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At least 1,495 records · Page 83
All sourcesNASA NTRSDOE OSTI
NASA NTRS1943

The induction of water to the inlet air as a means of internal cooling in aircraft-engine cylinders

Report presents the results of investigations conducted on a full-scale air-cooled aircraft-engine cylinder of 202-cubic inch displacement to determine the effects of internal cooling by water induction on the maximum permissible power and output of an internal-combustion engine. For a range of fuel-air and water-fuel ratios, the engine inlet pressure was increased until knock was detected aurally, the power was then decreased 7 percent holding the ratios constant. The data indicated that water was a very effective internal coolant, permitting large increases in engine power as limited by either knock or by cylinder temperatures.

Rothrock, Addison M↗
NASA NTRS1956

Theoretical analysis of incompressible flow through a radial-inlet centrifugal impeller at various weight flows

A method for the solution of the incompressible nonviscous flow through a centrifugal impeller, including the inlet region, is presented. Several numerical solutions are obtained for four weight flows through an impeller at one operating speed. These solutions are refined in the leading-edge region. The results are presented in a series of figures showing streamlines and relative velocity contours. A comparison is made with the results obtained by using a rapid approximate method of analysis.

Kramer, James J↗
NASA NTRS1947

Effect of Modifications to Induction System on Altitude Performance of V-1710-93 Engine III : Use of Parabolic Rotating Guide Vanes and NACA Designed Auxiliary-stage Inlet Elbow and Interstage Duct

Bench runs of a modified V-1710-93 engine equipped with a two-stage supercharger, interstage carburetor, aftercooler assembly, and backfire screens have been made at a simulated altitude of 29,000 feet to determine the effect of several induction-system modifications on the engine and supercharger performance. The standard guide vanes on the auxiliary- and engine-stage superchargers were replaced by rotating guide vanes with a parabolic blade profile. The auxiliary-stage inlet elbow and interstage duct were replaced with new units of NACA design. These modifications were made one at a time and data were obtained after each change to determine the effect of each modification. All runs were made at a constant engine speed of 3000 rpm at a simulated altitude of 29,000 feet and all changes in engine power were made by varying the speed of the auxiliary-stage supercharger.

Standahar, Ray M↗
NASA NTRS1993

Schlieren photographs and internal pressure distributions for three-dimensional sidewall-compression scramjet inlets at a Mach number of 6 in CF4

Three-dimensional sidewall-compression scramjet inlets with leading-edge sweeps of 30 deg and 70 deg were tested in the Langley Hypersonic CF4 Tunnel at a Mach number of 6 and a free-stream ratio of specific heats of 1.2. The parametric effects of leading-edge sweep, cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with static pressure orifices distributed on the sidewalls, baseplate, and cowl. Schlieren movies were made of selected tunnel runs for flow visualization of the entrance plane and cowl region. Although these movies could not show the internal flow, the effect of the internal flow on the external flow was evident by way of spillage. The purpose is to provide a preliminary data release for the investigation. The models, facility, and testing methods are described, and the test matrix and a tabulation of tunnel runs are provided. Line plots highlighting the stated parametric effects and a representative set of schlieren photographs are presented without analysis.

Holland, Scott D.↗
NASA NTRS1993

A comparative study of full Navier-Stokes and Reduced Navier-Stokes analyses for separating flows within a diffusing inlet S-duct

A three-dimensional implicit Full Navier-Stokes (FNS) analysis and a 3D Reduced Navier-Stokes (RNS) initial value space marching solution technique has been applied to a class of separate flow problems within a diffusing S-duct configuration characterized as vortex-liftoff. Both Full Navier-Stokes and Reduced Navier-Stokes solution techniques were able to capture the overall flow physics of vortex lift-off, however more consideration must be given to the development of turbulence models for the prediction of the locations of separation and reattachment. This accounts for some of the discrepancies in the prediction of the relevant inlet distortion descriptors, particularly circumferential distortion. The 3D RNS solution technique adequately described the topological structure of flow separation associated with vortex lift-off.

Anderson, B. H.↗
NASA NTRS1993

A comparative study of turbulence models in predicting hypersonic inlet flows

A computational study has been conducted to evaluate the performance of various turbulence models. The NASA P8 inlet, which represents cruise condition of a typical hypersonic air-breathing vehicle, was selected as a test case for the study; the PARC2D code, which solves the full two dimensional Reynolds-averaged Navier-Stokes equations, was used. Results are presented for a total of six versions of zero- and two-equation turbulence models. Zero-equation models tested are the Baldwin-Lomax model, the Thomas model, and a combination of the two. Two-equation models tested are low-Reynolds number models (the Chien model and the Speziale model) and a high-Reynolds number model (the Launder and Spalding model).

