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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 397 records · Page 22
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Pressure activated stability-bypass-control valves to increase the stable airflow range of a Mach 2.5 inlet with 40 percent internal contraction

The throat of a Mach 2.5 inlet with a coldpipe termination was fitted with a stability-bypass system. The inlet stable airflow range provided by various stability-bypass entrance configurations in alternate combination with several stability-bypass exit controls was determined for both steady-state conditions and internal transient pulses. Transient results were also obtained for the inlet with a choke point at the diffuser exit. Instart angles of attack were determined for the various stability-bypass entrance configurations. The response of the inlet-coldpipe system to internal and external oscillating disturbances was determined. Poppet valves at the stability-bypass exit provided an inlet stable airflow range of 28 percent or greater at all static and transient conditions.

Mitchell, G. A.↗
NASA NTRS1973

Minimizing boundary layer bleed for a mixed compression inlet

An experimental investigation of a full scale mixed compression inlet sized for the TF30-P-3 turbofan engine was conducted at Mach 2.5 and 2.0 operating conditions. The two cone axisymmetric inlet had minimum internal contraction consistent with high total pressure recovery and low cowl drag. At Mach 2.5, inlet recovery exceeded 0.90 with only 0.02 centerbody bleed mass-flow ratio and zero cowl bleed. A centerbody bleed of approximately 0.05 gave a maximum inlet unstart angle-of-attack of 6.85 deg. Inlet performance and angle-of-attack tolerance is presented for operation at Mach 2.5 and 2.0.

Wasserbauer, J. F.↗
NASA NTRS1974

Wind tunnel and flight performance of the YF-12 inlet system

The steady-state internal performance from 1/3-scale and full-scale wind tunnel models of the YF-12 inlet system is compared with results from flight tests. All systems were thoroughly instrumented for static and total pressure measurements. Results obtained included the inlet mass flows, surface pressure distribution, boundary layer profiles, and detailed total pressure measurements at the engine face. Inlet performance is compared at Mach numbers 2.1 and 2.8 and at several Reynolds numbers over a wide range of inlet operating conditions. Analysis generally shows good agreement in the results between the three systems when inlet conditions are closely matched.

Smeltzer, D. B.↗
NASA NTRS1975

Experimental investigation of a Mach 6 fixed-geometry inlet featuring a swept external-internal compression flow field

An investigation of a fixed-geometry, swept external-internal compression inlet was conducted at a Mach number of 6.0 and a test-section Reynolds number of 1.55 x 10 to the 7th power per meter. The test conditions was constant for all runs with stagnation pressure and temperature at 20 atmospheres and 500 K, respectively. Tests were made at angles of attack of -5 deg, 0 deg, 3 deg, and 5 deg. Measurements consisted of pitot- and static-pressure surveys in inlet throat, wall static pressures, and surface temperatures. Boundary-layer bleed was provided on the centerbody and on the cowl internal surface. The inlet performance was consistently high over the range of the angle of attack tested, with an overall average total pressure recovery of 78 percent and corresponding adiabatic kinetic-energy efficiency of 99 percent. The inlet throat flow distribution was uniform and the Mach number and pressure level were of the correct magnitude for efficient combustor design. The utilization of a swept compression field to meet the starting requirements of a fixed-geometry inlet produced neither flow instability nor a tendency to unstart.

Torrence, M. G.↗
NASA NTRS1975

Aeroacoustic performance of scale model sonic inlets

Various means can be used to vary engine inlet throat areas in order to reduce noise generated at takeoff and approach. Low-speed wind-tunnel tests were performed to evaluate the aeroacoustic performance of two takeoff geometries (cylindrical centerbody and bulb-shaped centerbody) and four approach geometries (bulb-shaped centerbody, annular ring, radial vanes, and step diffuser). The effects of inlet internal lip shape and diffuser length on sonic inlet performance were also investigated. It was found that the single-passage geometries, in particular a cylindrical centerbody takeoff geometry and a bulb-shaped centerbody approach geometry, provide the highest level of aeroacoustic performance. Increasing the inlet lip contraction ratio extends the maximum incidence angle for attached lip flow, while increasing the inlet diffuser length increases total pressure recovery for a given amount of noise suppression.

