Performance Of Supersonic Inlet Research Articles

Effects of Optimized Bleed System on Supersonic Inlet Performance and Buzz

This paper details the numerical investigations of bleed effects on supersonic inlet performance and buzz under various bleed conditions, including location, width, plenum static pressure, and inlet conditions (for example, inflow Mach number and throttling ratio). A bleed boundary condition model that reflects the effects of local flow expansion and porosity variation is proposed and validated. Flows around a supersonic inlet with bleeds are computed to show bleed effects on the inlet performance and buzz. By employing metamodeling in conjunction with a genetic algorithm, an optimal bleed condition maximizing the total pressure recovery is obtained, for each Mach number and throttling ratio, in the air chamber. A bleed condition that provides better performance for all Mach numbers and throttling ratios in the subcritical inlet operating range is then obtained by multipoint design optimization. Numerical results show that the optimal bleed condition, obtained by the multipoint design optimization, provides a much wider subcritical stable range for enhanced inlet operation by substantially increasing the total pressure recovery and suppressing inlet buzz.

Side Gust Effects on the Performance of a Supersonic Inlet with Bleed

A computational investigation on the effects of side gust on the performance of a supersonic inlet with a bleed system at Mach 1.8 is reported. This is an extension to a recent study (Halwas, H. K., and Aggarwal, S. K., “Effect of Side Gust on the Performance of External Compression Supersonic Inlet,” Journal of Aircraft, Vol. 56, No. 2, March 2019, pp. 569–582) that analyzed a similar inlet without bleed. The supersonic section is designed using the Taylor–Maccoll method, and the rest of the inlet is designed using a methodology from the literature. The bleed is positioned at the throat inner wall, and its shape is a combination of a ram scoop and a flush slot. ICEM CFD software is used to generate a three-dimensional (3-D) structured mesh. ANSYS Fluent 18.2 code is used to perform 3-D simulations. Grid sensitivity tests are performed using local grid refinements and by varying the total grid points from 3 to 10 million. The effects of side gust on the shock system and flow are analyzed by examining the Mach number and pressure distributions. The inlet performance is characterized in terms of the total pressure loss, flow distortion, and mass flow ratio parameters for a gust speed of and angles of 0, 30, and 60 deg. The results indicate that the inlet performance is significantly affected by the side gust, although the bleed system mostly removes the low-momentum flow.

Energy Addition–Based Virtual Cowl for Performance Enhancement of Scramjet Intake

AbstractEnergy addition has been considered a means to improve the performance of supersonic inlets. The energy addition to supersonic inlets can be considered as a virtual cowl, which can be imple...

Starting characteristics and phenomenon of a supersonic wind tunnel coupled with inlet model

Supersonic wind tunnel is regarded the most efficient experimental facility for studying the performance of supersonic inlets. In view of this background, the starting problem of a wind tunnel coupled with the inlet model is a critical experimental issue. However, inlet experiments cannot be performed when wind tunnels do not start. To satisfy this requirement, this study adopts numerical simulations to examine the starting characteristics of a wind tunnel coupled with the inlet model to serve as a basis for observing the flow at the starting procedure. Results of steady simulation show that the wind tunnel coupled with the medium-sized inlet model is still at risk in the starting procedure, although the inlet size of the model is considerably smaller than the physical limitation. Furthermore, imposing high total pressure on the wind tunnel inlet can make the wind tunnel restart. However, results also reveal that gradually increasing total pressure does not necessarily restart the wind tunnel, as wind tunnels can start only when total pressure amplitude is large enough. To improve the understanding of the phenomena, unsteady numerical simulations with different total pressure levels are performed on the wind tunnel inlet. Numerical results demonstrate the importance of generating a moving shock wave with adequate intensity that can overcome the structure of the separation zones near wind tunnel walls after sweeping the inlet model. Under this influence, the separation zones become relatively smaller, thereby weakening the intensity of the deduced shock wave. Several changes to improve the startup of the wind tunnel are also observed, such as the attachment of the detached shock wave mechanism situated in front of the inlet model and the disappearance of the shock wave/boundary-layer interactions near the lower wall. These changes can reduce total pressure loss in the wind tunnel and prompt the wind tunnel to restart. In conclusion, a sudden increase in total pressure can easily start the wind tunnel. The valve of the wind tunnel should therefore be opened rapidly regardless whether the wind tunnel is in the starting or non-starting procedure.

다공도 및 팽창파의 영향을 고려한 BLEED 경계조건 수치 모델링의 정확도 향상 연구

The present paper deals with accuracy improvement of a bleed boundary condition model used to improve the performance of supersonic inlets. In order to accurately predict the amount of bleed mass flow rates, this study performs a scaling of sonic flow coefficient data for 90-degree bleed holes in consideration of Prandtl-Meyer expansion theory. Furthermore, it is assumed that porosity varies with stream-wise location of the porous bleed plate to accurately predict downstream boundary layer profiles. The bleed boundary condition model is demonstrated through Computational Fluid Dynamics(CFD) simulations of bleed flows on a flat plate with/without an oblique shock. As a result, the bleed model shows the improved accuracy of bleed mass rates and downstream boundary layer profiles.

