent payload safety data) that is directly applicable to the Mission Operations Directorate (MOD) role in the payload mission. The primary objectives of the heat pipe performance (HPP) are to obtain quantitative data on the thermal performance of heat pipes in a microgravity environment. This information will increase understanding of the behavior of heat pipes in space and be useful for application to design improvements in heat pipes and associated systems. The purpose of HPP-2 is to establish a complete one-g and zero-g data base for axial groove heat pipes. This data will be used to update and correlate data generated from a heat pipe design computer program called Grooved Analysis Program (GAP). The HPP-2 objectives are to: determine heat transport capacity and conductance for open/closed grooved heat pipes and different Freon volumes (nominal, under, and overcharged) using a uniform heat load; determine heat transport capacity and conductance for single/multiple evaporators using asymmetric heat loads; obtain precise static, spin, and rewicking data points for undercharged pipes; investigate heat flux limits (asymmetric heat loads); and determine effects of positive body force on thermal performance.

Napp, Robert↗
NASA NTRS1988

Evaluation of heat sink materials for thermal management of lithium batteries

Aluminum, neopentyl glycol (NPG), and resins FT and KT are evaluated theoretically and experimentally as heat sink materials for lithium battery packs. The thermal performances of the two resins are compared in a thermal vacuum experiment. As solutions to the sublimation property were not immediately apparent, a theoretical comparison of the thermal performance of NPG versus KT, Al, and no material, is presented.

Dimpault-Darcy, E. C.↗
NASA NTRS1980

Evacuated-tube solar collector--performance evaluation

Report gives thermal performance test procedures and results for commercially produced, water-filled, 8-tube collectors. Tests include efficiency, time constant for temperature drop after solar flux is cut, change in efficiency as function of sun angle, and test to see if tubes break when filled with hot water.

Source record↗
NASA NTRS2005

Performance of a Light-Weight Ablative Thermal Protection Material for the Stardust Mission Sample Return Capsule

New tests and analyses are reported that were carried out to resolve testing uncertainties in the original development and qualification of a lightweight ablative material used for the Stardust spacecraft forebody heat shield. These additional arcjet tests and analyses confirmed the ablative and thermal performance of low density Phenolic Impregnated Carbon Ablator (PICA) material used for the Stardust design. Testing was done under conditions that simulate the peak convective heating conditions (1200 W/cm2 and 0.5 atm) expected during Earth entry of the Stardust Sample Return Capsule. Test data and predictions from an ablative material response computer code for the in-depth temperatures were compared to guide iterative adjustment of material thermophysical properties used in the code so that the measured and predicted temperatures agreed. The PICA recession rates and maximum internal temperatures were satisfactorily predicted by the computer code with the revised properties. Predicted recession rates were also in acceptable agreement with measured rates for heating conditions 37% greater than the nominal peak heating rate of 1200 W/sq cm. The measured in-depth temperature response data show consistent temperature rise deviations that may be caused by an undocumented endothermic process within the PICA material that is not accurately modeled by the computer code. Predictions of the Stardust heat shield performance based on the present evaluation provide evidence that the maximum adhesive bondline temperature will be much lower than the maximum allowable of 250 C and an earlier design prediction. The re-evaluation also suggests that even with a 25 percent increase in peak heating rates, the total recession of the heat shield would be a small fraction of the as-designed thickness. These results give confidence in the Stardust heat shield design and confirm the potential of PICA material for use in new planetary probe and sample return applications.

Covington, M. A.↗
NASA NTRS1980

Similarity tests of turbine vanes, effects of ceramic thermal barrier coatings

The role of material thermal conductivity was analyzed for its effect on the thermal performance of air-cooled gas turbine components coated with a ceramic thermal barrier material when tested at reduced temperatures and pressures. It is shown that the thermal performance can be evaluated reliably at reduced gas and coolant conditions; however, thermal conductivity corrections are required for the data at reduced conditions. Corrections for a ceramic thermal barrier coated vane are significantly different than for an uncoated vane. Comparison of uncorrected test data, therefore, would show erroneously that the thermal barrier coating was ineffective. When thermal conductivity corrections are applied to the test data these data are then shown to be representative of engine data and also show that the thermal barrier coating increases the vane cooling effectiveness by 12.5 percent.

