BLEED SYSTEM - NASA bleed plenum arrangement, bleed flow ducting and exits, and the prediction of...
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BLEED SYSTEM .S INLET
ERONAUTICS A N D SPACE ADMINISTRATION WASHINGTON, D. C. JANUARY 1973
1. Report No. 2. Government Accession No.
NASACR 2187 4. Title and Subtitle
DESIGN OF A BLEED SYSTEM FOR A MACH 3.5 INLET
J. Syberg and T. E. Hickcox
9. Performing Organization Name and Address The Boeing Company P.O. Box 3707 Seattle, Washington 98 124
3. Recipient's Catalog No.
5. Report Date
6. Performing Organization Code
8. Performing Organization Report NO.
D6-60168 10. Work Unit No.
11. Contract or Grant No. I NAS2-6643
Boundary layer Analytic design procedure Bleed systems Supersonic Axisymmetric
14. Sponsoring Agency Code F
13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address Contractor ReDort
National Aeronautics and Space Administration Washington, D.C., 20546
19. Security Classif. (of this report) 20. Security Classif. (of this page)
I 15. Supplementary Notes
An analytical bleed system design procedure is applied t o a Mach 3.5, axisymmetric, translating centerbody inlet. The design features a traveling centerbody bleed system with a ducting arrangement separating low-, medium-, and high-pressure bleed. Three forward plenums and nine throat plenums are installed on the centerbody to meet the bleed requirements in the started Mach range between 1.6 and 3.5. The cowl contains four stationary plenums, three for forward and one for throat bleed. To achieve maximum bleed plenum pressure and thereby minimize bleed drag all bleed holes are inclined 20" to the surface except in the cowl throat region. Here the requirement of high stability margin with minimum total pressure recovery penalty resulted in 90' bleed holes. The bleed hole diameter varies from bleed plenum t o bleed plenum to achieve the most efficient boundary layer control, while the bleed exits are sized to operate at the highest possible plenum pressure without unchoking the bleed holes. Bleed flow rates, bleed plenum pressures, and boundary layer development along the cowl and centerbody are predicted over the entire started Mach range.
21. NO. of Pages 22. Price'
8 8 $3.00
18. Distribution Statement I 17. Key Words (Suggested by Author(sl ) I Inlets Bleed holes
* For sale by the National Technical Information Service, Springfield, Virginia 22151
DESIGN OF A BLEED SYSTEM
FOR A MACH 3.5 INLET
By J . Syberg and T. E. Hickcox The Boeing Company
A bleed system has been designed for a Mach 3.5 inlet using analytical procedures which promise t o save many wind tunnel testing hours. The design is based on a model size compat- ible with the existing NASA Ames 1/3scale inlet model and on freestream conditions cor- responding to those in the NASA Ames 8- by 7-ft supersonic wind tunnel.
This report first describes the analytical procedures and tools used in the bleed system design. The operating characteristics of the inlet are then presented in the form of surface pressures and boundary layer development without boundary layer control. The proposed bleed systems for centerbody and cowl are then described, along with the design considera- tions for Mach 3.5 and off-design conditions. Next, the boundary layer growth is predicted throughout the started Mach number range with the bleed system operating. Finally, possible problem areas are identified and the overall system is evaluated.
Application of the analytical design procedure indicated the need for a bleed system with unusual and complex features that would be unlikely to evolve from normal wind tunnel development testing. The centerbody has a “traveling” bleed system with a ducting arrange- ment separating low-, medium-, and high-pressure bleed regions. Three forward bleed plenums and nine throat plenums are included, with the novel feature that one forward plenum is located among the throat plenums to provide sufficient forward bleed at off-design Mach numbers. The cowl contains four stationary plenums, three for forward bleed and one for throat bleed. The exit areas for the various bleed plenums are sized to operate at the highest possible plenum pressures without unchoking the bleed holes. All centerbody bleed holes and the forward cowl bleed holes are inclined 20” t o the surface, primarily to achieve maximum bleed plenum pressure and thereby minimize bleed drag. The cowl throat bleed holes are normal t o the surface to maximize the normal shock stability margin. The bleed hole diameter varies from bleed plenum to bleed plenum to achieve the most efficient boundary layer control.
With the proposed system the bleed flow rate is 13.3% of the lip flow (0.133 WL) at Mach 3.5, with 0.089 WL forward bleed and 0.044 WL throat bleed. Analysis indicates that this amount of forward bleed is near the minimum bleed rate required to achieve attached flow in the supersonic diffuser and provide a full boundary layer velocity profile in the throat ahead of the normal shock. Isentropic compressions and oblique shock wave reflections at off-design Mach numbers are in general well controlled. Boundary layer separation at oblique shock reflections is predicted at some off-design Mach numbers, but rapid reattachment is expected in all of these cases due t o favorable downstream pressure gradients coupled with high bleed rates near these critical regions.
The development of bleed systems for boundary layer control in supersonic inlets has in the past depended mainly on extensive wind tunnel tests. The tests were complex, time con- suming, and did not always allow optimization of the system. An analytical procedure (ref. 1 ) has now been developed for the design of bleed systems based on theoretical analysis and extensive experimental data. Using the procedure allows analytic definition of a bleed system, which may then be optimized in the wind tunnel with less testing.
The present work applies this procedure t o the design of a bleed system for a Mach 3.5 inlet with the objective of providing satisfactory operation across a wide range of “started” Mach numbers with adequate tolerance to transient disturbances in upstream Mach number and angle of incidence, and downstream corrected weight flow. The work included the design of internal cowl and centerbody contours, design of the bleed pattern, bleed hole geometry, bleed plenum arrangement, bleed flow ducting and exits, and the prediction of bleed system performance. The inlet and bleed system geometry is sized for an existing 1/3-scale supersonic inlet model (ref. 2) operating within the flow conditions available in the NASA Ames 8- by 7-ft supersonic wind tunnel. The present work was carried out under contract NAS 2-6643 (ref. 1).
Portions of the procedure and technology have been applied before (refs. 2 and 3), but never as an entire package. The present application of the procedure is the first time a com- pletely analytical design of the bleed system has been done prior to model testing. As such, a thorough validation of the procedure is needed from wind tunnel tests of the 1/3-scale inlet.
bleed hole area
cowl lip projected area
sonic area ratio
bleed hole diameter
wall temperature recovery-wall temperature/freestream total temperature
boundary layer shape factor-6*/8
aspect ratio of bleed hole-length/diameter
local Mach number
boundary layer power law velocity profile exponent
bleed hole mass flow coefficient-actual flow/ideal choked flow at same conditions
cowl lip radius
circumferential bleed hole spacing, Y/D
axial velocity component
theoretical freestream airflow through an area equal t o the cowl lip area
nondimensionalized bleed flow rate
nondimensionalized axial coordinate
AX/RL nondimensional forward translation of centerbody from Mach 3.5 position
Y half of the circumferential center-to-center spacing between holes in one row of bleed holes
Y boundary layer coordinate normal to wall
CY bleed hole angle relative to surface
6 boundary layer thickness
6" boundary layer displacement thickness
e boundary layer momentum thickness Subscripts
t h throat
TO, T, freestream total
property derived from analysis of Reyhner
property derived from analysis of Sasman and Cresci
1 upstream side of oblique shock reflection or normal shock
downstream side of normal shock
downstream side of oblique shock reflection
The first step in the bleed system design procedure for a new inlet contour is t o deter- mine a centerbody translation schedule versus freestream Mach number that will provide effi- cient internal compression as well as adequate tolerance against transient changes in Mach number and angle of incidence. These conflicting requirements can be met by maintaining the t