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  • 1
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    Analysis of developing laminar flows in circular pipes using a higher-order finite-difference technique (1995)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: A higher-order finite-difference technique is developed to calculate the developing-flow field of steady incompressible laminar flows in the entrance regions of circular pipes. Navier-Stokes equations governing the motion of such a flow field are solved by using this new finite-difference scheme. This new technique can increase the accuracy of the finite-difference approximation, while also providing the option of using unevenly spaced clustered nodes for computation such that relatively fine grids can be adopted for regions with large velocity gradients. The velocity profile at the entrance of the pipe is assumed to be uniform for the computation. The velocity distribution and the surface pressure drop of the developing flow then are calculated and compared to existing experimental measurements reported in the literature. Computational results obtained are found to be in good agreement with existing experimental correlations and therefore, the reliability of the new technique has been successfully tested.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: The Sixth Annual Thermal and Fluids Analysis Workshop; p 107-114
    Format: application/pdf
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  • 2
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    Heat transfer in a real engine environment (1985)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The hot section facility at the Lewis Research Center was used to demonstrate the capability of instruments to make required measurements of boundary conditions of the flow field and heat transfer processes in the hostile environment of the turbine. The results of thermal scaling tests show that low temperature and pressure rig tests give optimistic estimates of the thermal performance of a cooling design for high pressure and temperature application. The results of measuring heat transfer coefficients on turbine vane airfoils through dynamic data analysis show good comparison with measurements from steady state heat flux gauges. In addition, the data trends are predicted by the STAN5 boundary layer code. However, the magnitude of the experimental data was not predicted by the analysis, particularly in laminar and transitional regions near the leading edge. The infrared photography system was shown capable of providing detailed surface thermal gradients and secondary flow features on a turbine vane and endwell.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Turbine Engine Hot Section Technology, 1985; p 139-150
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  • 3
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    Thermal/structural analyses of several hydrogen-cooled leading-edge concepts for hypersonic flight vehicles (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The aerodynamic heating at high flight Mach numbers, when shock interference heating is included, can be extremely high and can exceed the capability of most conventional metallic and potential ceramic materials available. Numerical analyses of the heat transfer and thermal stresses are performed on three actively cooled leading-edge geometries (models) made of three different materials to address the issue of survivability in a hostile environment. These analyses show a mixture of results from one configuration to the next. Results for each configuration are presented and discussed. Combinations of enhanced internal film coefficients and high material thermal conductivity of copper and tungsten are predicted to maintain the maximum wall temperature for each concept within acceptable operating limits. The exception is the TD nickel material which is predicted to melt for most cases. The wide range of internal impingement film coefficients (based on correlations) for these conditions can lead to a significant uncertainty in expected leading-edge wall temperatures. The equivalent plastic strain, inherent in each configuration which results from the high thermal gradients, indicates a need for further cyclic analysis to determine component life.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: AIAA PAPER 90-0053
    Format: text
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  • 4
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    Summary of experimental heat-transfer results from the turbine hot section facility (1993)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experimental data from the turbine Hot Section Facility are presented and discussed. These data include full-coverage film-cooled airfoil results as well as special instrumentation results obtained at simulated real engine conditions. Local measurements of airfoil wall temperature, airfoil gas-path static-pressure distribution, and local heat-transfer coefficient distributions are presented and discussed. In addition, measured gas and coolant temperatures and pressures are presented. These data are also compared with analyses from Euler and boundary-layer codes.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TP-3250 , E-6615 , NAS 1.60:3250
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  • 5
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    Heat transfer in aerospace propulsion (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Presented is an overview of heat transfer related research in support of aerospace propulsion, particularly as seen from the perspective of the NASA Lewis Research Center. Aerospace propulsion is defined to cover the full spectrum from conventional aircraft power plants through the Aerospace Plane to space propulsion. The conventional subsonic/supersonic aircraft arena, whether commercial or military, relies on the turbine engine. A key characteristic of turbine engines is that they involve fundamentally unsteady flows which must be properly treated. Space propulsion is characterized by very demanding performance requirements which frequently push systems to their limits and demand tailored designs. The hypersonic flight propulsion systems are subject to severe heat loads and the engine and airframe are truly one entity. The impact of the special demands of each of these aerospace propulsion systems on heat transfer is explored.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-100874 , E-4105 , NAS 1.15:100874
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  • 6
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    HOST turbine heat transfer program summary (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The objectives of the HOST Turbine Heat Transfer subproject were to obtain a better understanding of the physics of the aerothermodynamic phenomena and to assess and improve the analytical methods used to predict the flow and heat transfer in high temperature gas turbines. At the time the HOST project was initiated, an across-the-board improvement in turbine design technology was needed. A building-block approach was utilized and the research ranged from the study of fundamental phenomena and modeling to experiments in simulated real engine environments. Experimental research accounted for approximately 75 percent of the funding with the remainder going to analytical efforts. A healthy government/industry/university partnership, with industry providing almost half of the research, was created to advance the turbine heat transfer design technology base.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-100280 , E-3876 , NAS 1.15:100280
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  • 7
