David O. Davis, Manan A. Vyas, and John W. Slater

Glenn Research Center, Cleveland, Ohio

Research on Supersonic Inlet Bleed

NASA/TM—2012-217620

August 2012

AIAA–2012–0272
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David O. Davis, Manan A. Vyas, and John W. Slater

Glenn Research Center, Cleveland, Ohio

Research on Supersonic Inlet Bleed

NASA/TM—2012-217620

August 2012

AIAA–2012–0272

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

50th Aerospace Sciences Meeting

sponsored by the American Institute of Aeronautics and Astronautics

Nashville, Tennessee, January 9–12, 2012
Acknowledgments

The authors would like to thank R.C. Clark, B.T. Seifert, R.G. Senyitko, J.M. Lucero, D.W. Hammett, C. Horn, and S.M. Velez 

of NASA Glenn Research Center for matters concerning the wind tunnel and gas bench operation. Funding from the Supersonics 

Project of the NASA Fundamental Aeronautics Program is gratefully acknowledged.

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Available electronically at http://www.sti.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2012-217620 1 

Research on Supersonic Inlet Bleed 

David O. Davis, Manan A. Vyas, and John W. Slater 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Phase I data results of the Fundamental Inlet Bleed Experiments project at NASA Glenn Research 

Center (GRC) are presented which include flow coefficient results for two single-hole boundary-layer 

bleed configurations. The bleed configurations tested are round holes at inclination angles of 90 and 20

both having length-to-diameter ratios of 2.0. Results were obtained at freestream Mach numbers of 1.33, 

1.62, 1.98, 2.46, and 2.92 and unit Reynolds numbers of 0.984, 1.89, and 2.46107

/m. Approach 

boundary-layer data are presented for each flow condition and the flow coefficient results are compared to 

existing multi-hole data obtained under similar conditions. For the 90 hole, the single and multi-hole 

distributions agree fairly well with the exception that under supercritical operation, the multi-hole data 

chokes at higher flow coefficient levels. This behavior is also observed for the 20 hole but to a lesser 

extent. The 20 hole also shows a markedly different characteristic at subcritical operation. Also 

presented are preliminary results of a Computational Fluid Dynamics (CFD) analysis of both 

configurations at the Mach 1.33 and a unit Reynolds number of 2.46107

/m. Comparison of the results 

shows the agreement to be very good. 

Nomenclature 

Ab Bleed hole area 

Amin Minimum area of bleed orifice 

D Bleed hole diameter 

gc Proportionality constant 

Hinc Incompressible shape factor 

L Bleed hole length 

M Mach number or molecular weight 

Mblk Nominal Mach number of wind tunnel fixed geometry convergent-divergent nozzle 

Mnoz Actual Mach number of wind tunnel fixed geometry convergent-divergent nozzle 

ppm parts per million (volume fraction) 

P Static pressure 

Pt Total pressure 

Q Sonic flow coefficient 

R Gas constant 

Re Unit Reynolds number 

Re Scaled Reynolds number (Re  1.0E-07 m) 

sccm standard cubic centimeters per minute 

slm standard liters per minute 

t Bleed plate thickness 

T Static temperature 

Tt Total temperature 
NASA/TM—2012-217620 2 

U Streamwise velocity component 

w Mass-flow rate 

x,y,z Cartesian coordinate system, x relative to nozzle exit plane 

xwin Coordinate relative to Test Section #1 top window leading edge 

Greek Symbols: 

 Bleed hole inclination angle 

 Boundary-layer thickness 

1 Boundary-layer displacement thickness 

 Ratio of specific heats 

 Volume fraction of gas 

Subscripts: 

0 Pertaining to the wind tunnel plenum chamber condition 

air Pertaining to air gas 

e Pertaining to boundary-layer edge condition 

i Pertaining to ideal conditions 

met Pertaining to methane gas 

plen Pertaining to the bleed plenum condition 

wall Pertaining to wall condition 

Superscripts: 

* Pertaining to choked conditions 

Introduction 

An ongoing effort at NASA Glenn Research Center (GRC) has been the development of improved 

bleed models for Computational Fluid Dynamics (CFD) codes used to aid in the design of supersonic 

inlets. Accurate prediction of the bleed orifice flow coefficient which relates bleed plenum pressure to 

mass flow removed is important to predicting inlet performance as well as estimating bleed drag. Much of 

the flow coefficient data at conditions of interest to inlet designers is based on bleed plates with multiple 

rows of holes. The flow coefficient for these plates is typically presented as a function of bleed plenum 

pressure normalized by the freestream total pressure. Further, the flow coefficient is typically referenced 

to the ideal sonic mass flow which is also based on the freestream total conditions. The use of freestream 

conditions is suitable for undisturbed core flow, but bleed regions are necessarily placed in the vicinity of 

shock waves where the freestream condition may not be relevant to the local conditions at the orifice 

entrance. Thus, a model that relates the flow coefficient to local conditions is desirable. Slater (Ref. 1) 

recently presented the results of such a model for 90 bleed holes based on the experimental data of Willis 

et al. (Ref. 2). The use of multi-hole data to develop flow coefficient models that are applied to individual 

holes based on local conditions may be misleading due the interaction of adjacent holes. Davis (Ref. 3) 

has shown that the overall flow coefficient of two adjacent normal holes at Mach 2.46 can vary be by as 

much as 6 percent at choked conditions depending on the orientation of the holes to one another. The 

conclusion to that study was that it was anticipated that multi-hole configurations would tend to exhibit 

higher flow coefficient levels than their single-hole counterpart. Reported herein are the results of singlehole bleed flow coefficient experiments and their comparison to the multi-hole data of Willis et al. 
NASA/TM—2012-217620 3 

