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Axisymmetric external compression double cone inlet - mach 3 to 0.8 - pressure recovery, mass flow ratio, cowl pressure and additive drag
The inlet was investigated with both a two-oblique-shock and an isentropic compression ramp. Provisions for a variety of bypass systems were incorporated and none decreased pressure recovery by more than about 1 percent. By utilizing throat boundary-layer control, critical pressure recoveries of 64 and 70 percent were obtained with the two-shock and isentropic ramps, respectively. Diffuser-exit air distortion was about 5 percent at critical operation Theoretical cowl pressure drag was equal to about 10 percent of the net thrust of an assumed engine.
Wind tunnel test results for a large scale inlet model designed for Mach 3.5 are presented and compared with analytical predictions. The inlet is an axisymmetric mixed-compression type with a lip diameter of 49.723 cm. The inlet design was developed using analytical procedures. Data are shown for freestream Mach numbers from 0.6 to 3.5. The test results indicate that boundary layer bleed requirements can be accurately predicted. Good agreement was obtained with analytical predictions of the flowfield structure and boundary layer development in the supersonic diffuser yielding high performance at the design Mach number. The highest engine face total pressure recovery at Mach 3.5 was 85.8%; this was obtained at 0.05 Mach tolerance with only 2.8% total pressure distortion and 13.4% bleed. In the started Mach number range from 1.6 to 3.5, the total pressure recovery in the throat, downstream of the terminal normal shock, ranged between 91% and 95%. Total pressure losses in the subsonic diffuser varied from 3% to 13%. The higher losses occuring between Mach 2.5 and 3.2 were believed to be caused by the rapid rate of increase in the area of the diffuser just downstream of the throat, possibly coupled with inadequate centerbody throat bleed. In the unstarted mode at transonic speeds, the maximum inlet flow was over 99% of the theoretical maximum capture mass-flow.
A finite difference computer program for turbulent compressible flow was used to establish the performance of several diffuser shapes for experimental testing. The diffusers were designed to have a linear change in Mach number, a linear change in pressure, or a curvature fitted by a quadratic equation. Testing was performed with M = 0.1 to 0.9 with and without boundary layer bleed. Above M = 0.6, data were obtained with a normal shock upstream of the diffuser entrance. Peak static pressure recovery occurred with a diffuser inlet M0.75. The quadratic diffuser yielded the highest total pressure recovery.
The results indicated increases in auxiliary-inlet pressure recovery with increases in scoop height relative to the boundary-layer thickness. The pressure recovery increased at about the same rate as theoretically predicted for an inlet in a boundary layer having a one-seventh power profile, but was only about 0.68 to 0.75 of the theoretically obtainable values. Under some operating conditions, flow from the primary jet was exhausted through the auxiliary inlet. This phenomenon could be predicted from the ejector pumping characteristics.
An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.This CD-ROM supplement contains inlet data including: Boundary layer data, Duct static pressure data, performance-AIP (fan face) data, Photos, Tunnel wall P-PTO data and definitions.
A finite difference computer program for turbulent compressible flow was used to establish the performance of several diffuser shapes for experimental testing. The diffusers were designed to have a linear change in Mach number, a linear change in pressure or a curvature fitted by a quadratic equation. Testing was performed with M = 0.1 to 0.9 with and without boundary layer bleed. Above M = 0.6, data were obtained with a normal shock upstream of the diffuser entrance. Peak static pressure recovery occurred with a diffuser inlet M-0.75. The quadratic diffuser yielded the highest total pressure recovery.
An experimental investigation was undertaken to measure the effect of various configurations of low-profile vortex generator arrays on the flow in a diffusing S-duct. Three parameters that characterize the vortex generator array were systematically varied to determine their effect: (1) the vortex generator height, (2) the streamwise location of the vortex generator array, and (3) the vortex generator spacing. Each array of vortex generators tested improved total pressure recovery. The configuration employing the largest vortex generators was the most effective in reducing total pressure distortion but did not produce the greatest total pressure recovery. No configuration of vortex generators completely eliminated the flow separation that naturally occurs in the S-duct; however the extent of the separated flow region was reduced.
