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At least 1,009 records · Page 56

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NASA NTRS1988

Effect of spatial inlet temperature and pressure distortion on turbofan engine stability

The effects of circumferential and radial inlet temperature distortion, circumferential pressure distortion, and combined temperature and pressure distortion on the stability of an advanced turbofan engine were investigated experimentally at simulated altitude conditions. With circumferential and radial inlet temperature distortion, a flow instability generated by the fan operating near stall caused the high-pressure compressor to surge at, or near, the same time as the fan. The effect of combined distortion was dependent on the relative location of the high-temperature and low-pressure regions; high-pressure compressor stalls occurred when the regions coincided, and fan stalls occurred with the regions separated.

Mehalic, Charles M.↗

NASA NTRS1988

Low pressure gas flow analysis through an effusive inlet using mass spectrometry

A mass spectrometric method for analyzing flow past and through an effusive inlet designed for use on the tethered satellite and other entering vehicles is discussed. Source stream concentrations of species in a gaseous mixture are determined using a calibration of measured mass spectral intensities versus source stream pressure for standard gas mixtures and pure gases. Concentrations are shown to be accurate within experimental error. Theoretical explanations for observed mass discrimination effects as they relate to the various flow situations in the effusive inlet and the experimental apparatus are discussed.

Brown, David R.↗

NASA NTRS1986

The effects of inlet turbulence and rotor/stator interactions on the aerodynamics and heat transfer of a large-scale rotating turbine model. Volume 3: Heat transfer data tabulation 65 percent axial spacing

This is Volume 3 - Heat Transfer Data Tabulation (65 percent Axial Spacing) of a combined experimental and analytical program which was conducted to examine the effects of inlet turbulence on airfoil heat transfer. The experimental portion of the study was conducted in a large-scale (approximately 5X engine), ambient temperature, rotating turbine model configured in both single stage and stage-and-a-half arrangements. Heat transfer measurements were obtained using low-conductivity airfoils with miniature thermocouples welded to a thin, electrically heated surface skin. Heat transfer data were acquired for various combinations of low or high inlet turbulence intensity, flow coefficient, first-stator/rotor axial spacing, Reynolds number and relative circumferential position of the first and second stators.

Dring, R. P.↗

NASA NTRS1988

Numerical simulation of hypersonic inlet flows with equilibrium or finite rate chemistry

An efficient numerical program incorporated with comprehensive high temperature gas property models has been developed to simulate hypersonic inlet flows. The computer program employs an implicit lower-upper time marching scheme to solve the two-dimensional Navier-Stokes equations with variable thermodynamic and transport properties. Both finite-rate and local-equilibrium approaches are adopted in the chemical reaction model for dissociation and ionization of the inlet air. In the finite rate approach, eleven species equations coupled with fluid dynamic equations are solved simultaneously. In the local-equilibrium approach, instead of solving species equations, an efficient chemical equilibrium package has been developed and incorporated into the flow code to obtain chemical compositions directly. Gas properties for the reaction products species are calculated by methods of statistical mechanics and fit to a polynomial form for C(p). In the present study, since the chemical reaction time is comparable to the flow residence time, the local-equilibrium model underpredicts the temperature in the shock layer. Significant differences of predicted chemical compositions in shock layer between finite rate and local-equilibrium approaches have been observed.

Yu, Sheng-Tao↗

NASA NTRS1987

Development of an effusive inlet for mass spectrometric gas analysis of hypersonic boundary layer

The use of a microchannel plate (MCP) as a mass spectrometer inlet device to allow nonintrusive sampling of flight vehicle boundary layers is investigated. Two possible configurations for mounting the inlet are studied: (1) flow coaxial with the channels; and (2) flow perpendicular to the channel axis. The test gases are pure Kr; pure Ne; and a mixture of 10 pct Kr, 10 pct Ne, and 80 pct N2. The pressure ranges studied vary from 500 to 10 microns. A mass discrimination at the quadrupole mass spectrometer is observed, indicating an enrichment in the heavier gas. Possible explanations for this enrichment are discussed. It is shown that an MCP is capable of acting as a nonintrusive sampling device. Further work that will enable quantitative determination of the species at the surface is discussed.

