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At least 253 records · Page 14

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NASA NTRS1959

Analytical Investigation of the Significance of Turbine-Inlet Temperature in High-Energy Rocket Turbodrive Applications

The effect of turbine-inlet temperature on rocket gross weight was investigated for three high-energy long-range rockets in order to explore the desirability of turbine cooling in rocket turbodrive applications. Temperatures above and below the maximum that is permissible in uncooled turbines were included. Turbine bleed rate and stage number were considered as independent variables. The gross weight of the hydrogen-reactor system was more sensitive to changes in turbine-inlet temperature than either the hydrogen-oxygen or the hydrogen-fluorine systems. Gross weight of the hydrogen-reactor system could be reduced by 2.6 percent by the use of cooling and a turbine-inlet temperature of 3000 R. The reductions in the first stages of the hydrogen-oxygen and hydrogen-fluorine systems were 0.7 and 0.2 percent, respectively. The effect of turbine-inlet temperature on rocket gross weight was small because the resulting turbine weight and bleed rate variations were small. Since these small gains must be balanced against considerations of greater cost, weight, and complexity as well as lessened reliability with a system utilizing a cooled turbine, none of the systems investigated showed gains warranting the use of turbine cooling.

Rohlik, Harold E.↗

NASA NTRS1997

Improved Hypersonic Inlet Performance Using Validated Strut Compression Designs

Aerojet is currently executing two Strutjet propulsion contracts: one a Rocket Based Combined Cycle (RBCC) engine for a NASA-Marshall Space Flight Center (MSFC) Advanced Reusable Transportation Technology (ARTT) program, the second a Dual Mode Ram/Scramjet engine for a USAF Wright Laboratories Storable Fuel Scramjet Flow Path Concepts program. The engines employed in both programs operate at supersonic and low hypersonic speeds and use inlets employing forebody external and sidewall compression. Aerojet has developed and validated a successful design methodology applicable to these inlet types. Design features include an integrated vehicle forebody, external side compression struts, strut sidewall and throat bleed, a throat shock trap, and variable geometry internal contraction. Computation Fluid Dynamic (CFD) predictions and test data show these inlets allow substantially increased flow turning angles over other designs. These increased flow turning angles allow shorter and lighter engines than current designs, which in turn enables higher performing vehicles with broad operating characteristics. This paper describes the designs of two different inlets evaluated by the NASA-MSFC and USAF programs, discusses the results of wind tunnel tests performed by NASA-Lewis Research Center, and provides correlations of test data with CFD predictions. Parameters of interest include low Mach number starting capability, start sensitivity as a function of back pressure at various contraction ratios, flow turning angles, strut and throat bleed effects, and pressure recovery at various Mach numbers.

Bulman, M. J.↗

NASA NTRS1999

Triboelectric Charging in Simulated Mars Environment

Triboelectric charging of nonconducting materials followed by sudden electrostatic discharge (ESD) can damage electronic equipment and become ignition hazard to combustible materials. Mars atmosphere has near zero humidity and therefore natural charge bleeding to surroundings is anticipated to be limited. Potential mitigation of ESD problems has been conjectured based upon strong extraterrestrial radiation on Mars compared to earth. A hypothesis was formulated that ESD problem is less significant in simulated Mars condition since strong radiation and presence of argon will generate an ionized environment; this will be conducive to rapid bleeding of static charge into the surroundings.

Lee, R.↗

NASA NTRS2000

On Flowfield Periodicity in the NASA Transonic Flutter Cascade: Experimental Study - Part 1

An extensive study to improve flow uniformity and periodicity in the NASA Transonic Flutter Cascade is presented here. The results are reported in two independent parts dealing with the experimental approach and the analytical approach. The first part, the Experimental Study, focuses first on the data sets acquired in this facility in the past and explains several discrepancies, particularly the questions of actual flow incidence and cascade back pressure levels. Next, available means for control and modifications of the cascade flowfield, boundary layer bleed and tailboard settings are presented in detail. This is followed by experimental data sets acquired in modified test facility configurations that were based on analytical predictions of the cascade flowfield. Finally, several important conclusions about improving the cascade flowfield uniformity and blade load periodicity are summarized. The important conclusions are: (1) boundary layer bleed does not improve the cascade flow periodicity; (2) tunnel wall contours must be carefully matched to the expected shape of cascade streamlines; (3) actual flow incidence for each cascade configuration rather must be measured instead of relying on the tunnel geometry; and (4) the current cascade configuration exhibits a very high blade load uniformity over six blades from blade #2 to blade #7, and the facility is now ready for unsteady pressure data acquisition.

