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Proceedings of the International Conference on Aerospace Science and Technology

26-28 June 2008, Bangalore, India

CHARACTERISTICS OF A SUPERSONIC AIR-INTAKE WITH BLEED

S. Das1 and J. K. Prasad2 1 Deptt of Space Engg & Rocketry, B. I. T, Mesra, Ranchi, India, [email protected] 2 Deptt of Space Engg & Rocketry, B. I. T, Mesra, Ranchi, India, [email protected]

ABSTRACT: A numerical investigation has been made to study the starting phenomena of a supersonic

air-intake with bleed. A mixed compression rectangular air-intake was designed using inviscid solutions.

Computational results with RANS solver using k-ω turbulence model, reveal the intake unstart, which

could be due to a large separation near the intake throat. It is observed that bleeding a small amount of air

at a location ahead of separation could start the intake. Studies were made for different bleeding rates.

Results indicate that bleed improves the flow quality in the diffuser duct and the performance.

Simulations with back pressure were also attempted and the basic flow structure is captured. Further the

effect of bleed is obtained for the back pressure. This suggests that boundary layer control through mass

flow removal avoids the starting problems of mixed compression intakes.

1. INTRODUCTION: Details of flow field inside air-intake is of vital importance to estimate the performance of air-breathing

engines operating in supersonic flight regime. Supersonic Air-intake is expected to compress the

freestream pressure to a level required for the engine. Various methods are adopted to accomplish this and

mixed compression mode is one of the method, which is generally adopted for operation at higher Mach

number due to various reasons. Overall efficiency of the mixed compression air-intake depends on the

level of pressure recovery, minimum flow distortion, and supply of desired mass flow rate. Critical issues

of such intakes are associated problems due to shock-boundary layer interactions leading to flow

separation, possible shock oscillations in the diffuser duct leading to intake-unstart, which might degrade

the performance of the intake. The control of flow separation in this shock-boundary layer interactionzone is generally accomplished through bleeding of a small percentage of air which is one of the most

sought method to improve the Intake performance. Bleed is a method by which a fraction of airflow is

removed through an orifice and helps to improve the quality of flow at a location where the flow stabilityis of prime importance. Other methods adopted to improve intake performance are wall perforation, fluid

injection, variable geometry, etc.

Improvement in starting characteristics of intake at supersonic speed is reported by Kubota et. al[1], witha bent cowl and a decrease in contraction ratio of the throat which could also improve the intake

performance. Van Wie et. al[2] studied the effect of Reynolds number, cowl length and height on the

starting characteristics of a rectangular intake at supersonic speed. The effect of fluid injection upstream

of unstarted intake was also investigated. Numerical investigation of different bleed models for a mixed

compression inlet is reported by Mizukami et. al[3]. Gawienowski [4] did experiments with different

bleed slot size and mass flow rates to assess the performance of an external compression intake at

supersonic speeds. Pressure recovery and distortion level were estimated and it was found that increasing

the bleed slot area as well as the mass flow, increases the intake performance. Selection of bleed hole

geometry and its inclination for an effective and efficient bleed system is reported by Syberg et. al[5].

Different bleed system at various locations on hypersonic intake is studied by Pandian et. al[6]. Shock

wave - boundary layer with bleed slot interaction studies are studied by Hamed et. al[7] and Kim et. al[8].

In the present study, numerical simulations are made using Fluent, on a typical 2D mixed compression

air-intake. Computations were performed using a Reynolds-Averaged Navier-Stokes solver with k-ω

turbulence model. Results indicate that k-ω turbulence model predicts an unstart which was not observed

with inviscid solutions. The purpose of the present study is to study the suppression of unstart by bleed

and to characterise the flow field with various bleeding rates and its effect on the overall flow field, with a

typical back pressure.

2. INTAKE MODEL:

A typical two-dimensional mixed compression air-intake was designed using inviscid analysis at Mach

2.2. The proposed intake has two compression ramps and two diffuser angles. Figure 1 shows the

geometry and also the computed flow field features obtained through inviscid simulations. The 1

st

and 2

nd

ramp angles are respectively 70 and 140 with respect to freestream direction. The diverging subsonic

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diffuser has a turning of 2.30 and 60 in downstream direction. The cowl inner surface is parallel to the

freestream direction. The overall length(L) of the model is 405.9mm with a capture height(hc) of 63.5mm.

3. COMPUTATIONS:Computations were made on this intake configuration using Compressible Reynolds – Averaged Navier-

Stokes equations with standard two-equation k-ω turbulence model available with FLUENT. ExplicitCoupled solver with upwind discretisation scheme for flow and transport equations was adopted. A

structured grid was created within the 2-D computational domain of interest. The minimum spacing in the

y-direction near the wall was kept as 0.15mm. A typical grid along with the extent of computational

domain is shown in Figure 2.

