Results for “Compressor Heat” — Page 14

NASA NTRS2019

TPSAS-NF1676L-32540-DND

The supersonic/hypersonic ground testing serves a multitude of purposes and objectives. From structural and material integrity, to force and moment models; from pressure mapping, thermal imaging, shock structure to scramjet/ramjet engine operability and ignition studies. Hypersonics has become more of a high profile area of late. Compared to flight testing, ground testing typically it is less expensive and usually is not a one single-test test program. It is more forgiving in the fact that a test program can utilize a single test article and get data from a number of tests. Ground test data is the input to the formation of research databases, as well as, validation of CFD models/tools. To understand the output, one must first understand the inputs. To achieve supersonic/hypersonic speeds the use of compressors or compressed air storage along with converging-diverging nozzles are needed. Flight temperatures also need to be achieved for the ground test to truly simulate flight conditions. At NASA Langley's 8' High Temperature Tunnel (HTT), compressed natural gas (methane) is burned to heat the air. To validate that the desired test condition is met, test instrumentation is used and data is collected from the combustor all the way through the nozzle exit plane. In this talk, NASA Langley's 8' HTT will be the basis for what kind of test capabilities hypersonic ground testing offers. It certainly does not offer every capability within ground testing, nor will the discussion cover all types. Specific test techniques used and data collection methods will be discussed and how these confirm certain flight conditions have been achieved. There will be subsequent thoughts on experimental methods that aid test articles such as thrust stands, model actuation, etc.

Aaron W Fuchs↗

NASA NTRS2017

Viscous Three-Dimensional Simulation of Flow in an Axial Low Pressure Compressor at Engine Icing Operating Points

Viscous three-dimensional simulations of the Honeywell ALF502R-5 low pressure compressor (sometimes called a booster) using the NASA Glenn code GlennHT have been carried out. A total of ten simulations were produced. Five operating points are investigated, with each point run with two different wall thermal conditions. These operating points are at, or near, points where engine icing has been determined to be likely. In the future, the results of this study will be used for further analysis such as predicting collection efficiency of ice particles and ice growth rates at various locations in the compressor. A mixing plane boundary condition is used between each blade row, resulting in convergence to steady state within each blade row. The k-omega turbulence model of Wilcox, combined with viscous grid spacing near the wall on the order of one, is used to resolve the turbulent boundary layers. For each of the operating points, heat transfer coefficients are generated on the blades and walls. The heat transfer coefficients are produced by running the operating point with two different wall thermal conditions and then solving simultaneously for the heat transfer coefficient and adiabatic wall temperature at each point. Average Nusselt numbers are calculated for the most relevant surfaces. The values are seen to scale with Reynolds number to approximately a power of 0.7. Additionally, images of surface distribution of Nusselt number are presented. Qualitative comparison between the five operating points show that there is relatively little change in the character of the distribution. The dominant observed effect is that of an overall scaling, which is expected due to Reynolds number differences. One interesting aspect about the Nusselt number distribution is observed on the casing (outer diameter) downstream of the exit guide vanes (EGVs). The Nusselt number is relatively high between the pairs of EGVs, with two lower troughs downstream of each EGV trailing edge. This is of particular interest since rather complex ice shapes have been observed in that region.

compressor↗

NASA NTRS1971

Experimental investigations of a uranium plasma pertinent to a self-sustaining plasma source

The research is pertinent to the realization of a self-sustained fissioning plasma for applications such as nuclear propulsion, closed cycle MHD power generation using a plasma core reactor, and heat engines such as the nuclear piston engine, as well as the direct conversion of fission energy into optical radiation (nuclear pumped lasers). Diagnostic measurement methods and experimental devices simulating plasma core reactor conditions are discussed. Studies on the following topics are considered: (1) ballistic piston compressor (U-235); (2) high pressure uranium plasma (natural uranium); (3) sliding spark discharge (natural uranium); (4) fission fragment interaction (He-3 and U-235); and (5) nuclear pumped lasers (He-3 and U-235).

Schneider, R. T.↗

NASA NTRS1948

Ice Protection of Turbojet Engines by Inertia Separation of Water II : Single-offset-duct System

Investigation of a single-offset-duct system designed to prevent entrance of water into a turbojet engine was conducted on a half-scale nacelle model. An investigation was made to determine ram-pressure recovery and radial velocity profiles at the compressor section and icing characteristics of such a duct system. At a design inlet velocity of 0.77, the maximum ram-pressure recovery attained with effective water-separating inlet was 77 percent, which is considerably less than attainable with a direct-ram inlet. Continuous heating of the accessory-housing surface would be required for inlets that have a small ice storage space.