Lepicovsky, J.↗

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 NTRS1947

Acceleration Characteristics of R-3350 Engine Equipped with NACA Injection Impeller

Qualitative investigations have shown that use of the NACA injection impeller with the R-3350 engine increases the inertia of the fuel-injection system and, when the standard fuel-metering system is used, this increase in inertia results in poor engine acceleration characteristics. This investigation was therefore undertaken to determine whether satisfactory acceleration characteristics of the engine equipped with the injection impeller could be obtained by simple modifications to the fuel-monitoring system. The engine was operated with two types of carburetor; namely, a hydraulic-metering carburetor incorporating a vacuum-operated accelerating pump and a direct-metering carburetor having a throttle-actuated accelerating pump. The vacuum-operated accelerating pump of the hydraulic-metering carburetor was modified to produce satisfactory accelerations by supplementing the standard air chamber with an additional 75-cubic spring. The throttle-actuated accelerating pump of the direct-metering carburetor was modified to produce satisfactory accelerations by replacing the standard 0.028-inch-diameter bleed in the load-compensator balance line with a smaller bleed of 0.0225-inch diameter. The results of this investigation indicated that both carburetors can be easily modified to produce satisfactory acceleration characteristics of the engine and no definite choice between the types of carburetor and accelerating pump can be made. Use of the direct-metering carburetor, however, probably resulted in better fuel distribution to the cylinders during the acceleration period and reduced the backfire hazard because all the fuel is introduced through the injection impeller.

Hickel, Robert O.↗

NASA NTRS1998

Inlet Performance Analysis Code Developed

The design characteristics of an inlet very much depend on whether the inlet is to be flown at subsonic, supersonic, or hypersonic speed. Whichever the case, the primary function of an inlet is to deliver free-stream air to the engine face at the highest stagnation pressure possible and with the lowest possible variation in both stagnation pressure and temperature. At high speeds, this is achieved by a system of oblique and/or normal shock waves, and possibly some isentropic compression. For both subsonic and supersonic flight, current design practice indicates that the inlet should deliver the air to the engine face at approximately Mach 0.45. As a result, even for flight in the high subsonic regime, the inlet must retard (or diffuse) the air substantially. Second, the design of an inlet is influenced largely by the compromise between high performance and low weight. This compromise involves tradeoffs between the mission requirements, flight trajectory, airframe aerodynamics, engine performance, and weight-all of which, in turn, influence each other. Therefore, to study the effects of some of these influential factors, the Propulsion System Analysis Office of the NASA Lewis Research Center developed the Inlet Performance Analysis Code (IPAC). This code uses oblique shock and Prandtl-Meyer expansion theory to predict inlet performance. It can be used to predict performance for a given inlet geometric design such as pitot, axisymmetric, and two-dimensional. IPAC also can be used to design preliminary inlet systems and to make subsequent performance analyses. It computes the total pressure, the recovery, the airflow, and the drag coefficients. The pressure recovery includes losses associated with normal and oblique shocks, internal and external friction, the sharp lip, and diffuser components. Flow rate includes captured, engine, spillage, bleed, and bypass flows. The aerodynamic drag calculation includes drags associated with spillage, cowl lip suction, wave, bleed, and bypass.