Fundamental Inlet Bleed Experiments (FIBE) 

The Fundamental Inlet Bleed Experiments (FIBE) project is a research effort at NASA GRC to 

provide a comprehensive experimental database and a better understanding of how bleed systems can be 

improved through better modeling, alternate bleed configurations, and bleed placement within supersonic 

and hypersonic inlets. A number of events lead to the initiation of the project and are summarized as 

follows: 

 June 1992—Inlet/Nozzle Technology for the 1990s Workshop was held at the Ohio Aerospace 

Institute (OAI). 

○ An outcome of that workshop was the need for a Bleed Specific Workshop 

 September 1993—Inlet Bleed Technology Workshop was held at OAI. 

○ Bleed research needs were identified. 

 1994 to 1997—NASA LeRC had an active bleed research program primarily in the 11 ft Supersonic 

Wind Tunnel (SWT) but also the 1515 cm SWT. 

○ A summary of this research was presented by Davis (Ref. 4) at the 1st Annual Shock 

Wave/Boundary Layer Interaction (SBLI) Workshop (2008). 

 2007—The Inlet and Nozzle Branch at NASA GRC made a decision to re-establish some of the bleed 

research capability in the 1515 cm SWT. 

 2011—Fundamental Inlet Bleed Experiments project initiated. 

The 1515 cm SWT was chosen for initial research due to its flexible test section and lower operating 

costs than some of GRC’s larger tunnels. Due to limited bleed vacuum capacity of the 1515 cm SWT, 

the research conducted in the 1990s was limited to single or small array orifice studies. In addition to 

detailed flow field and surface measurements in the vicinity of a bleed orifice (Refs. 5 to 8) flow 

coefficient data for various hole configurations were also obtained (Ref. 3). Since the initial testing under 

the FIBE project is focused on the flow coefficient behavior of single-hole configurations, the 1515 cm 

SWT is ideally suited. 

FIBE Phase Planning 

The FIBE project is planned to be conducted in three phases. Phase I will include facility checkout 

and acquisition of flow coefficient data using pre-existing test hardware. The checkouts will include 

verification of the recently upgraded trace gas system for mass-flow measurement (described below) as 

well as establishing the operating envelope of the current bleed system. This will include evaluation of the 

sample pump capacity, system response time and bleed vacuum capacity (mechanical Stokes pump and 

altitude exhaust). Also during Phase I, approach boundary-layer profiles will be established for each 

Mach number at several Reynolds numbers. Then, single-hole flow coefficient data will be obtained using 

existing 90 and 20 bleed holes for comparison with existing multi-hole data. 

Phase II will be an extension of Phase I and will investigate single-hole flow coefficient behavior 

including the effects of: 

 Hole inclination angle () 

 Hole length to diameter ratio (L/D) 

 Hole diameter to boundary-layer displacement thickness ratio (D/1) 

 Hole shape and/or entrance contouring 

 Detailed flow field measurements 

These results will be used to validate CFD results as well as to develop improved bleed models. 
NASA/TM—2012-217620 4 

Phase III will study multi-hole bleed regions with and without shock waves present. These results will 

be used to validate CFD results under inlet-like flow conditions. These studies will require upgrades to 

the existing bleed system of the 1515 cm SWT which will include larger bleed lines and increased 

vacuum capacity through the installation of an ejector system. 

Facilities 

Phase I and II testing will be performed in the NASA GRC 1515 cm Supersonic Wind Tunnel 

(SWT), which is described below. Phase III testing will be performed in either the 1515 cm SWT or the 

NASA GRC 11 ft SWT, depending on the scale required to meet the test objectives. The 11 ft SWT 

operates in a very similar manner to the 1515 cm SWT, but also has a 450 psig ejector system as a bleed 

vacuum source. In addition to unit bleed studies, the 11 ft SWT is also large enough for small inlet 

models that may incorporate bleed. The potential also exists to test bleed configurations in a large scale 

40/60 axisymmetric mixed compression inlet in the 1010 ft SWT. This test would be combined with an 

already proposed test to evaluate micro-vortex generators. 

FIBE Phase I Test Program 

FIBE Phase I testing began on April 29, 2011 and concluded on September 13, 2011. Based on the 

data obtained, improvements to the system were identified and a number of changes were made. Some of 

these improvements included complete re-alignment of the wind tunnel components, probe z-traverse 

mechanism alignment and sealing improvements, upgrade of the sample pumps to higher capacity, as well 

as some operational procedures to reduce data acquisition time. 

