Boundary-layer Bleed System Research Articles

Numerical Investigation of Aerodynamic and Electromagnetic Performances for S-Duct Caret Intake with Boundary-Layer Bleed System

This study presents the numerical results for the aerodynamic and electromagnetic performances of an S-duct caret intake. Using the multilevel fast multipole method (MLFMM) to solve Maxwell equations, the current on the intake surface is calculated, and the radar cross-section (RCS) is analyzed. Moreover, the intake flow field is numerically investigated using the SST k–ω turbulence model to solve the Reynolds-averaged Navier–Stokes equations. Compared to a straight intake, for an S-duct caret intake, the average RCS is lower by 7.65 dB, and the maximum RCS difference value caused by the blade rotation is lower by 6.75 dB. However, the flow capacity deteriorates when the total pressure recovery coefficient decreases by 0.004. Based on the analysis of the aerodynamic and electromagnetic characteristics of different intakes, a double S-duct intake is designed. Compared to a traditional S-duct intake, for the novel intake after model parameter modification, the average RCS is lower by 0.05 dB, and the total pressure distortion (TPD) is lower by 0.18. The analysis of the effects of different boundary-layer bleed systems shows that the symmetrical layout adversely affects the aerodynamic and electromagnetic performances of the S-duct intake, but the unilateral partial layout is beneficial, whose TPD is lower by 0.04 and average RCS is higher by −2.17 dB compared to a straight intake.

Numerical investigation of impact of fixed-exit boundary-layer bleed on supersonic intake performance

Results of numerical investigation of flow and performance characteristics of a rectangular supersonic intake with boundary-layer bleed system are presented. The effects of several flush slot bleed geometric parameters including entrance and exit area, plenum size, rear edge vertical displacement and the position relative to the incident shock have been studied to clarify their effects on the compression performance. The viscous flowfield has been obtained by solving Favre averaged Navier-Stokes equations with shear stress transport (SST) k-ω turbulence model. Moreover, the numerical methodology deployed is validated with the experimental data from literature. The analysis has been carried out in the freestream Mach number range of 1.8-2.5 and at 00 angle-of-attack, corresponding to a unit Reynolds number of approximately 1.9−2.7×107 m−1. The results show that with the use of bleed, the exit flowfield uniformity as well as the critical and peak total pressure recovery (TPR) improves considerably at all operating conditions. Furthermore, the analysis indicates that the exit throat area of a slot bleed operating at a particular design condition can be increased up to a point such that it is just sufficient to eliminate the upstream boundary-layer at its peak operation. The variation in the bleed entrance area or its position relative to the throat entrance only alters the critical performance depending on the supercritical boundary-layer mass flow loss; however, the peak TPR or the sustainable backpressure remains the same at all upstream conditions, controlled by the fixed-exit throat area. Further, it is found that for the removal of comparable amount of boundary-layer mass flow, the effects of pure flush slot or of a configuration with slight vertical displacement at its rear edge, located at the same position are similar.

Study of the Possibilities of Gasdynamic Air Flow Control in the Spatial Air Intake of a Light Supersonic Business-Class Aircraft by the RANS/ILES Method

The effect of the boundary-layer bleed system on the flow and characteristics of the spatial air intake of a supersonic business-class aircraft integrated with an airframe simulator was studied by the combined high-resolution Reynolds Averaged Navier–Stokes Implicit Large Eddy Simulation method (RANS/ILES ). Two variants of geometry with and without the boundary-layer bleed system were considered. The effect of gasdynamic flow control with the use of synthetic jets was studied for both air intake variants. The effect produced by the boundary-layer bleed system and the regime parameters of synthetic jets on the flow in the air intake and its characteristics, the level of turbulent fluctuations, and the surge boundary was studied. The level of velocity and pressure fluctuations was established to be lower at the outlet of the air intake without the boundary-layer bleed system. The application of synthetic jets made it possible to reduce the total pressure losses and level of turbulent fluctuations in the channel of the air intake and at its outlet, thus increasing the range of its stable operation.

