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At least 73 records · Page 4
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NASA NTRS2003

Supersonic Test of the 10-Inch Bifurcated Two-Stage Supersonic Inlet (BTSSI)

Experimental results from testing of a novel supersonic inlet model in NASA Glenn Research Center's 10- by 10-foot supersonic wind tunnel are presented. The patented inlet concept, called Two-Stage Supersonic Inlet (TSSI), incorporates a large cavity, or throat slot, in the supersonic diffuser intended to enhance the stability of the normal shock. The present embodiment of the concept is a bifurcated twin-duct) design. During the course of testing an unusual 'semi-started' mode of operation was encountered. The inlet was able to spill up to 30 percent of the captured airstream without fully expelling the normal shock. In this mode, the total pressure recovery dropped approximately 6 percent without increasing steady-state distortion. Dynamic instrumentation at the cowl lip station indicates the semi-start mode may be a series of unstart/restart cycles with frequency ranging from 0.2 to 20 Hz. Engine face total pressure measurements indicate a modest impact due to this event. However, since the current test article does not have a representative subsonic diffuser (and is in fact separated), it is unclear how this mode of operation would effect an engine. Further investigation of this phenomenon is required before it is fully understood. Prior testing of the TSSI concept allowed extension of fully started inlet operation to regions of significantly reduced supply flow without reducing recovery. The test article was a smaller scale than the present test and was a single duct design. In the present test, the expanded range of stable operation with high recovery was not realized.

Carlin, C. M.↗
NASA NTRS1957

Interference Effects of Fuselage-Stored Missiles on Inlet Duct Model of an Interceptor-Type Aircraft at Mach Numbers 1.5 to 1.9

The effect of missile armament on the performance of an interceptor-type aircraft model has been determined at Mach numbers 1.5, 1.7, and 1.9 and at angles of attack to 19 deg. With this configuration missiles were carried in a bay located on the bottom of the aircraft fuselage and mounted to a rotatable missile door. Rotation of the door then brought the missiles into the external or firing position. The aircraft model was characterized by triangular-shaped normal-shock inlets located at the wing roots. Relatively short and curved subsonic diffusers fed simulated twin side-by-side turbojet engines. Inasmuch as the missile bay extended considerably ahead of the inlet station, rotation of the missile door created considerable disturbance of the flow entering the inlets. In comparison with the internal missile arrangement, the external missile configurations increased the model lift, drag, and pitching moment. While the diffuser-exit flow distortion and static-pressure fluctuations were not greatly affected, diffuser total-pressure recovery was reduced as much as 0.058 at Mach number 1.9 for one missile configuration. The most detrimental effect of missile-door rotation occurred at the transient door positions, or with the door halfwzy between the missiles-in and -out conditions. At this door position the flow into the inlets was highly asymmetrical. Although the performance of both left and right ducts was generally reduced, the inlet duct on the cavity side of the missile door was most severely penalized, becoming unstable recovery losses and increases in flow distortion. The installation of fuselage fences along the missile bay was only partially effective in reducing these losses.

Piercy, T. G.↗
NASA NTRS1973

Some design considerations for supersonic cruise mixed compression inlets

A mixed compression inlet designed for supersonic cruise has very demanding requirements for high total pressure recovery and low bleed and cowl drag. However, since the optimum inlet for supersonic cruise performance may have other undesirable characteristics, it is necessary to establish trade-offs between inlet performance and other inlet characteristics. Some of these trade-offs between the amount of internal compression, aerodynamic performance and angle-of-attack tolerance are reviewed. Techniques for analysis of boundary layer control and subsonic diffuser flow are discussed.

Bowditch, D. N.↗
NASA NTRS1973

Some design considerations for supersonic cruise mixed compression inlets

A mixed compression inlet designed for supersonic cruise has very demanding requirements for high total pressure recovery and low bleed and cowl drag. However, since the optimum inlet for supersonic cruise performance may have other undesirable characteristics, it is necessary to establish trade-offs between inlet performance and other inlet characteristics. The paper will review some of these trade-offs between the amount of internal compression, aerodynamic performance and angle-of-attack tolerance. Also some techniques in use at the Lewis Research Center for analysis of boundary layer control and subsonic diffuser flow will be discussed.

Bowditch, D. N.↗
NASA NTRS1974

Distortion in a full-scale bicone inlet with internal focused compression and 45 percent internal contraction

The distortion characteristics were investigated at the subsonic diffuser exit of a full-scale, Mach 2.5, axisymmetric, mixed compression inlet. Performance and steady-state distortion characteristics were obtained at zero and maximum angle of attack and during an inlet unstart-restart sequence. For the configuration with no cowl bleed, steady-state distortion P(max)P(min)P(bar) ranged from 0.10 for critical inlet operation at 0 deg angle-of-attack to 0.306 for supercritical inlet operation at 6.84 deg angle-of-attack. Vortex generators provided a 50 percent reduction in steady-state distortion for critical operation. Bleed has a smaller effect on steady-stated distortion.