Supersonic pressure inlet boundary condition was specified at the inlet. At the exit, pressure outlet

boundary condition was assigned such that for supersonic outflow, all the variables were extrapolated

from the interior, and for subsonic outflow a typical pressure was specified at the exit. Non-reflecting

pressure outlet boundary condition was set at the top, meeting the cowl. No-slip boundary conditions

were specified at all the walls of the domain. All the computations were made either with supersonic exit

condition, which are called the free flow exit solution or with a back pressure, which are called the back pressure solution.

For all the computations, residuals of density and turbulent kinetic energy were monitored and suitable

convergence criterion was set for the solutions to converge. In addition net mass flux between inflow and

outflow boundaries and the y+ values at the ramp surfaces were monitored during the solution process.

4. RESULTS AND DISCUSSION:The present geometry of mixed compression air-intake, having two compression ramps and two diffuser

angles, designed for Mach 2.2, has been arrived at after making inviscid analysis. The actual flow field

inside the intake is governed by viscous effects and interactions, which may result to complex flow field

and even having the possibility of unstarts or buzz.

Compressible Viscous flow computations have been carried out using k-ω turbulence model. Simulation

initially indicated the focussed shocks at the cowl lip with shock reflections inside the subsonic diffuser.

The reflected shock from the cowl tip, impinges and interacts with the boundary layer near the throat,leading to a separation zone and expels the shock out of the throat. This leads to unstart of intake as seen

from the density contour presented in figure 3.

With an objective to start the intake, an attempt has been made to study the flow field by providing bleed

(% of captured mass flow) at a location upstream of the shock wave boundary layer interaction zone. This

indicated the suppression of intake unstart. Figure 4 shows the density contour inside the intake with

bleed of *** %, which clearly indicates start of intake. Further the effect of bleed on the flow field of

intake has been studied by making the simulation at different bleeds in the range of 0.6 to 2.4 % of

captured mass flow. Figure 5 shows the plot of Mach number at the location of throat at different bleed. It

is observed that, with increase in bleed, the throat Mach number reaches towards the design value. The

boundary layer profile downstream of the bleed and at the throat entry is shown in figure 6, where the

effect of bleed percentage on the profile indicates a better recovery for higher bleed mass flow rates.

Simulations have been also made by providing a back pressure at the exit of intake diffuser. The density

contours with back pressure (7.5 times freestream pressure), for different bleeds are shown in Figure 7.This shows the presence of normal shock at the diffuser and the separation on two walls. The effect of

bleed percentage can also be seen from the figures, which shows the different normal shock location in

the diffuser. Figure 8 shows the pressure distribution on the cowl surface with back pressure for various

bleed rates. The change in the location of normal shock is also seen with indication of steep rise in

pressure. Pressure recovery defined as ratio of total pressures, at the exit to freestream pressure is plotted

in figure 9, for different bleeds and for the inviscid case as well with a back pressure of P/Pinf=7.5. This

shows an improvement in pressure recovery with all the bleeds, however for different bleed rates the

change in pressure recovery is marginal. The simulations indicates the presence of complex flow field in

the intake and necessity of bleed to avoid unstart of intake. Increase in bleed rate could improve the flow

and hence the overall performance.

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hnology for Supersonic Inlets”, Journal of Aircraft, Vol.

10, No. 7, July 1973, pp. 407-413.

] Pandian, S., Jose, J., Patil, M. M and Srinivasa, P., “Hypersonic Air-Intake Performance Improvement Through

Different Bleed Systems”, ISABE-2001-1039.

[7] Hamed, A., Shih, S and Yeuan, J. J., “Investigation of Shock/Turbulent Boundary-Layer Bleed Interactions”,

Journal of Propulsion and Power, Vol. 10, No. 1, Jan-Feb 1994, pp. 16-24.

[8] Kim, H. D., Hingst, W. R and Davis, D. O., “Experimental Investigation of Crossing Shock Wave-Turbulent

Boundary Layer-Bleed Interaction”, AIAA-97-0608.

0.6 % Bleed 1.1 % Bleed

1.8 % Bleed 2.4 % Bleed

Figure 7: Density contour with different bleed for back pressure of P/Pinf = 7.5

Figure 9: Pressure RecoveryFigure 8: pressure distribution on Cowl

REFERENCES:

[1] Kubota, S., Tani, K and Masuya, G., “Aerodynamic Performances of a Combined Cycle Inlet”, Journal of

Propulsion and Power, Vol. 22, No. 4, July-August 2006, pp. 900-904.

[2] Van Wie, D. M., Kwok, F. T., and Walsh, R. F., 1996, “Starting Characteristics of Supersonic Inlets”, AIAAPaper 96-2914.

[3] Mizukami, M. and Saunders, J. D., “Parametrics on 2D Navier-Stokes Analysis of a Mach 2.68 Rectangular

Bifurcated Mixed Compression Inlet”, AIAA 95-2755, 1995.

[4] Gawienowski, J. J., “The Effect of Boundary-Layer Removal Through Throat Slots on the Internal Performance

of a Side Inlet at Mach Numbers of 2.0 and 2.3”, NASA TM-X-502.

] Syberg, J and Konesek, J. L., “Bleed System Design Tec[5

[6

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