A study of unsteady-state operating conditions of a supersonic inlet by the RANS/ILES method

Using the hybrid RANS/ILES method, a flow was investigated under unsteady-state conditions in a rectangular mixed-compression supersonic inlet at Md = 2. The computations were made for Mach numbers of the incident flow M0 = 1.8, 2, and 3. The geometry of two variants was investigated, viz., with the boundary layer bleed system and without it. Calculations were performed on the meshs containing (1.69–1.78) × 106 cells. The flow rate through the supersonic inlet was varied within a wide range. Under most of the conditions investigated, the flow in the duct was unsteady as a consequence of separation of the boundary layer upon its interaction with the shock waves. For all geometric variants investigated and M0, the throttle characteristics were constructed by the averaged flow parameters as well as the dependences of the flow rate through the boundary layer bleeding system and the static pressure pulsation level on the air inlet throttle ratio. Comparison of the computed results with the experimental data showed a good agreement in terms of both the averaged flow parameters and the pulsation characteristics.

Experimental Analyses of Inlet Characteristics of an Airbreathing Missile with Boundary-Layer Bleed

The performance of different boundary-layer bleed systems of a two-dimensional ramjet inlet has been investigated experimentally. The main objective of the study has been to identify the correlation between the inlet’s performance with different boundary-layer bleed geometries and the inlet’s internal flow. Experiments have therefore been carried out with varied boundary-layer bleed geometries at an angle of attack of 0 deg and at Mach numbers ranging from 2.5 to 3.5 in the Trisonic Wind Tunnel Cologne. The results are a continuation of an earlier study and show that the inlet’s performance and the effectiveness of boundary-layer bleed system depend on shock/boundary-layer interaction, unsteadiness of the flow, and flow separation.

Experimental Study of Boundary-Layer Bleed Impact on Ramjet Inlet Performance

The performance of different boundary-layer bleed systems of a two-dimensional ramjet inlet has been investigated experimentally. The main objective of the study has been to identify the boundary-layer bleed geometry that increases the inlet’s performance and to stabilize the subcritical condition (i.e., later buzzing). Experiments have therefore been carried out with varied boundary-layer bleed geometries at an angle of attack of 0 deg and at Mach numbers ranging from 2.5 to 3.5. The results show that the effectiveness of a boundary-layer bleed system depends on shock, boundary-layer interaction, unsteadiness of the flow, and flow separation.

Inlet Drag and Stability Considerations for Mo = 2.00 Design

The results of an experimental study to evaluate the performance and low-flow stability limits of several isolated inlet models are presented. The lowest drag approach to provide large-flow regulation was a bypass device for handling the excess airflow. The use of compression surface modulation or a boundary-layer bleed system to dispose of the excess airflow incurred similar drag penalties which were significantly greater than those for the bypass configuration. A good correlation of low-flow stability range data with a shock-lip geometry parameter is demonstrated. A by-product of this work is the introduction of a drag parameter that allows a more complete approach to quantifying inlet drag characteristics.

Transonic Performance of an Auxiliary Airflow System for Axisymmetric Inlets

A large-scale model of an axisymmetric inlet with a centerbody auxiliary airflow system has been tested in the wind tunnel at transonic speeds. The auxiliary system allows additional airflow (other than in the main duct formed by the cowl and translating centerbody) to pass through the centerbody of the inlet and combine with the main duct airflow on its way to the engine face. The results of the tests are presented, and the inlet performance is compared to a closely related alternative inlet with a ''traveling boundary-layer bleed system that precludes the use of a centerbody auxiliary airflow system. The comparison shows that the auxiliary airflow inlet can supply 7.7% more engine face airflow at Mach number 1.0 and is 26% shorter than the traveling bleed inlet. Even though maximum transonic airflow was not achieved at a comparable engine face mass-flow ratio of 0.580, a total-pressure distortion of 0.10, and a total-pressure recovery of 0.985 were achieved for the auxiliary airflow inlet, whereas a recovery of only 0.965 was achieved for the traveling bleed inlet. Nomenclature A aux = area of auxiliary airflow entrance annulus Ac = cowl-lip area (capture area) A 0 = freestream tube area entering the inlet D = cowl lip diameter Mdes = design Mach number A/th = throat Mach number