Brown, Kenneth G.↗

NASA NTRS1988

Effect of spatial inlet temperature and pressure distortion on turbofan engine stability

Mehalic, Charles M.↗

NASA NTRS1984

Experimental results for a two-dimensional supersonic inlet used as a thrust deflecting nozzle

Nearly all supersonic V/STOL aircraft concepts are dependent on the thrust deflecting capability of a nozzle. In one unique concept, referred to as the reverse flow dual fan, not only is there a thrust deflecting nozzle for the fan and core engine exit flow, but because of the way the propulsion system operates during vertical takeoff and landing, the supersonic inlet is also used as a thrust deflecting nozzle. This paper presents results of an experimental study to evaluate the performance of a supersonic inlet used as a thrust deflecting nozzle for this reverse flow dual fan concept. Results are presented in terms of nozzle thrust coefficient and thrust vector angle for a number of inlet/nozzle configurations. Flow visualization and nozzle exit flow survey results are also shown.

Johns, Albert L.↗

NASA NTRS1989

Mach 5 inlet CFD and experimental results

An experimental research program was conducted in the NASA Lewis Research Center 10 x 10 ft supersonic wind tunnel. The 2-D inlet model was designed to study the Mach 3.0 to 5.0 speed range for an over-under turbojet plus ramjet propulsion system. The model was extensively instrumented to provide both analytical code validation data as well as inlet performance information. Support studies for the program include flow field predictions with both 3-D parabolized Navier-Stokes (PNS) and 3-D full Navier-Stokes (FNS) analytical codes. Analytical predictions and experimental results are compared.

Weir, Lois J.↗

NASA NTRS1989

Space Shuttle Main Engine structural analysis and data reduction/evaluation. Volume 5: Main Injector LOX Inlet analysis

The analysis performed on the Main Injector LOX Inlet Assembly located on the Space Shuttle Main Engine is summarized. An ANSYS finite element model of the inlet assemably was built and executed. Static stress analysis was also performed.

Violett, Rebeca S.↗

NASA NTRS1988

Inlet development for the NFAC 80- by 120-foot indraft wind tunnel

This paper describes the inlet design requirements of the National Full-Scale Aerodynamic Complex (NFAC), a closed-loop 4-by-80 foot wind tunnel and indraft 80-by-120 foot wind tunnel. The requirements are based upon desired test section flow quality, atmospheric wind isolation, and noise attenuation. An aerodynamic design study of the inlet treatment is described along with the analysis method and the small-scale testing program used. Measurements taken in the full-scale facility are presented, and the flow quality is summarized. The results are compared with theoretical predictions and small-scale model results.

Van Aken, Johannes M.↗

NASA NTRS1989

Thermal analysis of engine inlet anti-icing systems

A hot air anti-icing system of a gas turbine engine inlet is analyzed numerically. A three-dimensional potential flow code, which accounts for compressibility effects, is used to determine the flowfield in and around the inlet. A particle trajectory code is developed using a local linearization technique. The trajectory code is used to calculate local water impingement rates. Energy balances are performed on both the surface runback water and the metallic skin to determine their temperature distributions. A variety of test cases are considered in order to validate the various numerical components of the process as well as to demonstrate the procedure.

Keith, Theo G., Jr.↗

NASA NTRS1989

Mach 5 inlet CFD and experimental results

An experimental research program was conducted in the NASA Lewis Research Center 10 ft. by 10 ft. supersonic wind tunnel. The two-dimensional inlet model was designed to study the Mach 3.0 to 5.0 speed range for an 'over-under' turbojet plus ramjet propulsion system. The model was extensively instrumented to provide both analytical code validation data as well as inlet performance information. Support studies for the program include flow field predictions with both three-dimensional parabolized Navier-Stokes (PNS) and three-dimensional full Navier-Stokes (FNS) analytical codes. Analytical predictions and experimental results are compared.

Weir, Lois J.↗

NASA NTRS1990

Multiple-Inlet/Single-Outlet Orifice Plate

Orifice plate for control of flow has multiple slanted inlets leading to single outlet. Multiple-inlet/single-outlet orifice less susceptible to blockage than single hole drilled through plate perpendicular to its surface. Easily calibrated for various flow rates and fits in place of simple orifice plate with no other modifications.