Test Hardware 

A schematic of a single bleed hole with reference parameters is shown in Figure 1. In this figure, we 

distinguish between a thick plate and a thin bleed plate. A thick bleed plate, which is illustrated in 

Figure 1(a), is defined as one in which the minimum flow area is equal to the area normal to the hole axis. 

If the diameter of the hole is maintained constant and the thickness of the plate is decreased, a point is 

reached where further reduction in the thickness results in a condition where the minimum flow area is 

not normal to the hole axis. This condition is defined as a thin bleed plate and is illustrated in Figure 1(b). 

The normalized thickness which separates a thick from a thin plate occurs when: 

cos

D

t (1) 

 

 a) “Thick” bleed plate b) “Thin” bleed plate 

Figure 1.—Schematic of bleed holes with reference parameters. 
NASA/TM—2012-217620 5 

Thus, for a round bleed hole in a “thick” plate, the minimum flow area is circular with the area defined as: 

2

thickmin, 4

A b DA   



 





cos

D

t (2) 

and, for a “thin” plate the minimum flow area is elliptical with the area defined as: 

2

2

thinmin, sin

cos1





























  

 



  D

t

D

t AA b 



 





cos

D

t (3) 

Note that as α → 90 for a thin plate, t/D → 0 and, as expected, any plate thickness is considered thick for 

normal holes. A plot of the normalized minimum flow area of a bleed hole as a function of normalized 

bleed plate thickness for various hole inclination angles is shown in Figure 2. Finally, the normalized 

length of the bleed hole is given by: 

  sin

1

D

t

D

L (4) 

Figure 2.—Area ratio versus normalized plate thickness. 

A /A min b
NASA/TM—2012-217620 6 

Figure 3.—Phase I test article configurations. 

For the Phase I testing, three pre-existing single-hole test articles were used to check out the bleed 

system operation. The test articles consist of one 90 (normal) hole, and two 20 hole configurations as 

shown in Figure 3. The bleed holes are machined into interchangeable 50.8 mm (2.0 in.) diameter round 

plugs that are installed into the upper window of the wind tunnel test section. The 90 (C01) and one of 

the 20 (C02) bleed holes have a nominal hole diameter of 6.0 mm and a length-to-diameter (L/D) ratio of 

2.0. The other 20 bleed hole (C03) has a nominal diameter of 5 mm and an L/D ratio of 2.92. Note that 

configuration C02, satisfying Equation (3), is considered to be a “thin” bleed plate and for its t/D = 0.684, 

the minimum flow area is 24.8 percent above the bleed hole area. It should also be noted that at shallow 

bleed hole inclination angles, the actual minimum area may be significantly higher due to breaking or 

radiusing of the sharp hole edge. 

Test Facility 

The FIBE Phase I testing was conducted in the 1515 cm SWT, which is a continuous flow facility 

with Mach number set by interchangeable fixed-geometry convergent-divergent (C-D) nozzle blocks. The 

current Mach number capability is M = 0.3 to 0.8 for subsonic operation and M = 1.4, 1.7, 2.0, 2.5, and 

3.0 for supersonic operation. The total temperature is nominally ambient. The wind tunnel is supplied 

with lab-wide 40 psig combustion air and exhausts to lab-wide Altitude Exhaust which is maintained 

below 13.8 kPa (2.0 psia). A sketch of the 1515 cm SWT is shown in Figure 4. 

Test Section #1 is configured to accept the test articles as shown in Figure 3. The top window insert 

of Test Section #1 is machined to accept the test articles shown in Figure 5. A plate is attached on the top 

of the window insert to form the bleed plenum. The bleed flow exhausts through a male barb fitting that 

screws into the plenum plate. The side window inserts have schlieren glass installed for flow visualization 

and the bottom window insert has three rows of axial pressure taps. 

C01

D=6.010mm

=90°

L/D=2.0

t/D=2.0

C02

D=6.029mm

=20°

L/D=2.0

t/D=0.684

C03

D=5.018mm

=20°

L/D=2.92

t/D=1.0
NASA/TM—2012-217620 7 

Figure 4.—1515 cm Supersonic Wind Tunnel. 

Figure 5.—Bleed setup in test section #1. 

Instrumentation 

The instrumentation consists primarily of pressure measurements to monitor the flow conditions 

approaching and within the test section and bleed plenum, and those associated with the trace-gas system 

to measure the bleed flow rate. 