Abbott, J. M.↗
NASA NTRS1975

Inlet performance of the integrated Langley scramjet module

The inlet concept for the Langley Scramjet Module has been developed and proven in Langley wind tunnels over a Mach number range from 2.3 to 6.0 (flight simulation of Mach 2.6 to 7.6). This modular engine concept is designed to integrate with the airframe, which results in precompression of the engine airflow by the vehicle bow shock and additional expansion of the nozzle exhaust gas by the afterbody of the vehicle. With these integration advantages, the inlet can be designed with modest contraction ratios and fixed geometry. Also, the module nozzle exit area can be equal to the capture area, which permits the cowl to be alined with the local flow producing minimum external drag. The inlet leading edges and planar compression surfaces are swept at 48 deg, which provides spillage at low Mach numbers for starting and which reduces the pressure gradient on the top surface to permit ingestion of the vehicle forebody boundary layer into the inlet without separating. Three fuel injection struts provide for the use of a short combustor having low internal cooling requirements. Schedules for mass capture ratio, contraction ratio, and total pressure recovery are well within the acceptable range for a good scramjet propulsion device. The fixed geometry, minimum external drag design has proven to be a practical, high-performance inlet concept.

Trexler, C. A.↗
NASA NTRS1975

Poppet valve control of throat stability bypass to increase stable airflow range of a Mach 2.5. inlet with 60 percent internal contraction

The throat of a Mach 2.5 inlet with a coldpipe termination was fitted with a stability-bypass system. System variations included several stability bypass entrance configurations. Poppet valves controlled the bypass airflow. The inlet stable airflow range achieved with each configuration was determined for both steady state conditions and internal pulse transients. Results are compared with those obtained without a stability bypass system. Transient results were also obtained for the inlet with a choke point at the diffuser exit and for the inlet with large and small stability bypass plenum volumes. Poppet valves at the stability bypass exit provided the inlet with a stable airflow range of 20 percent or greater at all static and transient conditions.

Mitchell, G. A.↗
NASA NTRS1975

Quiet Clean Short-Haul Experimental Engine (QCSEE) aerodynamic characteristics of 30.5 centimeter diameter inlets

A low speed test program was conducted in a 9- by 15-foot V/STOL wind tunnel to investigate internal performance characteristics and determine key design features required for an inlet to meet the demanding operational conditions of the QCSEE application. Four models each having a design average throat Mach number of 0.79 were tested over a range of incidence angle, throat Mach number, and freestream velocity. Principal design variable was internal lip diameter ratio. Stable, efficient inlet performance was found to be feasible at and beyond the 50 deg incidence angle required by the QCSEE application at its 41.2 m/sec (80 knot) nominal takeoff velocity, through suitably designed inlet lip and diffuser components. Forebody design was found to significantly impact flow stability via nose curvature. Measured inlet wall pressures were used to select a location for the inlet throat Mach number control's static pressure port that properly balanced the conflicting demands of relative insensitivity to flow incidence and sufficiently high response to changes in engine flow demand.

Paul, D. L.↗
NASA NTRS1975

Fan noise reduction at subsonic and supersonic tip speeds with high Mach number inlets

Results are presented for an experimental study of two fixed geometry inlets and one translating centerbody type inlet tested in an anechoic-chamber transonic-compressor facility. The test vehicle was a 12-inch tip diameter single-stage transonic compressor with 19 rotor blades and 30 stator blades located approximately two chord lengths downstream of the rotor. It is shown that the translating centerbody type of inlet is superior to the collapsing cowl type for noise reduction at high subsonic Mach numbers, both acoustically and aerodynamically. The feasibility of using a high inlet Mach number depends largely on the maximum area ratio requirement (normally at landing). The high inlet Mach numbers are more effective in reducing high-frequency noise than low frequencies and more effective in reducing the blade passage frequency than the overall noise level.

Lumsdaine, E.↗
NASA NTRS1976

Advanced inlet duct noise reduction concepts

A progress report is given on the implications of inlet noise reduction on aircraft direct operating costs (DOC). It considers treated inlet rings, various other inlet noise reduction concepts, and forward-speed effects. The report was limited to relatively well-established approaches to inlet noise reduction, such as acoustic liners and fixed-geometry/high-subsonic-speed inlets which are the focus of considerable current research activity. All of the concepts discussed are of a "passive" nature, i.e., no moving parts or electrical feedback systems.

Chestnutt, D.↗
NASA NTRS1977

Wind-tunnel tests on a 3-dimensional fixed-geometry scramjet inlet at M = 2.30 to 4.60

Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.

Mueller, J. N.↗
NASA NTRS1977

Analytical and experimental studies of acoustic performance of segmented liners in a compressor inlet

The performance of axially segmented (phased) acoustic treatment liners in the inlet of a compressor was investigated. Topics discussed include: (1) the validation of a theoretical procedure to predict propagation and suppression characteristics of duct liners; (2) the in-duct measurement of spinning modes; (3) investigation of phased treatment designs; (4) high Mach inlet acoustic tests; and (5) an experimental investigation of inlet turbulence. The analytical prediction for the multi-segmented treatment was found to provide the correct order of magnitude of suppression and was generally within 50% of that determined experimentally. Refinements required to improve the correlation are identified. Suppression due to high subsonic Mach number flow effects was found to become significant above an average throat Mach number of 0.65 to 0.7 and 20 PNdB was achieved with an average throat Mach number in the range of 0.80 to 0.85. For the measured turbulence in the inlet, including the axial and circumferential turbulence intensities and the axial integral length scale, data are presented with and without an inlet screen showing that the screen reduced the turbulence intensities and that the BPF noise was reduced as a consequence.