An inlet instantaneous distortion study program sponsored by NASA was recently completed using an F-15 fighter aircraft. Peak distortion data from subscale inlet model wind tunnel tests are shown to be representative of full-scale flight test peak distortion. The effects on peak distortion are investigated for engine presence, Reynolds number, scale and frequency content. Data are presented which show that: (1) the effect of engine presence on total pressure recovery, peak distortion, and turbulence is small but favorable, (2) increasing the Reynolds number increases total pressure recovery, decreases peak distortion, and decreases turbulence, and (3) increasing the filter cutoff frequency increases the calculated values of both peak distortion and turbulence.
In order to reduce the cost and complexity associated with fuel injection and mixing experiments for high-speed flows, and to further enable optical access to the test section for nonintrusive diagnostics, the Enhanced Injection and Mixing Project (EIMP) utilizes an open flat plate configuration to characterize inert mixing properties of various fuel injectors for hypervelocity applications. The experiments also utilize reduced total temperature conditions to alleviate the need for hardware cooling. The use of "cold" flows and non-reacting mixtures for mixing experiments is not new, and has been extensively utilized as a screening technique for scramjet fuel injectors. The impact of reduced facility-air total temperature, and the use of inert fuel simulants, such as helium, on the mixing character of the flow has been assessed in previous numerical studies by the authors. Mixing performance was characterized for three different injectors: a strut, a ramp, and a flushwall. The present study focuses on the impact of using an open plate to approximate mixing in the duct. Toward this end, Reynolds-averaged simulations (RAS) were performed for the three fuel injectors in an open plate configuration and in a duct. The mixing parameters of interest, such as mixing efficiency and total pressure recovery, are then computed and compared for the two configurations. In addition to mixing efficiency and total pressure recovery, the combustion efficiency and thrust potential are also computed for the reacting simulations.
A design methodology based on streamline-tracing is discussed for the design of external-compression, supersonic inlets for flight below Mach 2.0. The methodology establishes a supersonic compression surface and capture cross-section by tracing streamlines through an axisymmetric Busemann flowfield. The compression system of shock and Mach waves is altered through modifications to the leading edge and shoulder of the compression surface. An external terminal shock is established to create subsonic flow which is diffused in the subsonic diffuser. The design methodology was implemented into the SUPIN inlet design tool. SUPIN uses specified design factors to design the inlets and computes the inlet performance, which includes the flow rates, total pressure recovery, and wave drag. A design study was conducted using SUPIN and the Wind-US computational fluid dynamics code to design and analyze the properties of two streamline-traced, external-compression (STEX) supersonic inlets for Mach 1.6 freestream conditions. The STEX inlets were compared to axisymmetric pitot, two-dimensional, and axisymmetric spike inlets. The STEX inlets had slightly lower total pressure recovery and higher levels of total pressure distortion than the axisymmetric spike inlet. The cowl wave drag coefficients of the STEX inlets were 20% of those for the axisymmetric spike inlet. The STEX inlets had external sound pressures that were 37% of those of the axisymmetric spike inlet, which may result in lower adverse sonic boom characteristics. The flexibility of the shape of the capture cross-section may result in benefits for the integration of STEX inlets with aircraft.
A mixed compression axisymmetric inlet model with a capture diameter of 50 cm was tested at Mach numbers ranging from 0.8 to 2.65 at 0 deg angle of attack and a constant total pressure of approximately 1 atm. Analytical methods accounting for the effects of both viscous and inviscid flows and incorporating empirical bleed discharge coefficients were used in the procedure for designing the inlet contours and the bleed system. Experimental results are compared with analytic predictions and are also compared with results from earlier tests of an inlet with the same internal contours but with a bleed system developed by cut and try methods in the wind tunnel. With the bleed configuration predicted by the design procedure, maximum total pressure recovery at the engine face at the design Mach number of 2.65 was 93 percent, with a total pressure distortion less than 10 percent. Corresponding bleed mass flow was approximately 7.5 percent, which was about 1.3 percent less than predicted. At lower supersonic Mach numbers, pressure recovery and bleed were generally lower and distortion generally higher.