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    Thermal/structural analyses of several hydrogen-cooled leading-edge concepts for hypersonic flight vehicles (1990)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The aerodynamic heating at high flight Mach numbers, when shock interference heating is included, can be extremely high and can exceed the capability of most conventional metallic and potential ceramic materials available. Numerical analyses of the heat transfer and thermal stresses are performed on three actively cooled leading-edge geometries (models) made of three different materials to address the issue of survivability in a hostile environment. These analyses show a mixture of results from one configuration to the next. Results for each configuration are presented and discussed. Combinations of enhanced internal film coefficients and high material thermal conductivity of copper and tungsten are predicted to maintain the maximum wall temperature for each concept within acceptable operating limits. The exception is the TD nickel material which is predicted to melt for most cases. The wide range of internal impingement film coefficients (based on correlations) for these conditions can lead to a significant uncertainty in expected leading-edge wall temperatures. The equivalent plastic strain, inherent in each configuration which results from the high thermal gradients, indicates a need for further cyclic analysis to determine component life.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-102391 , E-4788 , NAS 1.15:102391 , AIAA PAPER 90-0053 , Aerospace Sciences Meeting; Jan 08, 1990 - Jan 11, 1990; Reno, NV; United States
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  • 8
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    A high heat flux experiment for verification of thermostructural analysis (1988)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A major concern in advancing the state of the art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of handling the high heat fluxes during flight. The leading edges of such systems must not only tolerate the maximum heating rates, but must also minimize distortions to the flow field due to excessive blunting and/or thermal warping of the compression surface to achieve the high inlet performance required. A combined analytical and experimental effort to study the aerothermodynamic loads on actively cooled structures for hypersonic applications was established. A hydrogen/oxygen rocket engine was modified to establish a high enthalpy high heat flux environment. The facility provides heat flux levels from about 200 up to 10000 Btu/sq ft/sec. Cross flow and parallel flow regeneratively cooled model can be tested and analyzed by using cooling fluids of water and hydrogen. Results are presented of the experiment and the characteristics of the Hot Gas Test Facility. The predicted temperature results of the cross flow model are compared with the experimental data on the first monolithic specimens and are found to be in good agreement. Thermal stress analysis results are also presented.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-100931 , E-4202 , NAS 1.15:100931 , Winter Annual Meeting of the ASME; Nov 28, 1988 - Dec 02, 1988; Chicago, IL; United States
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  • 9
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    Experience with advanced instrumentation in a hot section cascade (1989)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Lewis Research Center gas turbine Hot Section Test Facility was developed to provide a real engine environment with known boundary conditions for the aerothermal performance evaluation and verification of computer design codes. This verification process requires experimental measurements in a hostile environment. The research instruments used in this facility are presented, and their characteristics and how they perform in this environment are discussed. The research instrumentation consisted of conventional pressure and temperature sensors, as well as thin-film thermocouples and heat flux gages. The hot gas temperature was measured by an aspirated temperature probe and by a dual-element, fast-response temperature probe. The data acquisition mode was both steady state and time dependent. These experiments were conducted over a wide range of gas Reynolds numbers, exit gas Mach numbers, and heat flux levels. This facility was capable of testing at temperatures up to 1600 K, and at pressures up to 18 atm. These corresponded to an airfoil exit Reynolds number range of 0.5 x 10(6) to 2.5 x 10(6) based on the airfoil chord of 5.55 cm. The results characterize the performance capability and the durability of the instrumentation. The challenge of making measurements in hostile environments is also discussed. The instruments exhibited more than adequate durability to achieve the measurement profile. About 70 percent of the thin-film thermocouples and the dual-element temperature probe survived several hundred thermal cycles and more than 35 hr at gas temperatures up to 1600 K. Within the experimental uncertainty, the steady-state and transient heat flux measurements were comparable and consistent over the range of Reynolds numbers tested.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-102294 , E-4962 , NAS 1.15:102294 , Winter Annual Meeting of the American Society of Mechanical Engineers; Dec 10, 1989 - Dec 15, 1989; San Francisco, CA; United States
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  • 10
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    Hypersonic engine component experiments in high heat flux, supersonic flow environment (1993)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A major concern in advancing the state-of-the-art technologies for hypersonic vehicles is the development of an aeropropulsion system capable of withstanding the sustained high thermal loads expected during hypersonic flight. Even though progress has been made in the computational understanding of fluid dynamics and the physics/chemistry of high speed flight, there is also a need for experimental facilities capable of providing a high heat flux environment for testing component concepts and verifying/calibrating these analyses. A hydrogen/oxygen rocket engine heat source was developed at the NASA Lewis Research Center as one element in a series of facilities at national laboratories designed to fulfill this need. This 'Hot Gas Facility' is capable of providing heat fluxes up to 450 w/sq cm on flat surfaces and up to 5,000 w/sq cm at the leading edge stagnation point of a strut in a supersonic flow stream. Gas temperatures up to 3050 K can also be attained. Two recent experimental programs conducted in this facility are discussed. The objective of the first experiment is to evaluate the erosion and oxidation characteristics of a coating on a cowl leading edge (or strut leading edge) in a supersonic, high heat flux environment. Macrophotographic data from a coated leading edge model show progressive degradation over several thermal cycles at aerothermal conditions representative of high Mach number flight. The objective of the second experiment is to assess the capability of cooling a porous surface exposed to a high temperature, high velocity flow environment and to provide a heat transfer data base for a design procedure. Experimental results from transpiration cooled surfaces in a supersonic flow environment are presented.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-106273 , E-8002 , NAS 1.15:106273 , International Symposium on Optical Applied Science and Engineering; Jul 11, 1993 - Jul 16, 1993; San Diego, CA; United States
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