NOSE SHAPE - DUCTED BODIES↗

NASA NTRS1981

Rub energetics of compressor blade tip seals

The rub mechanics of aircraft gas turbine engine compressor abradable blade tip seals was studied at simulated engine conditions. In 12 statistically planned, instrumented rub tests using single titanium blades and fiber-metal rubstrips the rub velocity, incursion rate, incursion depth, blade thickness, and abradable strength were varied to determine the effects on rub energy, heat split between the blade, rubstrip surface and rub debris, and blade and seal wear. The rub energies were found to be most significantly affected by the incursion rate while rub velocity and blade thickness were of secondary importance. In five additional rub tests using single nickel alloy blades and multiple titanium alloy blades, rub energy and wear effects were found to be similar for titanium and nickel alloy blades while rub energies increased for multiple blades relative to single blade test results.

Laverty, W. F.↗

NASA NTRS1982

PURDU-WINCOF: A computer code for establishing the performance of a fan-compressor unit with water ingestion

A model for predicting the performance of a multi-spool axial-flow compressor with a fan during operation with water ingestion was developed incorporating several two-phase fluid flow effects as follows: (1) ingestion of water, (2) droplet interaction with blades and resulting changes in blade characteristics, (3) redistribution of water and water vapor due to centrifugal action, (4) heat and mass transfer processes, and (5) droplet size adjustment due to mass transfer and mechanical stability considerations. A computer program, called the PURDU-WINCOF code, was generated based on the model utilizing a one-dimensional formulation. An illustrative case serves to show the manner in which the code can be utilized and the nature of the results obtained.

Leonardo, M.↗

NASA NTRS1975

Structural response of fiber composite fan blades

A fiber composite airfoil, typical for high-tip speed compressor applications, is subjected to load conditions anticipated to be encountered in such applications, and its structural response is theoretically investigated. The analysis method used consists of composite mechanics embedded in pre- and post-processors and coupled with NASTRAN. The load conditions examined include thermal due to aerodynamic heating, pressure due to aerodynamic forces, centrifugal, and combinations of these. The various responses investigated include root reactions due to various load conditions, average composite and ply stresses, ply delaminations, and the fundamental modes and the corresponding reactions. The results show that the thermal and pressure stresses are negligible compared to those caused by the centrifugal forces. Also, the core-shell concept for composite blades is an inefficient design (core plies not highly stressed) and appears to be sensitive to interply delaminations. The results are presented in graphical and tabular forms to illustrate the types and amount of data required for such an analysis, and to provide quantitative data of the various responses which can be helpful in designing such composite blades.

Chamis, C. C.↗

NASA NTRS1976

Heat-Transfer Characteristics of Partially Film Cooled Plug Nozzle on a J-85 Afterburning Turbojet Engine

Plug nozzle film cooling data were obtained downstream of a slot located at 42 percent of the total plug length on a J-85 engine. Film cooling reduced the aft end wall temperature as much as 150 K, reduced total pressure loss in the upstream convection cooling passages by 50 percent, and reduced estimated compressor bleed flow requirement by 14 percent compared to an all convectively cooled nozzle. Shock waves along the plug surface strongly influenced temperature distributions on both convection and film cooled portions. The effect was most severe at nozzle pressure ratios below 10 where adverse pressure gradients were most severe.

Nosek, S. M.↗

NASA NTRS1989

Injected Water Augments Cooling In Turboshaft Engine

Report describes experiments in which water injected into compressor-bleed cooling air of aircraft turboshaft engine. Injection of water previously suggested as way to provide additional cooling needed to sustain operation at power levels higher than usual. Involves turbine-inlet temperatures high enough to shorten lives of first-stage high-pressure turbine blades. Latent heat of vaporization of injected water serves as additional heat sink to maintain blades at design operating temperatures during high-power operation.

