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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 1,351 records · Page 75
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Spontaneous ignition in afterburner segment tests at an inlet temperature of 1240 K and a pressure of 1 atmosphere with ASTM jet-A fuel

A brief testing program was undertaken to determine if spontaneous ignition and stable combustion could be obtained in a jet engine afterburning operating with an inlet temperature of 1240 K and a pressure of 1 atmosphere with ASTM Jet-A fuel. Spontaneous ignition with 100-percent combustion efficiency and stable burning was obtained using water-cooled fuel spraybars as flameholders.

Schultz, D. F.↗
NASA NTRS1974

Summary of recent investigations of inlet flow distortion effect on engine stability

A review is presented of recent experimental results, analytical procedures and test techniques employed to evaluate the effects of inlet flow distortion on the stability characteristics of representative afterburning turbofan and turbojet compression systems. Circumferential distortions of pressure and temperature, separately and in combination are considered. Resulting engine sensitivity measurements are compared with predictions based on simplified parallel compressor models and with several distortion descriptor parameters.

Graber, E. J., Jr.↗
NASA NTRS1974

Numerical solution of the time-dependent compressible Navier-Stokes equations in inlet regions

The results of a study to determine the effects of compressibility on the viscous flow through channels that have straight, parallel walls are presented. Two channel configurations are considered, the flow between two semi-infinite flat plates with uniform flow prescribed at the inlet plane and a cascade of semi-infinite flat plates with uniform flow introduced upstream. The flow field is modeled by using the time dependent, compressible Navier-Stokes equations. Time dependent solutions are obtained by using an explicit finite difference technique which advances the pressure on near field subsonic boundaries such that accurate steady state solutions are obtained. Steady state results at Reynolds number 20 and 150 are presented for Mach numbers between 0.09 and 0.36 and compared with the incompressible solutions of previous studies.

Olson, L. E.↗
NASA NTRS1974

Sound radiation from a high speed axial flow fan due to the inlet turbulence quadrupole interaction

A formula is obtained for the total acoustic power spectra radiated out the front of the fan as a function of frequency. The formula involves the design parameters of the fan as well as the statistical properties of the incident turbulence. Numerical results are calculated for values of the parameters in the range of interest for quiet fans tested at the Lewis Research Center. As in the dipole analysis, when the turbulence correlation lengths become equal to the interblade spacing, the predicted spectra exhibit peaks around the blade passing frequency and its harmonics. There has recently been considerable conjecture about whether the stretching of turbulent eddies as they enter a stationary fan could result in the inlet turbulence being the dominant source of pure tones from nontranslating fans. The results of the current analysis show that, unless the turbulent eddies become quite elongated, this noise source contributes predominantly to the broadband spectrum.

Goldstein, M. E.↗
NASA NTRS1974

Investigation of the stall-induced shock wave (hammershock) at the inlet to the engine

The peak static pressures measured at the inlet to the engine during stall are presented for a turbojet and two turbofan engines. It is shown for one turbofan and the turbojet that the static pressure ratio across the hammershock does not exceed significantly the normal shock pressure ratio necessary to stop the flow. The second turbofan engine did not follow this rule. Possible reasons for the departure are discussed. For the two turbofan engines the influence of the stall method on the hammershock intensity was investigated. Data related to the spatial distribution of pressure in the hammershock are also presented.

Kurkov, A. P.↗
NASA NTRS1974

Instantaneous distortion in a Mach 2.5, 40-percent-internal-contraction inlet and its effect on turbojet stall margin

An experimental wind tunnel investigation was conducted to determine the effects of time-variant distortions produced in a supersonic inlet on a J85-GE-13 turbojet engine. Results are presented principally in terms of instantaneous distortion amplitudes and total-pressure contours measured through compressor stall. They indicate that, although a time-averaged distortion may be far from a stall-inducing value, corresponding instantaneous distortion amplitudes can approach or exceed this value. Seven engine stall events were studied. In six of these events instantaneous distortion peaks of sufficient magnitude to cause stall were observed in the time period shortly before stall was detected.

Burstadt, P. L.↗
NASA NTRS1974

Effect of inlet ingestion of a wing tip vortex on turbojet stall margin

A two-dimensional inlet and a turbojet engine were investigated in a Mach 0.4 stream so as to ingest the tip vortex of a forward mounted wing. Results show that ingestion of a wing tip vortex by a turbojet engine can cause a large reduction in engine stall margin. The loss in stall compressor pressure ratio was primarily dependent on vortex location and rotational direction and not on total-pressure variations across the compressor face.

Mitchell, G. A.↗
NASA NTRS1973

The effect of inlet temperature and pressure distortion on turbojet performance

The effects on stability of steady-state, 180 degree extent circumferential distortions of inlet total temperature and pressure were experimentally determined for a turbojet engine. Results for both individual and combined temperature and pressure distortions are presented showing the losses incurred in stall pressure ratio and are compared with results predicted using a simplified parallel compressor model. The loss due to combined distortions was dependent upon the relative orientation between the low pressure and high temperature regions. Reasonable agreement was achieved between the predicted and observed loss in stall pressure ratio when based on a constant corrected speed relationship.

Braithwaite, W. M.↗
NASA NTRS1974

Summary of recent investigations of inlet flow distortion effects on engine stability

A review is presented of recent experimental results, analytical procedures and test techniques employed to evaluate the effects of inlet flow distortion on the stability characteristics of representative afterburning turbofan and turbojet compression systems. Circumferential distortions of pressure and temperature, separately and in combination, are considered. Resulting engine sensitivity measurements are compared with predictions based on simplified parallel compressor models and with several distortion descriptor parameters.