Test Section Instrumentation 

The flow condition entering the test section is determined from measurements of total pressure, total 

temperature, and dew point inside the wind tunnel plenum and with wall static pressure taps located near 

the exit plane of the C-D nozzle. Local conditions in the vicinity of the test article are measured with a 

translatable Pitot probe and wall static pressure taps on the upper and lower test section window inserts. A 

view of the flow side of the upper test section window insert is shown in Figure 6. With reference to 

Figure 6, 18 surface static taps are located in the window insert and two are located on the test article 

itself. The taps are located along the centerline (z = 0) and at z = 2.22 cm (0.875 in.). The numbers 

beside the taps denote the port number of ESP pressure module number 5 that the taps were connected to. 
NASA/TM—2012-217620 8 

The Pitot probe shown in Figure 6 is remotely actuated in the y and z-directions, but must be manually 

positioned in the x-direction (streamwise). For Phase I testing, only approach profiles in a transverse 

plane coinciding with the second row of static taps (x = 2.700 cm (1.065 in.)) were obtained. The tip of 

the Pitot probe is made from 0.508 mm (0.020 in.) tube flattened to 0.3302 mm (0.0130 in.). The 

approximate position of the probe relative to the wall is established by an electrical touch circuit. The 

axial location of where the bleed hole centerline intersects the wall surface relative to the window leading 

edge varies depending on the bleed configuration and is shown in Figure 7. 

Figure 6.—Upper test section window insert instrumentation. 

Figure 7.—Axial location of bleed holes. 

M5P(2)

7

14

17

8

M1P(20)

21

20

3 4

5 6

15 16

18 19

9 10

11 12 13

z

y

x win

C03

C02

C01

B.L. Measurement Plane

2.700

6.048

6.276

7.620

Dimensions in Centimeters
NASA/TM—2012-217620 9 

Bleed Mass-Flow Measurement 

The bleed mass-flow rate is measured by the method of mixtures (Refs. 9 and 10) where the mixer 

gas is methane and its concentration is measured with a Flame Ionization Detector (FID). A schematic of 

the setup is shown in Figure 8. The bleed flow exits the bleed plenum through the male barb fitting 

installed in the plenum cover. Through a small port in the side of the barb fitting, methane is injected 

(Teledyne Hastings model HFC-302, 0 to 100 sccm) into the bleed flow at a known flow rate. The 

methane mixes with the bleed flow and a sample is delivered to the FID (Signal Instruments 3000 HM, 

0 to 1000 ppm) which measures the methane concentration. Sample concentrations are typically 

maintained between 500 to 1000 ppm. The bleed flow rate can then be calculated from the following 

relation: 

met

met

met

met

w

M

M

w  

















  













  air 1

air (5) 

The bleed flow can also be routed through a calibrated electronic flow meter (Omega model FMA 

875-V, 0 to 200 slm range) which is used for measuring low bleed flow rates. Due to the low pressure in 

the bleed line, the flow meter will choke before the bleed hole does. The planned procedure for 

establishing a flow coefficient as a function of plenum pressure is to measure the bleed flow with the flow 

meter until it chokes. At this point the methane flow is established and the Omega flow meter bypass 

valve is opened. The bleed flow is then measured by the method of mixtures until the bleed hole chokes. 

Figure 8.—The 15x15 cm SWT test schematic. 
NASA/TM—2012-217620 10 

TABLE 1.—TEST OPERATING CONDITIONS (RE = RE 1.0E-07 m) 

Cond. no. Mblk Re Mnoz Me 

(mm) 

1

(mm) 

Hinc

001 1.4 0.984 1.381 1.328 14.0 2.68 1.333 

002 1.4 1.89 1.387 1.334 14.0 2.55 1.324 

003 1.4 2.46 1.389 1.333 13.2 2.37 1.323 

004 1.7 0.984 1.671 1.617 12.5 2.75 1.341 

005 1.7 1.89 1.677 1.625 13.7 2.50 1.305 

006 1.7 2.46 1.679 1.618 12.3 2.41 1.320 

007 2.0 0.984 1.987 1.954 12.4 2.97 1.341 

008 2.0 1.89 1.992 1.966 12.2 2.77 1.326 

009 2.0 2.46 1.993 1.971 12.1 2.66 1.320 

010 2.5 0.984 2.493 2.455 14.2 3.92 1.344 

011 2.5 1.89 2.503 2.461 13.4 3.57 1.320 

012 2.5 2.46 2.506 2.463 13.1 3.44 1.311 

013 3.0 0.984 3.015 2.918 13.3 4.28 1.338 

014 3.0 1.89 3.031 2.956 12.5 3.99 1.316 

015 3.0 2.46 3.035 2.960 12.2 3.78 1.306 

Operating Conditions 

Data were accumulated at nominal Mach numbers of 1.4, 1.7, 2.0, 2.5, and 3.0. At each Mach 

number, data were obtained at Reynolds numbers of 0.984, 1.89, and 2.46E+07/m (3.00, 5.75, and 

7.50E+06/ft). Thus, there were a total of 15 different conditions run. For bookkeeping purposes, a 

sequential number was assigned to each condition, which is summarized in Table 1. Also in Table 1 is the 

actual Mach number calculated at the nozzle exit plane as well as the local boundary-layer edge Mach 

number measured at the boundary-layer measurement plane shown in Figure 7. The distance from the 

nozzle exit to the boundary-layer measurement plane is 36.36 cm (14.32 in.) so that the edge Mach 

number is slightly below the nozzle exit Mach number. 