Jules, Kenol↗

NASA NTRS2004

Air Separation Using Hollow Fiber Membranes

The NASA Glenn Research Center in partnership with the Ohio Aerospace Institute provides internship programs for high school and college students in the areas of science, engineering, professional administrative, and other technical areas. During the summer of 2004, I worked with Dr. Clarence T. Chang at NASA Glenn Research Center s combustion branch on air separation using hollow fiber membrane technology. . In light of the accident of Trans World Airline s flight 800, FAA has mandated that a suitable solution be created to prevent the ignition of fuel tanks in aircrafts. In order for any type of fuel to ignite, three important things are needed: fuel vapor, oxygen, and an energy source. Two different ways to make fuel tanks less likely to ignite are reformulating the fuel to obtain a lower vapor pressure for the fuel and or using an On Board Inert Gas Generating System (OBIGGS) to inert the Central Wing Tank. goal is to accomplish the mission, which means that the Air Separation Module (ASM) tends to be bulky and heavy. The primary goal for commercial aviation companies is to transport as much as they can with the least amount of cost and fuel per person, therefore the ASM must be compact and light as possible. The plan is to take bleed air from the aircraft s engines to pass air through a filter first to remove particulates and then pass the air through the ASM containing hollow fiber membranes. In the lab, there will be a heating element provided to simulate the temperature of the bleed air that will be entering the ASM and analysis of the separated air will be analyzed by a Gas Chromatograph/Mass Spectrometer (GC/MS). The GUMS will separate the different compounds in the exit streams of the ASM and provide information on the performance of hollow fiber membranes. Hopefully I can develop ways to improve efficiency of the ASM. different types of jet fuel were analyzed and data was well represented on SAE Paper 982485. Data consisted of the concentrations of over 300 different hydrocarbons commonly found in JP- 8, Jet A, and JP-5 fuels. I researched the major hydrocarbons that has a concentration of greater than 50 parts per million and found the vapor pressure data coefficients for a specific temperature range. The coefficients were applied to Antoine s Equation and Riedel's Equation to calculate the vapor pressures for that specific hydrocarbon in the specific temperature range. With the vapor pressure data scientists can formulate a fuel composition that has a lower vapor pressure profile, therefore making jet fuels less flammable. work, learn how to operate and examine the data from Gas Chromatograph and Mass Spectrometer, and develop new ways in applying hollow fiber membrane technology to other areas of environmental engineering. The United States military currently uses air separation technology and their primary The other side of making air travel safer is to reformulate the fuel. Analyses of three My goal this summer is to learn about hollow fiber membrane technologies and how they

Huang, Stephen E.↗

NASA NTRS2001

Compressor Stall Recovery Through Tip Injection Assessed

Aerodynamic stability is a fundamental limit in the compressor design process. The development of robust techniques for increasing stability has several benefits: enabling higher loading and fewer blades, increasing safety throughout a mission, increasing tolerance to stage mismatch during part-speed operation and speed transients, and providing an opportunity to match stages at the compressor maximum efficiency point, thus reducing fuel burn. Mass injection upstream of the tip of a high-speed axial compressor rotor is a stability enhancement approach known to be effective in suppressing stall in tip-critical rotors if the injection is activated before stall occurs. This approach to stall suppression requires that a reliable stall warning system be available. Tests have recently been performed to assess whether steady injection can also be used to recover from fully developed stall. If mass injection is effective in recovering from stall quickly enough to avoid structural damage or loss of engine power, then a stall warning system may not be required. The stall recovery tests were performed on a transonic compressor rotor at its design tip speed of 1475 ft/sec using four injectors evenly spaced around the compressor case upstream of the rotor. The injectors were connected to an external air source. In an actual engine application, the injected air would be supplied with compressor bleed air. The injectors were isolated from the air source by a fast-acting butterfly valve. With the injectors turned off, the compressor was throttled into stall. Air injection was then activated with no change in throttle setting by opening the butterfly valve. The compressor recovered from stall at a fixed throttle setting with the aid of tip injection. The unsteady operating characteristic of the rotor was measured during these tests using high-response pressure sensors located upstream and downstream of the rotor. The figure shows the results, where the unsteady pressure and mass flow are superimposed on the steady operating characteristic. The total injected mass flow was equal to 1.3 percent of the compressor flow. The solid line with no solid squares on it denotes the operating point during the beginning of throttle closure and the initial drop into stall. The gray traces denote the operating point during an additional throttle closure that occurred over the next 1200 rotor revolutions (4 sec). The dashed line denotes the recovery from stall that occurred during 90 rotor revolutions (0.3 sec) after the injectors were activated with no change in throttle setting. Tip injection not only recovers the compressor from stall, but also restores the compressor to its pre-stall level of pressure rise. In contrast, standard stall recovery schemes such as compressor bleed, stator vane actuation, or engine throttle modulation result in a loss of pressure rise across the compressor, which results in a loss of engine power.