Godown, Robert L.↗

NASA NTRS1989

Engine inlet distortion in a 9.2 percent scale vectored thrust STOVL model in ground effect

Advanced Short Takeoff/Vertical Landing (STOVL) aircraft which can operate from remote locations, damaged runways, and small air capable ships are being pursued for deployment around the turn of the century. To achieve this goal, NASA Lewis Research Center, McDonnell Douglas Aircraft, and DARPA defined a cooperative program for testing in the NASA Lewis 9- by 15-foot low speed wind tunnel (LSWT) to establish a database for hot gas ingestion, one of the technologies critical to STOVL. Results are presented which show the engine inlet distortions (both temperature and pressure) in a 9.2 percent scale vectored thrust STOVL model in ground effects. Results are shown for the forward nozzle splay angles of 0 degrees, -6 degrees, and 18 degrees. The model support system had 4 degrees of freedom, heated high pressure air for nozzle flow, and a suction system exhaust for inlet flow. The headwind (freestream) velocity was varied from 8 to 23 knots.

Johns, Albert L.↗

NASA NTRS1989

Engine inlet distortion in a 9.2 percent scaled vectored thrust STOVL model in ground effect

Advanced Short Takeoff/Vertical Landing (STOVL) aircraft which can operate from remote locations, damaged runways, and small air capable ships are being pursued for deployment around the turn of the century. To achieve this goal, a cooperative program has been defined for testing in the NASA Lewis 9- by 15-foot Low Speed Wind Tunnel (LSWT) to establish a database for hot gas ingestion, one of the technologies critical to STOVL. This paper presents results showing the engine inlet distortions (both temperature and pressure) in a 9.2 percent scale Vectored Thrust STOVL model in ground effects. Results are shown for the forward nozzle splay angles of 0, -6, and 18 deg. The model support system had 4 deg of freedom, heated high pressure air for nozzle flow, and a suction system exhaust for inlet flow. The headwind (freestream) velocity was varied from 8 to 23 kn.

Johns, Albert L.↗

NASA NTRS1990

Three-dimensional shock-shock interactions on the scramjet inlet

The effects of shock impingement on the inlet of a scramjet engine are investigated numerically. The impinging shock is caused by the vehicle forebody. The interaction of this forebody shock with the inlet leading edge shock results in a very complex fully three-dimensional flowfield containing local regions of high pressure and intense heating. In the present investigation, this complex flowfield is calculated by solving the thin-layer Navier-Stokes equations using a finite-volume flux splitting technique due to van Leer. For zero or small sweep angles a Type IV interaction occurs while for moderate sweep of about 25 deg, a Type V interaction occurs. Both Type IV and Type V interactions are investigated.

Singh, D. J.↗

NASA NTRS1990

A flux-split solution procedure for unsteady inlet flows

The unstart and restart of an axisymmetric inlet is investigated using a flux-split procedure applied to the Euler and Reduced Navier Stokes (RNS) equations. A time consistent direct sparse matrix solver is applied to compute the transient flow field both internal and external to the inlet. Time varying oblique and normal shocks are captured. The code is quite general and is applicable for subsonic, transonic and supersonic free streams. The current analysis is concerned with supersonic flight conditions.

Pordal, H. S.↗

NASA NTRS1990

Comparison of 3-D viscous flow computations of Mach 5 inlet with experimental data

The accuracy of the time-marching three-dimensional Navier-Stokes code PARC3D (Cooper et al., 1987) is evaluated using experimental data on a Mach 5 mixed-compression inlet configuration, obtained in the 10 x 10-ft working section of the supersonic wind tunnel at NASA Lewis. The basic principles of the PARC3D computations (in which the inlet and tunnel geometry and the bleed conditions are fully modeled) and the experimental setup are described, and the results are presented in extensive graphs and characterized in detail. PARC3D is shown to give reasonably accurate predictions of the bleed-zone elimination of low-energy vortical flow due to shock/boundary-layer interaction; various possible improvements are briefly considered.

Reddy, D. R.↗