Wasserbauer, J. F.↗
NASA NTRS1976

Study of two axisymmetric inlets designed for Mach 3.5

Results from wind-tunnel tests of two large-scale models of axisymmetric mixed-compression inlet systems designed for Mach number 3.5 are compared. One inlet incorporated a 'traveling'-bleed system in an effort to achieve maximum transonic engine airflow supply. The other inlet required only a 'fixed'-bleed system, but had 21 percent less transonic airflow supply. The inlet with fixed bleed appears more attractive, if auxiliary airflow systems are used to increase the transonic airflow supply, because it can be 45 percent shorter and would be considerably lighter than the traveling-bleed inlet. In addition, the fixed-bleed inlet offers more operating-control margin at supersonic Mach numbers when the inlet is started. Further, its off-design performance is higher because separation of the flow in the subsonic diffuser can be avoided - something that apparently cannot be done with a traveling-bleed inlet without reducing the transonic airflow supply. Finally, it appears that the management and efficiency of bleed airflow for the fixed-bleed inlet can be improved using analytical methods verified by the tests.

Sorensen, N. E.↗
NASA NTRS1980

Stability aspects of diverging subsonic flow

The paper describes the computation of two-dimensional, subsonic, diverging internal flows and how they differ from the corresponding converging flows. Such diverging or decelerating flows occur in such obvious places as subsonic diffusers and inlets; however, such flows also occur in supersonic nozzles in the presence of a normal shock. The flow instability and its relation to the numerical method used, boundary conditions, and viscous effects are assessed both analytically and numerically. The inviscid flow is shown to be physically unstable and a poor representation of the true viscous flow.

Cline, M. C.↗
NASA NTRS1979

Calculation of three-dimensional turbulent subsonic flows in transition ducts

A method for computing three-dimensional turbulent subsonic flow in curved ducts is being developed. A set of tube-like surface oriented coordinates is employed for a general class of geometries applicable to subsonic diffusers with offset bends. The geometric formulation is complex and no previous treatment of this class of viscous flow problems is known to the authors. The duct centerline is a space curve specified by piecewise polynomials. A Frenet frame is located on the centerline at each axial location. The cross sections are described by superellipses imbedded in the Frenet frame. Duct surfaces are also coordinate surfaces, which greatly simplifies the boundary conditions. The resulting coordinates are nonorthogonal. An approximate set of governing equations is employed for viscous flows having strong flow in a primary flow direction. The derivation is coordinate invariant and the resulting equations are expressed in tensor form. These equations are solved by an efficient alternating direction implicit (ADI) method. This numerical method is generally stable and permits solution in difficult geometries using the general tensor formulation.

Levy, R.↗
NASA NTRS1984

Computation of viscous flow in planar and axisymmetric ducts by an implicit marching procedure

A streamwise marching procedure, approximately 200 times faster than a full Navier-Stokes procedure with comparable accuracy, is presented for solving problems of compressible viscous subsonic flow. Results are presented and compared with experimental data for the cases of developing turbulent flow in a circular pipe; turbulent flow in a two-dimensional S-duct; and turbulent flow in a typical subsonic diffuser. Prior to each main marching step, a preliminary marching step is taken in which the integral continuity equation and an uncoupled form of the streamwise momentum equation are solved simultaneously to obtain the viscous pressure correction. During the main marching step the equations for continuity, streamwise momentum, cross-flow momentum, and energy are solved simultaneously as a coupled system using an implicit finite-difference method, with the viscous pressure correction treated as a source term. The analysis may be used for flows with both favorable and adverse pressure gradients and to predict the location of flow separation.

Towne, C. E.↗
NASA NTRS1984

Numerical simulation of flow through inlets/diffusers with terminal shocks

Two-dimensional Euler and Navier-Stokes solutions of the flow through three inlet/diffuser configurations with terminal shock systems are reported. Calculations without bleed indicate that the terminal shock location is very sensitive to the outflow back pressure. For cases where there are no available experimental results, it becomes difficult to estimate the back pressure that will result in a terminal shock. Estimates based on quasi-one-dimensional analysis are not found adequate for complex two-dimensional flows. It is found that since the flow downstream of the terminal shock is subsonic, and what happens at the outflow boundary affects the flow inside the inlet, enough of the subsonic diffuser must be modeled to accurately predict the terminal shock region. The diffuser portion should be fairly long with the outflow boundary occurring in a region of more or less uniform flow to be able to prescribe a uniform back pressure. The third configuration studied was investigated with and without incorporating bleed in the code. It is found that the use of bleed stabilizes the shock location and allows solutions which without bleed result in unstarting of the inlet. Comparisons are made with available experimental data.