An experimental investigation of a flush-mounted, S-duct inlet with large amounts of boundary layer ingestion has been conducted at Reynolds numbers up to full scale. The study was conducted in the NASA Langley Research Center 0.3-Meter Transonic Cryogenic Tunnel. In addition, a supplemental computational study on one of the inlet configurations was conducted using the Navier-Stokes flow solver, OVERFLOW. Tests were conducted at Mach numbers from 0.25 to 0.83, Reynolds numbers (based on aerodynamic interface plane diameter) from 5.1 million to 13.9 million (full-scale value), and inlet mass-flow ratios from 0.29 to 1.22, depending on Mach number. Results of the study indicated that increasing Mach number, increasing boundary layer thickness (relative to inlet height) or ingesting a boundary layer with a distorted profile decreased inlet performance. At Mach numbers above 0.4, increasing inlet airflow increased inlet pressure recovery but also increased distortion. Finally, inlet distortion was found to be relatively insensitive to Reynolds number, but pressure recovery increased slightly with increasing Reynolds number.
A matrix of simulations of hypersonic flow over blunt entry vehicles with steady and pulsing retropropulsion jets is presented. Retropropulsion in the supersonic domain is primarily designed to reduce vehicle velocity directly with thrust. Retropropulsion in the hypersonic domain may enable significant pressure recovery through unsteady, oblique shocks while providing a buffer of reactant gases with relatively low total temperature. Improved pressure recovery, a function of Mach number squared and oblique shock angle, could potentially serve to increase aerodynamic drag in this domain. Pulsing jets are studied to include an additional degree of freedom to search for resonances in an already unsteady flow domain with an objective to maximize the time-averaged drag coefficient. In this paradigm, small jets with minimal footprints of the nozzle exit on the vehicle forebody may be capable of delivering the requisite perturbations to the flow. Simulations are executed assuming inviscid, symmetric flow of a perfect gas to enable a rapid assessment of the parameter space (nozzle geometry, plenum conditions, jet pulse frequency). The pulsed-jet configuration produces moderately larger drag than the constant jet configuration but smaller drag than the jet-off case in this preliminary examination of a single design point. The fundamentals of a new algorithm for this challenging application with time dependent, interacting discontinuities using the feature detection capabilities of Walsh functions are introduced.
Propulsion integration for low-boom supersonic aircraft requires careful inlet selection, placement, and tailoring to achieve acceptable propulsive and aerodynamic performance, without compromising vehicle sonic boom loudness levels. In this investigation, an inward-turning streamline-traced and axisymmetric spike inlet are designed and independently installed on a conceptual low-boom supersonic demonstrator aircraft. The airframe was pre-shaped to achieve a target ground under-track loudness of 76.4 PLdB at cruise using an adjoint-based design optimization process. Aircraft and inlet performance characteristics were obtained by solution of the steady-state Reynolds-averaged Navier-Stokes equations. Isolated cruise inlet performance including total pressure recovery and distortion were computed and compared against installed inlet performance metrics. Evaluation of vehicle near-field pressure signatures, along with under- and off-track propagated loudness levels is also reported. Results indicate the integrated axisymmetric spike design offers higher inlet pressure recovery, lower fan distortion, and reduced sonic boom. The vehicle with streamline-traced inlet exhibits lower external wave drag, which translates to a higher lift-to-drag ratio and increased range capability.
Unsteady shock-induced separation of the ramp boundary layer was reduced and stabilized more effectively by external perforations than by external or internal slots. At Mach 2.0 peak total-pressure recovery was increased from 0.802 to 0.89 and stable mass-flow range was increased 185 percent over that for the solid ramp. Peak pressure recovery occurred just before instability. The 7 and one-third-diameter duct ahead of the engine reduced large total-pressure distortions but was not as successful for small distortions as obtained with throat bleed. By removing boundary-layer air the bypass nearly recovered the total-pressure loss due to the long duct.
An idealized axisymmetric, all-internal compression inlet was designed for a Mach number of 3.75. The objective of the design was to obtain a steady, one-dimensional transonic flow and a high over-all total pressure recovery. Boundary-layer removal was employed in the vicinity of the inflection point of the supersonic contour. Static and total-pressure fluctuations were measured in the transonic flow region. A total pressure recovery of about 90 percent was obtained with a boundary-layer-removal mass flow rate of 15 percent of the inlet mass flow rate. The accompanying root-mean-square total pressure fluctuation in the throat region was only 1 percent of the free stream total pressure. The test Mach number was 3.80 and the Reynolds number based on inlet diameter was 2.63 x 10(exp 6).
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.