Vortex Generators for a Normal Shock/Boundary Layer Interaction with a Downstream Diffuser

T HE interaction of a shock wavewith a turbulent boundary layer constitutes a fundamental problem of high-speed fluid mechanics. A detailed survey of past work on high-speed interactions has been carried out by Settles and Dolling [1] and Smits and Dussauge [2]. The shock interaction problem is particularly germane to the design of supersonic inlets. In such supersonic inlets, deceleration of the flow is achieved through a succession of oblique shock waves followed by a terminal normal shock. Boundary layers form on the inlet surfaces and interact with the shock system, giving rise to various shock/boundary-layer interactions (SBLIs). Each interaction of oblique/normal shock waves with the boundary layer causes stagnation pressure losses and downstream spatial distortions seen by the engine. An inlet must be carefully designed to minimize these losses and distortions during the compression process since they affect overall propulsion performance. In mixed-compression inlets, shock-induced separation can lead to engine unstart, which requires that the entire propulsion system undergo a restart sequence during flight. In external compression inlets, specifically axisymmetric configurations, a thick hubside boundary layer increases blockage and can decrease compressor performance. Thus, successfully controlling SBLIs has the potential to significantly improve supersonic inlet performance. As will be discussed in the following, various techniques of flow control for SBLIs have been proposed. However, it is often difficult to interpret the results because the flowfield may be too specific to an individual inlet configuration or too basic such that a relationship to inlet performance is not clear. To address this issue, a newwind-tunnel flowfield has been proposed [3] that captures much of the key shock boundary-layer interaction physics of supersonic external compression inlets. Thisflowfieldwill be used to study the novel flow control methods introduced herein. The conventionalflow control technique for SBLI conditions in an engine inlet employs a bleed of the boundary layer [4,5]. This bleed Presented as Paper 2010-4464 at the 40th AIAA Fluid Dynamics Conference and Exhibit, Chicago, IL, 28 June–1 July 2010; received 31 January 2011; revision received 26 July 2011; accepted for publication 11 August 2011. Copyright©2011 by theAmerican Institute ofAeronautics and Astronautics, Inc. All rights reserved. Copies of this paper may be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 0748-4658/12 and $10.00 in correspondence with the CCC. Ph.D. Candidate, Aerospace Engineering, 104 S. Wright Street. Member AIAA. Professor of Aerodynamics, Department of Engineering, Trumpington Street. Associate Fellow AIAA. Professor, Mechanical and Aerospace Engineering, 122 Engineer’s Way. Associate Fellow AIAA. JOURNAL OF PROPULSION AND POWER Vol. 28, No. 1, January–February 2012

Numerical study on performance of supersonic inlets with various three-dimensional bumps

Numerical investigations were performed with a supersonic inlet system installed with a three-dimensional bump which was substituted for a diverter or conventional ramp-type compression systems at Mach 2. The modified inlets were designed to have two oblique shocks and a terminal normal shock followed by a subsonic diffuser, with a circular cross-section throughout. A numerical analysis was conducted to understand the three-dimensional flow field including shock/boundary layer interactions that occur around a three-dimensional bump and to evaluate the performance of the supersonic inlets. The current numerical simulations showed a bump-type inlet based on a conventional ramp-type inlet can provide an improvement in total pressure recovery downstream of the shock/boundary-layer interaction over a ramp-type inlet.

Survey of validation data base for shockwave boundary-layer interactions in supersonic inlets

The performance of supersonic inlets is strongly affected by the boundary-layer development over its internal surfaces. Boundary-layer bleed is used to suppress separation and to provide the desired inlet performance. The gain in pressure recovery and stability is accompanied, however, with a loss in mass flow and an increase in drag that must be minimized by optimizing the amount of bleed and bleed configuration. The purpose of this work is to review and assess the data base for shock boundary-layer interaction that is pertinent to the flow prediction in supersonic inlets. The first part of the review concerns mixed compression supersonic inlets and their bleed system performance at design and off-design conditions. Based on the assessment of this data, specific areas related to shock wave/boundary-layer interaction bleed, for flow-separation control, are identified, the last part of the review addresses this phenomena in various twoand three-dimensional flow configurations. The effect of bleed in the interaction zone is especially emphasized.

Effects of Airframe-Inlet Integration on Half -Axisymmetric and Two-Dimensional Supersonic Inlet Performance

Diagnostic data from large-scale, airframe-inlet model wind-tunnel tests have been analyzed to facilitate understanding of both axisymmetric and two-dimensional inlet performance in different installation configurations of highly maneuverable supersonic aircraft. Distortion pattern shifts have been used to identify duct swirl, apparently due to duct offset and loft. Examinations of total pressure surveys through the ducts of various inlet installations have been used to show the influence of flow separation at the inlet aperture and in the throat on performance degradation in various inlet configurations. Benefits and limitations of inlet shielding in supersonic maneuvering flight have been explored, showing performance advantages of half-axisymmetric inlets over two-dimensional inlets in the wing-shielded flowfields and more general performance advantages associated with fuselage shielding. Data comparisons have incorporated total pressure recovery, duct turbulence, and compressor face steady-state and dynamic flow distortion index showing, in particular, the importance of dynamic flow distortion measurement and sophisticated, high-speed analysis capability.

Possibilities for Improved Supersonic Inlet Performance

The results from detailed large-scale inlet tests were examined to assess the penalties, in terms of vehicle cruise range, for boundary-layer bleed, less than ideal engine-face pressure recovery, inlet weight, and external cowl drag. The assessment suggested specific improvements in the design of the inlet system that may increase the range of a typical supersonic transport approximately 6.9%. While no single design improvement can account for a large increase in range, careful attention to each design detail can yield a substantial total improvement. Because of noise considerations, future engines for advanced supersonic transports may be approximately 50% larger than used for the present study, making the effects of improved inlet performance even more important.