Approach Boundary-Layer Profiles 

Approach boundary-layer profiles were taken at an axial position of xwin = 2.700 cm (1.065 in.). This 

station coincides with the second axial row of pressure taps shown in Figure 6. For data reduction 

purposes, the static pressure was assumed to be constant across the boundary layer and equal to the 

average of the two taps on either side of the centerline (taps labeled M5P(9) and M5P(15) in Figure 6). 

The center tap, M5P(3), was directly under the probe and is influenced by the presence of the probe until 

it is well away from the wall. Beyond the point where the tap is no longer influenced, which is a function 

of Mach number, the center tap was in excellent agreement with the two outboard taps. 

The measured boundary-layer profiles are plotted in terms of Mach number, normalized velocity, and 

in van Driest scaled variables (Ref. 11) (law-of-the-wall) in Figure 9. As expected, the profiles become 

fuller as the Reynolds number is increased indicating a reduction in displacement thickness. There are 

also two things to note about the Mach 3.0 profiles. First, is that the core flow is not uniform, and second, 

that the Re = 0.984 case does not recover to the same core Mach number as the higher Reynolds number 

cases. The flow quality of the Mach 3.0 block will be investigated further in the future. All of the profiles 

exhibit law-of-the-wall behavior indicating fully-turbulent boundary layers. 

The calculated boundary-layer thickness, displacement thickness and incompressible shape factor for 

each operating condition is tabulated in Table 1. Note for current test articles, the bleed-hole diameter-todisplacement thickness (D/1) is nominally D/1 ~ 2, whereas traditionally for bleed systems this ratio is 

chosen to be closer to unity. The boundary-layer edge Mach number, displacement thickness, and 

incompressible shape factor are plotted as a function of nozzle Mach number for the various Reynolds 

numbers in Figure 10. The boundary-layer edge Mach number is largely independent of Reynolds number 

except for the aforementioned lower core Mach number for the lowest Reynolds number case at 

Mach 3.0. The displacement thickness decrease with increasing Reynolds number is as expected. 
NASA/TM—2012-217620 11 

a) Mach number profiles. 

b) Normalized velocity profiles. 

c) Van driest scaled profiles. 

Figure 9.—Approach boundary-layer profiles. 

M =1.4 1.7 2.0 2.5 3.0 blk

Re'=0.984

Re'=1.89

Re'=2.46

M =1.4 1.7 2.0 2.5 3.0 blk

Re'=0.984

Re'=1.89

Re'=2.46

M =1.4 blk

2.5

3.0

1.7

2.0

Re'=0.984

Re'=1.89

Re'=2.46

a) Boundary-layer edge Mach number. 

b) Displacement thickness. 

c) Incompressible shape factor. 

Figure 10.—Approach boundary-layer parameters. 

1.0

1.5

2.0

2.5

3.0

3.5

1.0 1.5 2.0 2.5 3.0 3.5

Boundary‐Layer Edge Mach Number

Nozzle Mach Number

Re=0.984E+07 /m

Re=1.89E+07 /m

Re=2.46E+07 /m

0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

4.5

1.0 1.5 2.0 2.5 3.0 3.5

Boundary‐Layer Displacement Thickness, mm

Nozzle Mach Number

Re=0.984E+07 /m

Re=1.89E+07 /m

Re=2.46E+07 /m

1.00

1.05

1.10

1.15

1.20

1.25

1.30

1.35

1.40

1.45

1.50

1.0 1.5 2.0 2.5 3.0 3.5

Incompressible Shape Factor

Nozzle Mach Number

Re=0.984E+07 /m

Re=1.89E+07 /m

Re=2.46E+07 /m
NASA/TM—2012-217620 12 

Although some dependence on Mach number was anticipated, the S-shaped trend with Mach number was 

not expected and is not fully understood. The incompressible shape factor is relatively independent of 

Mach number and varies with Reynolds number in a manner expected. The nominal value of 1.32 for all 

the profiles is consistent with a fully-turbulent compressible boundary layer. 

Flow Coefficient Data 

Flow coefficient data were obtained with the C01 and C02 configurations at the conditions listed in 

Table 1. Flow coefficient data are typically presented in terms of a sonic flow coefficient defined as: 

* wi

w Q  (6) 

where the numerator is the measured mass-flow rate and the denominator is the ideal mass-flow rate 

under choked conditions: 

   

 





 















 



       12

1

air 2

1

R

g

T

AP

w c

e,t

* be,t

i (7) 

The sonic flow coefficient is typically presented as a function of the ratio of bleed plenum pressure to 

freestream (boundary-layer edge) total pressure (Pplen/Pt,e). The total temperature at the boundary-layer 

edge (Tt,e) is assumed to be the same as the total temperature in the wind tunnel plenum chamber (Tt,0). 

With reference to Figure 8, a typical flow coefficient curve was obtained by initially passing all the 

bleed flow through the calibrated electronic flowmeter. The bleed control valve is opened and data 

recorded until the flow meter chokes. At this point, the bleed control valve is closed slightly, the bypass 

valve is opened, and methane flow is established. The measurements are then continued with the trace gas 

system by opening the bleed valve, recording the methane flow rate and concentration, and then 

calculating the mass flow via Equation (5). The bleed control valve was typically backed-off so as to 

establish a measure of overlap from the two measurement techniques. 