Biesiadny, Thomas J.↗

NASA NTRS2010

Investing American Recovery and Reinvestment Act Funds to Advance Capability, Reliability, and Performance in NASA Wind Tunnels

The National Aeronautics and Space Administration's (NASA) Aeronautics Test Program (ATP) is implementing five significant ground-based test facility projects across the nation with funding provided by the American Recovery and Reinvestment Act (ARRA). The projects were selected as the best candidates within the constraints of the ARRA and the strategic plan of ATP. They are a combination of much-needed large scale maintenance, reliability, and system upgrades plus creating new test beds for upcoming research programs. The projects are: 1.) Re-activation of a large compressor to provide a second source for compressed air and vacuum to the Unitary Plan Wind Tunnel at the Ames Research Center (ARC) 2.) Addition of high-altitude ice crystal generation at the Glenn Research Center Propulsion Systems Laboratory Test Cell 3, 3.) New refrigeration system and tunnel heat exchanger for the Icing Research Tunnel at the Glenn Research Center, 4.) Technical viability improvements for the National Transonic Facility at the Langley Research Center, and 5.) Modifications to conduct Environmentally Responsible Aviation and Rotorcraft research at the 14 x 22 Subsonic Tunnel at Langley Research Center. The selection rationale, problem statement, and technical solution summary for each project is given here. The benefits and challenges of the ARRA funded projects are discussed. Indirectly, this opportunity provides the advantages of developing experience in NASA's workforce in large projects and maintaining corporate knowledge in that very unique capability. It is envisioned that improved facilities will attract a larger user base and capabilities that are needed for current and future research efforts will offer revenue growth and future operations stability. Several of the chosen projects will maximize wind tunnel reliability and maintainability by using newer, proven technologies in place of older and obsolete equipment and processes. The projects will meet NASA's goal of integrating more efficient, environmentally safer, and less energy consuming hardware and processes into existing tunnel systems. These include Environmental Protection Agency-approved refrigerants, energy efficient motors, and faster, flexible tunnel data systems.

Sydnor, Goerge H.↗

NASA NTRSDate not supplied

Experimental and Numerical Investigation of the NASA High Efficiency Centrifugal Compressor Vaned Stage Geometry and Aerodynamic Performance

Since its inception in the early 2010s, the NASA High Efficiency Centrifugal Compressor (HECC) has been enigmatic for the propulsion research community: experimental data and numerical simulations of the stage have generally not aligned in their quantifications of performance metrics. Typically, the fault for these disagreements is assigned to the numerical simulations as the simulations are models, and models are inherently incomplete representations of the experiment. This “incompleteness” may manifest in assumptions regarding roughness or heat transfer, simplifications of the flow path (i.e., neglecting bleed flows), or the oft-scapegoated turbulence model. In the case of HECC, recent work showed unexpected discrepancies between the intended impeller geometry defined in the design report (termed Design-Intent) and the manufactured impeller used in the experimental campaigns (termed As-Manufactured). That work used numerical simulations to establish that the geometric differences between the Design-Intent and As-Manufactured impellers were significant enough to result in drastically different performance predictions for the HECC vaneless diffuser configuration. The Design-Intent impeller simulations over predicted the performance relative to the experiment, whereas the As-Manufactured simulations better represented the experimental data, both in terms of one-dimensional performance metrics and spanwise flow profiles. This effort expands on that work by examining in detail the geometry and aerodynamic performance of the HECC vaned diffuser configuration. Further differences between the Design-Intent and As-Manufactured geometries have been discovered in the vaned diffuser and exit guide vanes, and these differences are documented herein. The summations of the geometric differences for all of the components were used to create two numerical models of HECC vaned diffuser configuration: the Design-Intent simulations which are generated from the original geometry definitions given in the design report and the As-Manufactured simulations which are the best available representation of the manufactured compressor hardware used in the experimental test campaigns. In congruence with the earlier vaneless diffuser work, the numerical predictions of the Design-Intent choked mass flow rate, total pressure ratio, and efficiency were notably greater than that of the As-Manufactured simulations. To increase confidence in the experimental dataset, measurements from a recent test campaign conducted in 2024 are used to validate the original experimental data acquired from 2012 to 2014 with good repeatability overall, especially considering the passage of time and differences in the data acquisition systems between the test campaigns. Both numerical simulations were then extensively evaluated against the experimental data. The As-Manufactured simulations provided better estimates of the stage performance than the Design-Intent cases in terms of most performance metrics. Nonetheless, more detailed results still show opportunities for improvement. Despite a more accurate representation of the physical hardware, characterization of the impeller work input remains a challenge even for rigorously developed numerical models.