Motsinger, R. E.↗
NASA NTRS1977

Use of finite element techniques in the determination of the acoustic properties of turbofan inlets

This paper describes the application of the finite-element method in combination with Galerkin's method in the determination of the acoustic properties of turbofan inlets containing high-subsonic-Mach-number flows. An approximate solution to the steady inviscid flow field is obtained using an integral method for calculating the potential flow field in the inlet with a correction to account for compressibility effects. The accuracy of the finite-element technique in predicting the acoustic properties of annular ducts has been checked by comparison with available analytical solutions for the problems of planeand spinning-wave propagation through a hard-walled annular duct with a constant mean flow. Results are presented comparing low-frequency plane-wave propagation through a hard-walled turbofan inlet containing a one-dimensional flow with the same inlet containing a fully two-dimensional axisymmetric mean flow. It is shown that when one-dimensional mean flow is assumed to exist in the duct, the plane wave propagates with relatively little distortion. However, propagation of a plane wave through the fully two-dimensional flow field in the inlet produces severe distortion due to the excitation of higher-order modes.

Sigman, R. K.↗
NASA NTRS1977

Fan inlet for a V/STOL airplane

An inlet was designed for a tilt nacelle lift/cruise fan. A complete set of low speed data was obtained with a 38-cm diameter cold-flow model, and a partial set of data was obtained with a 140-cm-diameter inlet coupled to a high bypass turbofan. The results indicate that a fixed-geometry inlet is a viable concept for the proposed application. Analyses of the data demonstrate that at high angles of incidence and at low airflow rates inlet-flow separation is strongly sensitive to Reynolds number and, consequently, the large-scale inlet has a significantly greater range of separation-free operation.

Koncsek, J. L.↗
NASA NTRS1977

Atmospheric effects on inlets for supersonic cruise aircraft

Time response effects of atmospheric-type disturbances on inlets and some implications regarding inlet controls are considered. A study involving the simulation of a mixed-compression inlet is described. The study employs a linear dynamic analysis, modified for a significant nonlinearity. Transient disturbances in ambient temperature and pressure and wind gusts are considered independently at the inlet cruise Mach number of 2.5. The presented results show the maximum amplitude of a triangular-wave disturbance, that does not cause inlet unstart, as a function of disturbance pulse width.

Cole, G. L.↗
NASA NTRS1978

A combined potential and viscous flow solution for V/STOL engine inlets

A potential flow routine and a viscous boundary layer routine have been combined into a single routine for estimating the flow in and around subsonic inlets. In this combined routine, the viscous flow solution about the inlet body is obtained by adding the viscous displacement thickness to the inlet geometry. Combination of the two flow solutions has resulted in cost savings, both in preparation time and in computer time. This routine is a useful tool in optimizing lip shapes for V/STOL inlets. The method of combining the routine, comparison with NASA test data, and utilization of the routine for V/STOL inlet design are presented.

Ybarra, A. H.↗
NASA NTRS1979

Critical mass flux through short Borda type inlets of various cross sections

Mass flux measurements associated with chocked flows through four Borda type inlet geometries: circular, square, triangular and rectangular (two-dimensional) and two sharp edged geometries taken over a very wide range of inlet stagnation conditions. The measurements indicate that: (1) the mass flux is independent of the inlet cross-section geometry and (2) the mass flux is dependent only on the inlet stagnation conditions. Also by using choked flow results found in the literature, the reduced mass flux is independent of working fluid. Two implications are drawn which remain to be verified: (1) since seal leak rates are weakly dependent on geometry but pressure distribution is strongly dependent on geometry, seal design efforts should be directed more toward controlling the dynamics, and (2) high-L/D ducts of arbitrary cross section and Borda type inlets can possess free jets.

Hendricks, R. C.↗
NASA NTRS1979

Conceptual study of a turbojet/ramjet inlet

An inlet concept for separate turbojet and ramjet engines was defined and compared with an equivalent inlet for a wraparound turboramjet engine. The comparison was made for a typical high altitude hypersonic cruise vehicle where the turbojet inlet capture area was required to be half as large as the ramjet inlet capture area at cruise. The use of a shorter nacelle having substantially lower cooling requirements at cruise for the inlet concept for separate turbojet and ramjet engines is suggested. The separate engine concept better isolates the turbojet from the ramjet, requires no special close off mechanisms within the turbojet, and avoids the circumferential heat load imposed by a wraparound ramjet. A more variable geometry is required.

Weidner, J. P.↗