Kapoor, Kamlesh↗
NASA NTRS1993

Experimental investigation of generic three-dimensional sidewall-compression scramjet inlets at Mach 6 in tetrafluoromethane

Three-dimensional sidewall-compression scramjet inlets with leading-edge sweeps of 30 deg and 70 deg were tested in the Langley Hypersonic CF4 Tunnel at Mach 6 and with a ratio of specific heats of 1.2. The parametric effects of leading-edge sweep, cowl position, contraction ratio, and Reynolds number were investigated. The models were instrumented with 42 static pressure orifices that were distributed on the sidewalls, base plate, and cowl. Schlieren movies were made of each test for flow visualization of the effects of the internal flow spillage on the external flow field. To obtain an approximate characterization of the flow field, a modification to two-dimensional, inviscid, oblique shock theory was derived to accommodate the three-dimensional effects of leading-edge sweep. This theory qualitatively predicted the reflected shock structure (i.e., sidewall impingement locations) and the observed increase in spillage with increasing leading-edge sweep. The primary effect of moving the cowl forward was capturing the flow that would have otherwise spilled out ahead of the cowl. Increasing the contraction ratio increases the number of internal shock reflections and hence incrementally increases the sidewall pressure distribution. Significant Reynolds number effects were noted over a small range of Reynolds number.

Holland, Scott D.↗
NASA NTRS1994

Inlet distortion in engines on VSTOL aircraft

This report presents the results of a research program on inlet distortion in engines on VSTOL aircraft carried out at the MIT Gas Turbine Laboratory during the period Oct. 1989 - Dec. 1993. The program focused on the development of three dimensional flow computational methodology for predicting the effects of nonuniform flow on the performance of aircraft engines in VSTOL aircraft, the development of a three dimensional instability analysis of flow in multistage axial compressors, and the preliminary applications of these newly developed methodologies for elucidating the effects of flow three dimensionality. The accomplishments of the program are brought out when the current status of predictive capabilities for three dimensional flow instabilities in compressors is assessed against that in 1989.

Tan, Choon S.↗
NASA NTRS1994

A Full Navier-Stokes Analysis of Subsonic Diffuser of a Bifurcated 70/30 Supersonic Inlet for High Speed Civil Transport Application

A full Navier-Stokes analysis was performed to evaluate the performance of the subsonic diffuser of a NASA Lewis Research Center 70/30 mixed-compression bifurcated supersonic inlet for high speed civil transport application. The PARC3D code was used in the present study. The computations were also performed when approximately 2.5 percent of the engine mass flow was allowed to bypass through the engine bypass doors. The computational results were compared with the available experimental data which consisted of detailed Mach number and total pressure distribution along the entire length of the subsonic diffuser. The total pressure recovery, flow distortion, and crossflow velocity at the engine face were also calculated. The computed surface ramp and cowl pressure distributions were compared with experiments. Overall, the computational results compared well with experimental data. The present CFD analysis demonstrated that the bypass flow improves the total pressure recovery and lessens flow distortions at the engine face.

Kapoor, Kamlesh↗
NASA NTRS1994

Analytical investigation of adaptive control of radiated inlet noise from turbofan engines

An analytical model has been developed to predict the resulting far field radiation from a turbofan engine inlet. A feedforward control algorithm was simulated to predict the controlled far field radiation from the destructive combination of fan noise and secondary control sources. Numerical results were developed for two system configurations, with the resulting controlled far field radiation patterns showing varying degrees of attenuation and spillover. With one axial station of twelve control sources and error sensors with equal relative angular positions, nearly global attenuation is achieved. Shifting the angular position of one error sensor resulted in an increase of spillover to the extreme sidelines. The complex control inputs for each configuration was investigated to identify the structure of the wave pattern created by the control sources, giving an indication of performance of the system configuration. It is deduced that the locations of the error sensors and the control source configuration are equally critical to the operation of the active noise control system.

Risi, John D.↗
NASA NTRS1994

Neural net controller for inlet pressure control of rocket engine testing

Many dynamic systems operate in select operating regions, each exhibiting characteristic modes of behavior. It is traditional to employ standard adjustable gain proportional-integral-derivative (PID) loops in such systems where no apriori model information is available. However, for controlling inlet pressure for rocket engine testing, problems in fine tuning, disturbance accommodation, and control gains for new profile operating regions (for research and development) are typically encountered. Because of the capability of capturing I/O peculiarities, using NETS, a back propagation trained neural network is specified. For select operating regions, the neural network controller is simulated to be as robust as the PID controller. For a comparative analysis, the higher order moment neural array (HOMNA) method is used to specify a second neural controller by extracting critical exemplars from the I/O data set. Furthermore, using the critical exemplars from the HOMNA method, a third neural controller is developed using NETS back propagation algorithm. All controllers are benchmarked against each other.

Trevino, Luis C.↗
NASA NTRS1995

Mach 10 computational study of a three-dimensional scramjet inlet flow field

The present work documents the computational results for a combined computational and experimental parametric study of the internal aerodynamics of a generic three-dimensional sidewall-compression scramjet inlet configuration at Mach 10. The three-dimensional Navier-Stokes code SCRAMIN was chosen for the computational portion of the study because it uses a well-known and well-proven numerical scheme and has shown favorable comparison with experiment at Mach numbers between 2 and 6. One advantage of CFD was that it provided flow field data for a detailed examination of the internal flow characteristics in addition to the surface properties. The experimental test matrix at mach 10 included three geometric contraction ratios (3, 5, and 9), three Reynolds numbers (0.55 x 10(exp 6) per foot, 1.14 x 10(exp 6) per foot, and 2.15 x 10(exp 6) per foot), and three cowl positions (at the throat and two forward positions). Computational data for two of these configurations (the contraction ratio of 3, Re = 2.15 x 10 (exp 6) per foot, at two cowl positions) are presented along with a detailed analysis of the flow interactions in successive computational planes.