vaned diffuser↗

NASA NTRSDate not supplied

Experimental and Numerical Investigation of the NASA High Efficiency Centrifugal Compressor Vaned Stage Geometry and Aerodynamic Performance

Since its inception in the early 2010s, the NASA High Efficiency Centrifugal Compressor (HECC) has been enigmatic for the propulsion research community: experimental data and numerical simulations of the stage have generally not aligned in their quantifications of performance metrics. Typically, the fault for these disagreements is assigned to the numerical simulations as the simulations are models, and models are inherently incomplete representations of the experiment. This “incompleteness” may manifest in assumptions regarding roughness or heat transfer, simplifications of the flow path (i.e., neglecting bleed flows), or the oft-scapegoated turbulence model. In the case of HECC, recent work showed unexpected discrepancies between the intended impeller geometry defined in the design report (termed Design-Intent) and the manufactured impeller used in the experimental campaigns (termed As-Manufactured). That work used numerical simulations to establish that the geometric differences between the Design-Intent and As-Manufactured impellers were significant enough to result in drastically different performance predictions for the HECC vaneless diffuser configuration. The Design-Intent impeller simulations over predicted the performance relative to the experiment, whereas the As-Manufactured simulations better represented the experimental data, both in terms of one-dimensional performance metrics and spanwise flow profiles. This effort expands on that work by examining in detail the geometry and aerodynamic performance of the HECC vaned diffuser configuration. Further differences between the Design-Intent and As-Manufactured geometries have been discovered in the vaned diffuser and exit guide vanes, and these differences are documented herein. The summations of the geometric differences for all of the components were used to create two numerical models of HECC vaned diffuser configuration: the Design-Intent simulations which are generated from the original geometry definitions given in the design report and the As-Manufactured simulations which are the best available representation of the manufactured compressor hardware used in the experimental test campaigns. In congruence with the earlier vaneless diffuser work, the numerical predictions of the Design-Intent choked mass flow rate, total pressure ratio, and efficiency were notably greater than that of the As-Manufactured simulations. To increase confidence in the experimental dataset, measurements from a recent test campaign conducted in 2024 are used to validate the original experimental data acquired from 2012 to 2014 with good repeatability overall, especially considering the passage of time and differences in the data acquisition systems between the test campaigns. Both numerical simulations were then extensively evaluated against the experimental data. The As-Manufactured simulations provided better estimates of the stage performance than the Design-Intent cases in terms of most performance metrics. Nonetheless, more detailed results still show opportunities for improvement. Despite a more accurate representation of the physical hardware, characterization of the impeller work input remains a challenge even for rigorously developed numerical models.

centrifugal compressor↗

NASA NTRS2000

Numerical Simulation of Non-Rotating and Rotating Coolant Channel Flow Fields

Future generations of ultra high bypass-ratio jet engines will require far higher pressure ratios and operating temperatures than those of current engines. For the foreseeable future, engine materials will not be able to withstand the high temperatures without some form of cooling. In particular the turbine blades, which are under high thermal as well as mechanical loads, must be cooled. Cooling of turbine blades is achieved by bleeding air from the compressor stage of the engine through complicated internal passages in the turbine blades (internal cooling, including jet-impingement cooling) and by bleeding small amounts of air into the boundary layer of the external flow through small discrete holes on the surface of the blade (film cooling and transpiration cooling). The cooling must be done using a minimum amount of air or any increases in efficiency gained through higher operating temperature will be lost due to added load on the compressor stage. Turbine cooling schemes have traditionally been based on extensive empirical data bases, quasi-one-dimensional computational fluid dynamics (CFD) analysis, and trial and error. With improved capabilities of CFD, these traditional methods can be augmented by full three-dimensional simulations of the coolant flow to predict in detail the heat transfer and metal temperatures. Several aspects of turbine coolant flows make such application of CFD difficult, thus a highly effective CFD methodology must be used. First, high resolution of the flow field is required to attain the needed accuracy for heat transfer predictions, making highly efficient flow solvers essential for such computations. Second, the geometries of the flow passages are complicated but must be modeled accurately in order to capture all important details of the flow. This makes grid generation and grid quality important issues. Finally, since coolant flows are turbulent and separated the effects of turbulence must be modeled with a low Reynolds number turbulence model to accurately predict details of heat transfer.