Development Program for the F-15 Inlet

This paper describes the F-15 variable capture area inlet and the associated control system. Factors that influenced integration of the inlet/inlet control system with the total airframe are discussed. The methodology of using instantaneous total pressure patterns at the engine face to evaluate the effect of inlet distortion on engine operation is briefly described. Several inlet design variables that were tested are discussed. In addition to its effects on inlet behavior, the variable capture area reduces cruise drag, favorably influences aircraft longitudinal stability at low speeds, and reduces loads on the inlet backup structure at certain supersonic conditions. The boundary-layer bleed system optimization demonstrated that a slot in the throat was the single most effective location for bleed removal. Comparative results are presented for the final configuration between model and full-scale wind-tunnel testing, and subsequent flight testing. Results show that full-scale instantaneous distortion typically is equal to or less than that measured on a l/6th scale model.

Reduction of fan noise by annulus boundary layer removal

Much of the noise produced by a fan is due to varying forces caused by the unsteady flow field through which it passes. In the absence of inlet guide vanes, support struts and other mechanical obstructions, the flow irregularities are caused by large scale intake turbulence and temporal and spatial variations in the annulus boundary layer. In many cases the most effective source is the annulus boundary layer, as this interacts with the blade tip which is the fastest part of the blade and is therefore the most effective noise source. This noise can be reduced if the annulus boundary layer is smoothly bled away from the rotor so that the blade tip sees a thin, uniform boundary layer. A one metre diameter ventilation fan has been run in a duct with and without a boundary layer bleed system and a reduction of far field sound power of approximately 5 dB has been obtained for a bleed of 5 % of the main flow with some far field tone reductions of more than 15 dB. Measurements of far field directivities, in-duct acoustic modes and aerodynamic distortions have confirmed the suggested explanation for this phenomenon. The tests have shown that great care must be taken in the design of a bleed system so that residual boundary layer distortions are not in the correct wavelength range to produce propagating acoustic waves at important blade passing harmonic frequencies. If such distortions are present, the far field noise can increase instead of decrease.

Bleed System Design Technology for Supersonic Inlets

A boundary-layer bleed system design procedure for supersonic inlets, with emphasis on the selection of bleed hole geometry, is described. Available experimental bleed hole performance data, coupled with bleed drag calculations, show that holes with shallow inclination are superior to holes normal to the surface in terms of over-all inlet performance. Recent test results from large-scale inlet models indicate that bleed hole size, bleed hole length, and boundary-layer velocity profile upstream of the bleed region are important parameters in the design of an effective and efficient bleed system. Nomenclature A = area A/A* = sonic area ratio Ca - drag coefficient D = bleed hole diameter HI - boundary-layer shape factor L/D = bleed hole length to diameter ratio m = mass flow M = Mach number P = pressure Q = sonic mass flow coefficient; ratio of actual mass flow to theoretical maximum mass flow at local total pressure and temperature X/D = spacing between rows of bleed holes Y/D - bleed hole spacing a = bleed hole inclination to inlet surface /8 = exit nozzle ramp angle 6* = boundary-layer displacement thickness rj = bleed efficiency factor Subscripts bl = bleed epl = bleed exit plenum / = cowl lip loc = local condition pi — bleed plenum se = sonic exit to — freestream stagnation condition 00 = freestream

Performance Estimates for a Supersonic Axisymmetric Inlet System

Results of recent tests of a large-scale axisymmetric mixed-compression inlet system designed for Mach number 2.65 indicate that the performance of the supersonic diffuser and required boundary-layer bleed system can be accurately estimated with relatively new analytic methods. Thus, many wind tunnel testing hours can be saved by avoiding much of the usual ?cut and try' wind tunnel development.