Gladden, H. J.↗
NASA NTRS1980

Similarity tests of turbine vanes - Effects of ceramic thermal barrier coatings

The role of material thermal conductivity was analyzed for its effect on the thermal performance of air-cooled gas turbine components coated with a ceramic thermal barrier material when tested at reduced temperatures and pressures. This study shows that the thermal performance can be evaluated reliably at reduced gas and coolant conditions. However, thermal conductivity corrections are required for the data at reduced conditions. These corrections for a ceramic thermal barrier coated vane are significantly different than the corrections for an uncoated vane. Comparison of uncorrected test data, therefore, would show erroneously that the thermal barrier coating was ineffective. When thermal conductivity corrections are applied to the test data these data are then shown to be representative of engine data and also show that the thermal barrier coating increases the vane cooling effectiveness by 12.5 precent.

Gladden, H. J.↗
NASA NTRS2003

System for Testing Thermal Insulation of Pipes

An apparatus and method have been developed for measuring the rates of leakage of heat into pipes carrying liquids, the purpose of the measurements being to quantify the thermal performance of the insulation system. The apparatus is designed primarily for testing pipes used to carry cryogenic liquids, but can also be used for measuring the thermal performance of other insulated pipes or piping systems. The basic measurement principle is straightforward: The outer surface of the pipe insulation is maintained at a fixed warmer temperature. The interior of the pipe is maintained in a narrow fixed lower-temperature range by means of a regular liquid (e.g., water) that is pumped through the pipe at a known flow rate or a cryogenic liquid (e.g., nitrogen) that is saturated at atmospheric pressure and replenished until steady-state conditions are achieved. In the case of water or another liquid pumped through, the inlet and outlet temperatures are measured and heat-leak power is calculated as the mass flow rate of the liquid multiplied by the specific heat of the liquid multiplied by the inlet-to-outlet temperature rise of the liquid. In the case of liquid nitrogen or another low-temperature boiling liquid, the heat-leak power is calculated as the rate of boil-off multiplied by the latent heat of vaporization of the liquid. Then the thermal-insulation performance of the pipe system can be calculated as a function of the measured heat-leak power, the inner and outer boundary temperatures, and the dimensions of the pipe. The apparatus can test as many as three pipes simultaneously. The pipes can have inner diameters up to .15 cm and outer diameters up to .20 cm. The lengths of the pipes may vary; typical lengths are of the order of 18 m. Two thermal guard boxes . one for each end of the pipe(s) under test . are used to make the inlet and outlet fluid connections to the pipe(s) (see figure). The connections include bellows that accommodate thermal expansion and contraction of the pipes. The guard boxes and pipe(s) are positioned so that the pipe(s) slope upward from the upstream to the downstream end at an angle of at least 2 . The upward slope allows vapor bubbles to accumulate at the downstream end. The thermal guard boxes keep the ends of the pipes at the lower interior temperature to prevent spurious lengthwise leakage of heat into the pipes. It is important to prevent this spurious heat leakage because, if it were allowed to occur, it could contribute a large error in the measured heat-leak power. The upstream thermal guard box includes a heat exchanger through which liquid flowing into the pipe(s) is subcooled to the saturation temperature corresponding to the ambient pressure. Conversely, this heat exchanger can also be used to warm the flowing liquid to a desired fixed temperature. The apparatus includes a temperature control device that is placed around each pipe under test. Each device is operated under thermostatic control to maintain the outer surface of the pipe insulation at the specified test temperature. All measurements are recorded on a portable data-acquisition system.

Fesmire, James E.↗