Suder, Ken L.↗

NASA NTRS2006

Optimal Control of Shock Wave Turbulent Boundary Layer Interactions Using Micro-Array Actuation

The intent of this study on micro-array flow control is to demonstrate the viability and economy of Response Surface Methodology (RSM) to determine optimal designs of micro-array actuation for controlling the shock wave turbulent boundary layer interactions within supersonic inlets and compare these concepts to conventional bleed performance. The term micro-array refers to micro-actuator arrays which have heights of 25 to 40 percent of the undisturbed supersonic boundary layer thickness. This study covers optimal control of shock wave turbulent boundary layer interactions using standard micro-vane, tapered micro-vane, and standard micro-ramp arrays at a free stream Mach number of 2.0. The effectiveness of the three micro-array devices was tested using a shock pressure rise induced by the 10 shock generator, which was sufficiently strong as to separate the turbulent supersonic boundary layer. The overall design purpose of the micro-arrays was to alter the properties of the supersonic boundary layer by introducing a cascade of counter-rotating micro-vortices in the near wall region. In this manner, the impact of the shock wave boundary layer (SWBL) interaction on the main flow field was minimized without boundary bleed.

Anderson, Bernhard H.↗

NASA NTRS1954

Preliminary Data on the Effects of Inlet Pressure Distortions on the J57-P-1 Turbojet Engine

An investigation to determine the steady-state and surge characteristics of the J57-P-1 two-spool turbojet engine with various inlet air-flow distortions was conducted in the altitude wind tunnel at the NACA Lewis laboratory. Along with a uniform inlet total-pressure distribution, one circumferential and three radial pressure distortions were investigated. Data were obtained over a complete range of compressor speeds both with and without intercompressor air bleed at a flight Mach number of 0.8 and at altitudes of 35,000 and 50,000 feet. Total-pressure distortions of the magnitudes investigated had very little effect on the steady-state operating line for either the outer or inner compressor. The small radial distortions investigated also had engine over that obtained with the uniform inlet pressure distribution. The circumferential distortion, however, raised the minimum speed at which the engine could operate without encountering surge when the intercompressor bleeds were closed. This increase in minimum speed resulted in a substantial reduction in the operable speed range accompanied by a reduction in the altitude operating limit.

Wallner, Lewis E.↗

NASA NTRS2012

Development and Applications of a Stage Stacking Procedure

The preliminary design of multistage axial compressors in gas turbine engines is typically accomplished with mean-line methods. These methods, which rely on empirical correlations, estimate compressor performance well near the design point, but may become less reliable off-design. For land-based applications of gas turbine engines, off-design performance estimates are becoming increasingly important, as turbine plant operators desire peaking or load-following capabilities and hot-day operability. The current work develops a one-dimensional stage stacking procedure, including a newly defined blockage term, which is used to estimate the off-design performance and operability range of a 13-stage axial compressor used in a power generating gas turbine engine. The new blockage term is defined to give mathematical closure on static pressure, and values of blockage are shown to collapse to curves as a function of stage inlet flow coefficient and corrected shaft speed. In addition to these blockage curves, the stage stacking procedure utilizes stage characteristics of ideal work coefficient and adiabatic efficiency. These curves are constructed using flow information extracted from computational fluid dynamics (CFD) simulations of groups of stages within the compressor. Performance estimates resulting from the stage stacking procedure are shown to match the results of CFD simulations of the entire compressor to within 1.6% in overall total pressure ratio and within 0.3 points in overall adiabatic efficiency. Utility of the stage stacking procedure is demonstrated by estimation of the minimum corrected speed which allows stable operation of the compressor. Further utility of the stage stacking procedure is demonstrated with a bleed sensitivity study, which estimates a bleed schedule to expand the compressors operating range.