Talcott, N. A., Jr.↗
NASA NTRS1985

Application of computational fluid dynamics to complex inlet ducts

A three-dimensional parabolic Navier-Stokes code, PEPSIG, was used to analyze the flow in the subsonic diffuser section of a typical modern inlet design. The effect of curvature of the diffuser centerline and transitioning cross sections was studied to determine the primary cause of flow distortion in the duct. Total pressure values at the engine compressor face are reported.

Towne, C. E.↗
NASA NTRS1985

Application of computational fluid dynamics to complex inlet ducts

A three-dimensional parabolic Navier-Stokes code, PEPSIG, was used to analyze the flow in the subsonic diffuser section of a typical modern inlet design. The effect of curvature of the diffuser centerline and transitioning cross sections was studied to determine the primary cause of flow distortion in the duct. Total pressure values at the engine compressor face are reported.

Towne, C. E.↗
NASA NTRS1986

Low-speed performance of an axisymmetric, mixed-compression, supersonic inlet with auxiliary inlets

A test program was conducted to determine the aerodynamic performance and acoustic characteristics associated with the low-speed operation of a supersonic, axisymmetric, mixed-compression inlet with auxiliary inlets. Blow-in-auxiliary doors were installed on the NASA Ames P inlet. One door per quadrant was located on the cowl in the subsonic diffuser selection of the inlet. Auxiliary inlets with areas of 20 and 40 percent of the inlet capture area were tested statically and at free-stream Mach numbers of 0.1 and 0.2. The effects of boundary layer bleed inflow were investigated. A JT8D fan simulator driven by compressed air was used to pump inlet flow and to provide a characteristic noise signature. Baseline data were obtained at static free-stream conditions with the sharp P-inlet cowl lip replaced by a blunt lip. Auxiliary inlets increased overall total pressure recovery of the order of 10 percent.

Trefny, C. J.↗
NASA NTRS1986

Multispecies CARS measurements in turbulent combustion

A coherent anti-Stokes Raman scattering (CARS) instrument was upgraded to more accurately measure the number density of oxygen molecules. This instrument is designed to simultaneously measure nitrogen number density, oxygen number density, and the ro-vibrational temperature of nitrogen. The CARS is a nonintrusive diagnostic technique which utilizes a pulsed Nd:YAG laser, broadband dye lasers, a complex optical system, and an intensified photodiode array detector to sample hostile combustion environments. Measurements were made in a flame produced by a coaxial subsonic diffusion hydrogen-air burner. These data were used as a data base for comparison with the results of two separate computational fluid dynamics calculations. Complications which had arisen in previous studies with regard to interpretation of the oxygen data were eliminated.

Antcliff, Richard R.↗
NASA NTRS1992

Three-dimensional compressible turbulent computations for a diffusing S-duct

The purpose of the present study was to evaluate the capability of the computational fluid dynamics computer program PARC3D to model flow in a typical diffusing subsonic S-duct, with strong secondary flows. This evaluation is needed to provide confidence in the analysis of aircraft inlets, which have similar geometries. The performance predictions include total pressure profiles, static pressures, velocity profiles, boundary layer data, and skin friction data. Flow in the S-duct is subsonic, and the boundary layers are assumed to be turbulent. The results for both H and O grid solutions, are compared with existing test data.

Smith, C. F.↗
NASA NTRS1950

Investigation of Operating Pressure Ratio of a Supersonic Wind Tunnel Utilizing Distributed Boundary-layer Suction in Test Section

Effect of distributed boundary-layer suction on operating pressure ratio of a supersonic wind tunnel was investigated. Investigation was made in 3.84- by 10-inch supersonic tunnel operating at Mach number 2.0 and suction was applied in neighborhood of the normal shock to two walls of a constant-area extension of test section. A reduction of 4 percent of operating pressure ratio was attributed to improved flow conditions at subsonic-diffuser inlet. The theoretical normal shock was, in practice, replaced by a multiple-branch shock configuration across which the flow parameters changed in approximate accordance with the Rankine-Hugoniot values.

Cohen, C B↗
NASA NTRS1955

Preliminary Investigation of Effect on Performance of Dividing Conical-spike Nose Inlets into Halves at Mach Numbers 1.5 to 2.0

Inserting a splitter plate in the subsonic diffuser caused a pressure-recovery loss of about 1 percent for an inlet with a long nearly constant-area throat section. The loss was due to the increased surface area. Another inlet, which had a comparatively rapid area increase immediately after the throat, experienced pressure-recovery losses of 5 and 6 percent at Mach numbers of 1.8 and 2.0, respectively, and about 1 percent at Mach 1.5.

Allen, John L↗