Initial checkout of the trace gas system was performed at Mach zero conditions for the three bleed 

configurations as shown in Figure 11. In this figure, the solid symbols are the results from the electronic 

flowmeter and the open symbols are the results from the trace gas system. Under these conditions, the 

agreement between the two methods is observed to be very good. All three configurations initially exhibit 

similar behavior as the plenum pressure is reduced. With reference to Figure 7, the C03 configuration, 

which turns the flow the full 70 to flow normal to the minimum area, lags behind C01 and C02 

configurations, but eventual chokes at the same level (Q~0.8) as the C01 configuration. The C02 

configuration, which turns the flow 57 to flow normal to the minimum area, follows the C01 closely then 

deviates where it chokes at a flow coefficient near unity. The approximately 25 percent higher flow 

coefficient is consistent with the minimum area being about 25 percent higher than the hole area for 

t/D = 0.684 and  = 20. 

Under flow conditions, the agreement between the two methods was not as good. Figure 12 shows 

flow coefficient distributions for the C01 configuration at Mblk = 1.4 at the three Reynolds numbers tested, 

although all Mach numbers tested displayed similar behavior. The overlaid curves to the far right show 

the trace gas results for all three Reynolds numbers. The flowmeter and trace gas results are then 

presented to the left on shifted scales so the difference between the two techniques can be discerned. 
NASA/TM—2012-217620 13 

Figure 11.—Flow coefficient for Mach zero condition. 

Figure 12.—Comparison of electronic flowmeter and trace gas system for C01 at Mblk = 1.4. 

0.00

0.20

0.40

0.60

0.80

1.00

0.00 0.20 0.40 0.60 0.80 1.00

Q

Pplen/Pt,e

C01 Flowmeter

C01 Trace Gas System

C02 Flowmeter

C02 Trace Gas System

C03 Flowmeter

C03 Trace Gas System

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

0.18

0.20

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70

Q

Pplen/Pt,e

Re'=0.984 Flowmeter

Re'=0.984 Trace Gas

Re'=1.89 Flowmeter

Re'=1.89 Trace Gas

Re'=2.46 Flowmeter

Re'=2.46 Trace Gas

0.40 0.40            0.40
NASA/TM—2012-217620 14 

For the case presented, trace gas results were obtained which fully overlapped the electronic flowmeter 

results. The first thing of note is that the electronic flowmeter reads a slight positive flow when put under 

vacuum conditions even when there was no bleed flow. These points are the solid symbols furthest to the 

right in the plots. Whereas the trace gas results (far right curves) are relatively independent of Reynolds 

number, at least in the lower flow range, the electronic flowmeter shifts increasingly to the right as the 

Reynolds number is reduced (decreasing bleed line pressures). This suggests that there is a Reynolds 

number effect on the calibration of the flowmeter. Prior to testing, the flowmeter was calibrated at 

approximately between one and two atmospheres. Under flow conditions the bleed line pressure varied 

between 1.7 to 55 kPa (0.25 to 8.0 psia). The reason for using the electronic flowmeter was to avoid 

having to use multiple range methane flow controllers to resolve the lower flow rates as well as the 

improved response time over the trace gas system at low flow rates. Two improvements will be 

considered for future testing. First, multiple methane flow controllers of different ranges that can be 

remotely switched in and out of the circuit will be added to ensure that the methane flow rate is 

appropriate for the bleed flow rate (resulting concentration between 500 to 1000 ppm), and second, the 

electronic flow meter, which was calibrated at above atmospheric conditions, will be calibrated under 

vacuum conditions. Inasmuch as it appears that the electronic flowmeter was being operated outside the 

range of its calibration, only the trace gas results will be presented from here forward. This means that the 

lower flow rates won’t be resolved as well as they otherwise might have been. 

Subsequent to the accumulation of the current flow coefficient data, flow coefficient data at 

Mach 2.46 from Bodner et al. (Refs. 5 and 6) was located. In this study, Bodner used the same bleed 

plugs in the same facility but used a sidewall of the tunnel rather than the top wall of the tunnel. The same 

electronic flowmeter/method-of-mixtures technique was used to obtain his data, but the gas bench was 

comprised of all different components. A comparison of these data with the current data at M = 2.46 is 

shown in Figure 13. For the C02 case (Fig. 13(b)), the data are in very good agreement. For the C01 

configuration (Fig. 13(a)), the trends are the same but Bodner’s curve has a higher slope than the present 

data. After this discrepancy was discovered, the present C01 data were repeated with the same results. 

Inasmuch as the present data were carefully repeated and reviewed with no anomalies discovered, it is 

assumed here that the present data are correct. 

The flow coefficient distributions for all of the operating conditions are shown in Figure 14. With 

reference to Figure 14(a), a feature of note is the tail-up behavior at the left end of the curves, primarily at 

the lower Reynolds numbers. This behavior is attributed to the inability of the sample pump to provide 

the sample at a sufficient flow rate to the FID unit. Much of the data were accumulated with a single stage 

vacuum pump that failed midway through the investigation. It was replaced temporarily with a borrowed 

two-stage pump, and then ultimately a new four stage pump was installed. There was not time to repeat 

these curves with the four stage pump, but some spot checks of the data indicated that the problem will be 

resolved at all conditions but the low Reynolds number cases at Mach 2.5 and 3.0. 