Graber, E. J., Jr.↗
NASA NTRS1974

Investigation of the stall-induced shock wave /hammershock/ at the inlet to the engine

The peak static pressures measured at the inlet to the engine during stall are presented for a turbojet and two turbofan engines. It is shown for one turbofan and the turbojet that the static pressure ratio across the hammershock does not exceed significantly the normal shock pressure ratio necessary to stop the flow. The second turbofan engine did not follow this rule. Possible reasons for the departure are discussed. For the two turbofan engines, the influence of the stall method on the hammershock intensity has been investigated. Data related to the spatial distribution of pressure in the hammershock are also presented.

Kurkov, A. P.↗
NASA NTRS1975

Experimental and analytical study of an inlet forebody for an airframe-integrated scramjet concept

Preliminary analytical and experimental inlet forebody investigations have been conducted at Mach numbers of 6.0 and 8.5. The forebody design concept consisted of a sharp-nosed right circular cone followed by elliptical cross sections. This concept resulted in swept isentropic compression which would allow swept cowl leading edges. Measurements were made to define the condition of the inviscid flow field developed by the forebody, including flow profiles in the vicinity of cowl leading-edge stations, and the three-dimensional boundary-layer effects. The investigation verified some of the expected differences between the predicted and the experimental results.

Andrews, E. H., Jr.↗
NASA NTRS1975

Evaluation of some control-volume techniques for analysis of shock-boundary layer interactions in supersonic inlets

Various control-volume models used to analyze the shock boundary layer bleed interaction were investigated. The bleed assumptions of the models and their influence on the analytical solutions are discussed. The results of the analysis using these models are compared with experimental boundary-layer data taken in a supersonic inlet. The experimental Mach number upstream of the interaction was 1.66, and the oblique-shock pressure ratio was 1.33. The boundary layer data included bleed flow rates up to approximately 0.6 of the upstream boundary layer mass flow rate. The first model assumed the bleed was removed from the control volume with a momentum that was characterized by a pressure intermediate between the upstream and downstream pressures. The second model assumed the control volume was bounded by a streamline dividing the bleed and residual flows and eliminated the need to specify the momentum of the bleed flow. Comparison of the results using the models showed that specifying the bleed pressure in one model was equivalent to specifying the pressure along the dividing streamline in the other.

Hingst, W. R.↗
NASA NTRS1975

Performance of vortex generators in a Mach 2.5 low-bleed full scale 45-percent-internal-contraction axisymmetric inlet

Steady-state and dynamic flow characteristics associated with two sets of vortex generators having different mixing criteria were determined. The inlet performance with and without these vortex generators is presented. The vortex generators were successful in eliminating separation, increasing area-weighted total pressure recovery, and decreasing distortion. Transmission times obtained from cross-correlations of the wall static pressures and the diffuser exit total pressure showed no effect of the upstream flow characteristics on the diffuser exit pressures when generators were used. Without generators, separation occurred and the upstream pressure characteristics had immediate effects on the diffuser exit pressure characteristics.

Neumann, H. E.↗
NASA NTRS1975

Two-stage fan. 3: Data and performance with rotor tip casing treatment, uniform and distorted inlet flows

A two stage fan with a 1st-stage rotor design tip speed of 1450 ft/sec, a design pressure ratio of 2.8, and corrected flow of 184.2 lbm/sec was tested with axial skewed slots in the casings over the tips of both rotors. The variable stagger stators were set in the nominal positions. Casing treatment improved stall margin by nine percentage points at 70 percent speed but decreased stall margin, efficiency, and flow by small amounts at design speed. Treatment improved first stage performance at low speed only and decreased second stage performance at all operating conditions. Casing treatment did not affect the stall line with tip radially distorted flow but improved stall margin with circumferentially distorted flow. Casing treatment increased the attenuation for both types of inlet flow distortion.

Burger, G. D.↗
NASA NTRS1975

Transonic off-design drag and performance of three mixed-compression axisymmetric inlets

An experimental investigation was conducted to determine the off-design drag and pressure performance of three axisymmetric supersonic inlets in the transonic speed range. For typical engine airflows at Mach 0.8 the drag coefficient varied from 0.045 to 0.09; at Mach 1.2 the largest drag coefficient measured was 0.25. Below Mach 0.9 a lower drag resulted when all or at least part of the excess weight flow was spilled over the cowl rather than through the bypass doors; above Mach 1.1 the lowest drag was obtained by bypassing excess flow.

Woollett, R. R.↗
NASA NTRS1975

Evaluation by step response tests of prototype relief valves designed for YF-12 inlet stability bleed system

Two stability bleed system relief valves were tested in a special dynamic test facility. These poppet valves are prototypes for a stability bleed system designed for use in a YF-12 flight inlet. One valve is unshielded, while the other has a special shield to eliminate the flow effect pressures on the piston. The tests determined the size of a damping orifice to be used during wind tunnel tests of the bleed system and verified an analog simulation of the valves. The effects of initial pressure level, pressure step size, and spring rate were investigated.

Dustin, M. O.↗
NASA NTRS1975

Control of shock-wave boundary-layer interactions by bleed in supersonic mixed compression inlets

An experimental investigation has been conducted to determine the effect of bleed region geometry and bleed rate on shock wave-boundary layer interactions in an axisymmetric, mixed-compression inlet at a Mach number of 2.5. The full realizable reduction in transformed form factor is obtained by bleeding off about half the incident boundary layer mass flow. Bleeding upstream or downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise. Slanted holes are more effective than normal holes. Two different bleed hole sizes were tested without detectable difference in performance.

Fukuda, M. K.↗