Rigby, David L.↗

NASA NTRS1983

Cascade Joule-Thomson refrigerators

The design criteria for cascade Joule-Thomson refrigerators for cooling in the temperature range from 300 K to 4.2 K were studied. The systems considered use three or four refrigeration stages with various working gases to achieve the low temperatures. Each stage results in cooling to a progressively lower temperature and provides cooling at intermediate temperatures to remove the substantial amount of parasitic heat load encountered in a typical dewar. With careful dewar design considerable cooling can be achieved with moderate gas flows. For many applications, e.g., in the cooling of sensitive sensors, the fact that the refrigerator contains no moving parts and may be remotely located from the gas source is of considerable advantage. A small compressor suitable for providing the gas flows required was constructed.

Tward, E.↗

NASA NTRS1983

New perspectives for advanced automobile diesel engines

Computer simulation results are presented for advanced automobile diesel engine performance. Four critical factors for performance enhancement were identified: (1) part load preheating and exhaust gas energy recovery, (2) fast heat release combustion process, (3) reduction in friction, and (4) air handling system efficiency. Four different technology levels were considered in the analysis. Simulation results are compared in terms of brake specific fuel consumption and vehicle fuel economy in km/liter (miles per gallon). Major critical performance sensitivity areas are: (1) combustion process, (2) expander and compressor efficiency, and (3) part load preheating and compound system. When compared to the state of the art direct injection, cooled, automobile diesel engine, the advanced adiabatic compound engine concept showed the unique potential of doubling the fuel economy. Other important performance criteria such as acceleration, emissions, reliability, durability and multifuel capability are comparable to or better than current passenger car diesel engines.

Tozzi, L.↗

NASA NTRS1981

Advanced Gas Turbine (AGT) powertrain system development for automotive applications

Preliminary layouts were made for the exhaust system, air induction system, and battery installation. Points of interference were identified and resolved by altering either the vehicle or engine designs. An engine general arrangement evolved to meet the vehicle engine compartment constraints while minimizing the duct pressure losses and the heat rejection. A power transfer system (between gasifier and power turbines) was developed to maintain nearly constant temperatures throughout the entire range of engine operation. An advanced four speed automatic transmission was selected to be used with the engine. Performance calculations show improvements in component efficiencies and an increase in fuel economy. A single stage centrifugal compressor design was completed and released for procurement. Gasifier turbine, power turbine, combustor, generator, secondary systems, materials, controls, and transmission development are reported.

Source record↗

NASA NTRS1948

Altitude-Wind-Tunnel investigation of Westinghouse 19B-2, 19B-8, and 19XB-1 jet-propulsion engines V : combustion chamber performance

Pressure losses through the combustion chamber and the combustion efficiency of the 19B-2 and 19B-8 jet-propulsion engines and the combustion efficiency of the 19XB-1 jet-propulsion engine are presented.Data were obtained from an investigation of the complete engine in the NACA Cleveland altitude wind tunnel over a range of simulated altitudes from 5000 to 30,000 feet and tunnel Mach numbers from less than 0.100 to 0.455. The combustion-chamber pressure loss due to friction was higher for the 19B-2 combustion chamber than for the 19B-8. The 19B-2 combustion chamber had a screen of 40-percent open area interposed between the compressor outlet and the combustion-chamber inlet. The screen for the 19B-8 combustion chamber had a 60-percent open area, which except for a small difference in tail-pipe-nozzle outlet area represents the only point of difference between the standard 19B-2 and 19B-8 combustion chambers. The pressure loss due to heat addition to the flowing gases in the combustion chamber was approximately the same for the 19B-2 and 19B-8 configurations. Altitude and tunnel Mach number had no significant effect on the over-all total-pressure loss through the combustion chamber. A decrease in tail-pipe-nozzle outlet area (tail cone out) resulted in a decrease in combustion-chamber total-pressure loss at high engine speeds.

Boyd, Bemrose↗

NASA NTRS1991

Recent progress in inverse methods in France

Given the current level of jet engine performance, improvement of the various turbomachinery components requires the use of advanced methods in aerodynamics, heat transfer, and aeromechanics. In particular, successful blade design can only be achieved via numerical design methods which make it possible to reach optimized solutions in a much shorter time than ever before. Two design methods which are currently being used throughout the French turbomachinery industry to obtain optimized blade geometries are presented. Examples are presented for compressor and turbine applications. The status of these methods as far as improvement and extension to new fields of applications is also reported.

Bry, Pierre-Francois↗