With reference to Figure 14(b), the behavior of the 20 hole is markedly different than the 90 hole. 

At zero flow rates, there is a slight pressurization of the bleed plenum. At low flow rates, the plenum 

pressure is observed to be invariant with flow rate. As the flow approaches a choked condition (upper part 

of the curves), the lowest Reynolds number data are observed to ride below the higher two cases except 

for the lowest Mach number case (Mblk = 1.4). For this case only, the curve is shifted to the right. The 

reason for the shift is not yet known, but it is repeatable. 

The low Reynolds number data were taken to exercise the system over as large of a range as possible, 

but practically speaking it isn’t as of much interest as the higher Reynolds number data. If we exclude the 

low Reynolds number data for both the C01 and C02 configurations, then the remaining two Reynolds 

number cases agree fairly well. While there is no definite trend observed for the C01 configuration, the 

C02 configuration consistently shows slightly higher flow rates as the Reynolds number increases. 
NASA/TM—2012-217620 15 

a) C01 bleed configuration (90 hole) 

b) C02 bleed configuration (20 hole) 

Figure 13.—Comparison of flow coefficient with data from Bodner et al. at M = 2.46. 

0.000

0.005

0.010

0.015

0.020

0.025

0.030

0.035

0.00 0.01 0.02 0.03 0.04 0.05 0.06 0.07

Q

Pplen/Pt,e

Bodner et al., M=2.46, Re'=1.69

Present,M=2.46, Re'=1.89

Present,M=2.46, Re'=2.46

0.000

0.020

0.040

0.060

0.080

0.100

0.120

0.140

0.00 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08

Q

Pplen/Pt,e

Bodner et al., M=2.46, Re'=1.69

Present,M=2.46, Re'=1.89

Present,M=2.46, Re'=2.46
NASA/TM—2012-217620 16 

a) C01 bleed configuration 

b) C02 bleed configuration 

Figure 14.—Sonic flow coefficient as a function of normalized bleed plenum pressure. 

 

M =1.4 blk

3.0

2.0

2.5

1.7

M =1.4 blk

3.0

2.0

2.5

1.7
NASA/TM—2012-217620 17 

TABLE 2.—PARAMETERS OF BLEED CONFIGURATIONS 

OF WILLIS et al. (REF. 2) 

Variable C1 C3 

 (degrees) 90.0 20.0 

D (mm) 6.35 6.35 

L/D 1.0 2.92 

D/1 ~1.0 ~1.0 

No. of holes 75 75 

No. of rows 6 6 

No. of columns 12.5 12.5 

x/D spacing 2.0 2.78 

z/D spacing 2.0 2.0 

Porosity (percent) 19.6 14.1 

The present data for Re = 1.89 were compared with the multi-hole data of Willis et al. (Ref. 2). 

The configurations compared are the C1 ( = 90) and C3 ( = 20) bleed plates of Willis et al. The 

parameters of these plates are summarized in Table 2. The data were accumulated at similar Mach 

numbers (Me = 1.27, 1.58, 1.98, 2.46) to the present test. Notable differences between the present hole 

parameters and those in the Willis plates are D/1 and L/D. The comparison of the flow coefficient 

curves are shown in Figure 15. With reference to Figure 15(a), the data trends are very similar with the 

exception that the single-hole data tend to show a more defined choke at a lower level than that of the 

multi-hole data. This is best illustrated by the Mach 2.0 cases, where the edge Mach numbers are 

almost the same and the curves are nearly coincident until the choke points are reached. The present 

Mach 2.0 (Me = 1.97) data would be expected to be slightly to the right of the multi-hole Mach 2.0 

(Me = 1.98) data, but the larger L/D ratio of the single-hole would tend to move it back to the left. The 

higher choke point of the multi-hole data implies that each successive row of holes sees a slightly larger 

local pressure. The present data at M = 2.5 is also similar but shifted slightly to the left which is likely 

due to the higher hole length-to-diameter ratio. For the C02 configuration (Fig. 13(b)), the difference 

between the choke level is still evident, but to a much smaller degree. The Mach 2.0 and 2.5 data agree 

well with the multi-hole data of Willis. The more significant difference between the distributions is in 

the subcritical flow region (low flow rates) where the pressurization of the bleed plenum is much more 

pronounced for the multi-hole configuration. 
NASA/TM—2012-217620 18 

a) C01 bleed configuration ( = 90) 

b) C02 bleed configuration ( = 20) 

Figure 15.—Present data compared to data of Willis et al. 