Holland, Scott D.↗
NASA NTRS1995

Mach 10 computational study of a three-dimensional scramjet inlet flow field

The present work documents the computational results for a combined computational and experimental parametric study of the internal aerodynamics of a generic three-dimensional sidewall-compression scramjet inlet configuration at Mach 10. The three-dimensional Navier-Stokes code SCRAMIN was chosen for the computational portion of the study because it uses a well-known and well-proven numerical scheme and has shown favorable comparison with experiment at Mach numbers between 2 and 6. One advantage of CFD was that it provided flow field data for a detailed examination of the internal flow characteristics in addition to the surface properties. The experimental test matrix at Mach 10 included three geometric contraction ratios (3, 5, and 9), three Reynolds numbers (0.55 x 10(exp 6) per foot, 1.14 x 10(exp 6) per foot, and 2.15 x 10(exp 6) per foot), and three cowl positions (at the throat and two forward positions). Computational data for two of these configurations (the contraction ratio of 3, Re = 2.15 x 10(exp 6) per foot, at two cowl positions) are presented along with a detailed analysis of the flow interactions in successive computational planes.

Holland, Scott D.↗
NASA NTRS1995

Minimum fan turbine inlet temperature mode evaluation

Measured reductions in turbine temperature which resulted from the application of the F-15 performance seeking control (PSC) minimum fan turbine inlet temperature (FTIT) mode during the dual-engine test phase is presented as a function of net propulsive force and flight condition. Data were collected at altitudes of 30,000 and 45,000 feet at military and partial afterburning power settings. The FTIT reductions for the supersonic tests are less than at subsonic Mach numbers because of the increased modeling and control complexity. In addition, the propulsion system was designed to be optimized at the mid supersonic Mach number range. Subsonically at military power, FTIT reductions were above 70 R for either the left or right engines, and repeatable for the right engine. At partial afterburner and supersonic conditions, the level of FTIT reductions were at least 25 R and as much as 55 R. Considering that the turbine operates at or very near its temperature limit at these high power settings, these seemingly small temperature reductions may significantly lengthen the life of the turbine. In general, the minimum FTIT mode has performed well, demonstrating significant temperature reductions at military and partial afterburner power. Decreases of over 100 R at cruise flight conditions were identified. Temperature reductions of this magnitude could significantly extend turbine life and reduce replacement costs.

Orme, John S.↗
NASA NTRS1996

A Pre-Mixed Shock-Induced-Combustion Approach to Inlet and Combustor Design for Hypersonic Applications

The need for efficient access to space has created interest in airbreathing propulsion as a means of achieving that goal. The NASP program explored a single-stage-to-orbit approach which could require scramjet airbreathing propulsion out to Mach 16 to 20. Recent interest in global access could require hypersonic cruise engines operating efficiently in the Mach 10 to 12 speed range. A common requirement of both these types of propulsion systems is that they would have to be fully integrated with the aero configuration so that the forebody becomes a part of the external compression inlet and the nozzle expansion is completed on the vehicle aftbody.

Weidner, John P.↗
NASA NTRS1996

Dynamic and Transient Performance of Turbofan/Turboshaft Convertible Engine With Variable Inlet Guide Vanes

A convertible engine called the CEST TF34, using the variable inlet guide vane method of power change, was tested on an outdoor stand at the NASA Lewis Research Center with a waterbrake dynamometer for the shaft load. A new digital electronic system, in conjunction with a modified standard TF34 hydromechanical fuel control, kept engine operation stable and safely within limits. All planned testing was completed successfully. Steady-state performance and acoustic characteristics were reported previously and are referenced. This report presents results of transient and dynamic tests. The transient tests measured engine response to several rapid changes in thrust and torque commands at constant fan (shaft) speed. Limited results from dynamic tests using the pseudorandom binary noise technique are also presented. Performance of the waterbrake dynamometer is discussed in an appendix.

McArdle, Jack G.↗
NASA NTRS1959

Effect on Inlet Performance of a Cowl Visor and an Internal-Contraction Cowl for Drag Reduction at Mach Numbers 3.07 and 1.89

Two methods for reducing the external cowl angle, and hence the cowl pressure drag, were investigated on a two-dimensional model. One method used at both on- and off-design Mach numbers was the addition of a cowl visor that had the inner surface parallel to the free stream at 0 deg angle of attack. The other method investigated consisted in replacing the original cowl by a flatter cowl that also provided internal contraction. Both the visor and the internal-contraction cowl reduced the cowl pressure drag 64 percent or more. The visor had little effect on inlet performance at the design Mach number except to reduce the stability range slightly. At off-design, the visor caused an increase in critical pressure recovery.

Gertsma, Laurence W.↗