Kulkarni, Sameer↗

NASA NTRS2014

Thermal Analysis of Cryogenic Hydrogen Liquid Separator

During launch for the new Space Launch System (SLS) liquid hydrogen is bleed through the engines during replenish, pre-press, and extended pre-press to condition the engines prior to launch. The predicted bleed flow rates are larger than for the shuttle program. A consequence of the increased flow rates is having liquif hydrogen in the vent system, which the facilities was never designed to handle. To remedy the problem a liquid separator is being designed in the system to accumulated the liquid propellant and protect the facility flare stack (which can only handle gas). The attached document is a presentation of the current thermalfluid analysis performed for the separator and will be presented at the Thermal and Fluid Analysis Workshop (NASA workshop) next week in Cleveland, Ohio.

liquid separator↗

NASA NTRS2002

LEWICE 2.2 Capabilities and Thermal Validation

A computational model of bleed air anti-icing and electrothermal de-icing have been added to the LEWICE 2.0 software by integrating the capabilities of two previous programs, ANTICE and LEWICE/ Thermal. This combined model has been released as LEWICE version 2.2. Several advancements have also been added to the previous capabilities of each module. This report will present the capabilities of the software package and provide results for both bleed air and electrothermal cases. A comprehensive validation effort has also been performed to compare the predictions to an existing electrothermal database. A quantitative comparison shows that for deicing cases, the average difference is 9.4 F (26%) compared to 3 F for the experimental data while for evaporative cases the average difference is 2 F (32%) compared to an experimental error of 4 F.

Wright, William B.↗

NASA NTRSDate not supplied

Some Design Considerations for Two-Dimensional, Mixed-Compression Inlets for Mach 2.5 to 3.5.

A computational design study was performed to explore design factors for two-dimensional, mixed-compression inlets designed for cruise conditions at Mach 2.5, 2.75, 3.0, 3.25, and 3.5 and their influence on the aerodynamic performance of the inlets. The design factors included the split of external and internal compression, capture aspect ratio, throat section and subsonic diffuser geometry, and location and rates of porous bleed regions. The performance of the inlet was measured by the total pressure recovery, cowl wave drag, bleed flow rates, and total pressure distortion at the engine face. The results provide useful information for the design of inlets for future supersonic aircraft.

Supersonic Inlet↗

NASA NTRSDate not supplied

Scoping of an Air Supply Configuration for AFC on a Commercial Transport Airplane High Lift System

A scoping study is being conducted on the air supply fora microjet-based active flow control (AFC) system on a twin-turbofan commercial transport airplane. Microjets provide circulation control using small surface-normal pneumatic jets located near the trailing edge of a lifting surface such as a wing or flap. When located on the pressure side of the lifting surface they increase the lift, and when located on the suction side they decrease lift. In this study, microjets are considered for installation in the flaps of the high-lift version of the Common ResearchModel (CRM-HL). Two different architectures to supply the air for the microjets are considered: (1) bleed air from the airplane’s auxiliary power unit (APU) plus ram air, and (2) engine fan bleed air plus ram air. A model based on the 1D compressible flow equations is applied to analyze the air supply system architectures and predict the microjet flow rate with the resulting airplane performance changes based on Reynolds-averaged Navier-Stokes modeling of microjets on the CRM-HL. The results of this scoping study are encouraging in that pressurized air from the APU or the engine fan can be used to entrain ram air and thereby increase the AFC mass flow rate to achieve effective lift control and airplane performance enhancement during takeoff and landing.

AATT↗

NASA NTRS1957

Performance Comparison at Mach Numbers 1.8 and 2.0 of Full Scale and Quarter Scale Translating-Spike Inlets

The performance of a full-scale translating-spike inlet was obtained at Mach numbers of 1.8 and 2.0 and at angles of attach from 0 deg to 6 deg. Comparisons were made between the full-scale production inlet configuration and a geometrically similar quarter-scale model. The inlet pressure-recovery, cowl pressure-distribution, and compressor-face distortion characteristics of the full-scale inlet agreed fairly well with the quarter-scale results. In addition, the results indicated that bleeding around the periphery ahead of the compressor-face station improved pressure recovery and compressor-face distortion, especially at angle of attack.

Anderson, B. H.↗

NASA NTRS1965

Miniature bearings lubricated by sonic dispersion method

Evenly distributing a monomolecular film over the balls and tracks of miniature precision ball bearings by sonic dispersion results in precise lubrication which prevents lubricant bleed out to adjacent components. Varying the lubricant to solvent ratio of the mixture causes varying lubricant coating thicknesses.

Source record↗