0.00

0.05

0.10

0.15

0.20

0.25

0.00 0.10 0.20 0.30 0.40 0.50 0.60

Q

Pplen/Pt,e

Willis et al., M=1.27, Re'=1.35

Willis et al., M=1.58, Re'=1.34

Willis et al., M=1.98, Re'=1.77

Willis et al., M=2.46, Re'=1.75

Present,M=1.33, Re'=1.89

Present,M=1.62, Re'=1.89

Present,M=1.97, Re'=1.89

Present,M=2.46, Re'=1.89

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70

Q

Pplen/Pt,e

Willis et al., M=1.27, Re'=1.35

Willis et al., M=1.58, Re'=1.34

Willis et al., M=1.98, Re'=1.77

Willis et al., M=2.46, Re'=1.75

Present,M=1.33, Re'=1.89

Present,M=1.62, Re'=1.89

Present,M=1.97, Re'=1.89

Present,M=2.46, Re'=1.89
NASA/TM—2012-217620 19 

Flow Coefficient Scaling 

The similarity of the curves for the different Mach numbers has led the authors to seek a scaling that 

would collapses the curves to a single distribution. Initial efforts by Davis (Ref. 4) and Slater (Ref. 1) 

focused on the 90 data of Willis et al. Both investigators took the approach of normalizing the bleed 

plenum pressure by the local surface static pressure, but Davis also included a coefficient to account for 

the slight overpressure of the bleed plenum at zero flow rates. For scaling the flow coefficient, the 

investigators took different approaches. Whereas Slater assumed for the 90 hole case that the total 

pressure in the hole was approximately equal to the local surface static pressure, Davis established a 

purely empirical scaling based on the extrapolated choked value of the bleed plate. Slater’s correlation 

takes the form of: 

   2

scaled 597997350 030693460 plen scaled, 593614200 Pplen scaled,  .Q  .  P  .  (8) 

where 

2 1

scaled,

M

2

1 1 







 



   





















































e

e,t

plen

e,t

w

e,t

plen

plen

P

P

P

P

P

P

P (9) 

and 

2 1

scaled

M

2

1 1 







 



   

















 

e e,t

w

Q

P

P

Q Q (10) 

and Davis’ scaled correlation takes the form of: 







































 

20744313

scaled,

scaled

914732540

1

73971571 741772710 .

plen

.

P

. .Q (11) 

where 

2 1

scaled,

M

2

1 10591 0591 







 



   























































e

e,t

plen

e,t

w

e,t

plen

plen

.

P

P

P

P .

P

P

P (12)

and 

  e .. e.

Q Q

e

2 M scaled

   95328695M956987758852416 

 (13) 
NASA/TM—2012-217620 20 

Willis’ data scaled by both methods and the above correlations derived from these data are shown in 

Figure 16. Both methods collapse the 90 data fairly well. To illustrate deviations of the present singlehole data, the present data are compared to these above correlations in Figure 17. This plot shows Slater’s 

correlation fits the present data better which isn’t unexpected since Davis’ correlation was based on an 

empirical fit of the extrapolated choke points of the multi-hole data. Slater’s correlation seems to fit the 

lower Mach number data better with increasing deviation as the Mach number is increased. However, 

even the lowest Mach number data deviation from the Slater correlation near the choke point, implying 

that an adjustment for the number of bleed hole rows, and potentially the hole spacing, may be required. 

Finally, the 20 present single-hole and multi-hole data of Willis are plotted scaled by the parameters 

suggested by Slater (Eqs. (9) and (10)) in Figure 18. For low flow rates (Qscaled  1.0), it is clear that the 

single and multi-hole data have a significantly different behavior. It is also evident that the single-hole 

distributions choke at slightly lower values than the multi-hole data, but more significantly the choked 

data do not collapse as well as the 90 data. This suggests that the assumption that the total pressure in the 

hole is equal to the surface static is not valid for shallow angled holes. This is expected since the shallow 

angled holes do not produce a strong barrier shock like the steeper angled holes do. Developing a 

correlation that includes the effect of hole angle will require additional data at intermediate hole angles 

which will be the subject of future work. 

Figure 16.—Correlations of Davis and Slater derived from 90 multi-hole data of Willis et al. 

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0.00 0.20 0.40 0.60 0.80 1.00 1.20 1.40

Qscaled

Pplen,scaled

Willis et al., M=1.27, Re'=1.35

Willis et al., M=1.58, Re'=1.34

Willis et al., M=1.98, Re'=1.77

Willis et al., M=2.46, Re'=1.75

Davis Correlation

Slater Correlation

1.00
NASA/TM—2012-217620 21 

Figure 17.—Comparison of present 90 hole data to correlations of Davis and Slater. 

Figure 18.—Single and multi-hole 20 data scaled per Slater correlation. 

0.00

0.20

0.40

0.60

0.80

1.00

1.20

0.00 0.20 0.40 0.60 0.80 1.00 1.20 1.40

Qscaled

Pplen,scaled

Present,M=1.33, Re'=1.89

Present,M=1.62, Re'=1.89

Present,M=1.97, Re'=1.89

Present,M=2.46, Re'=1.89

Present,M=2.96, Re'=1.89

Present,M=1.33, Re'=2.46

Present,M=1.62, Re'=2.46

Present,M=1.97, Re'=2.46

Present,M=2.46, Re'=2.46

Present,M=2.96, Re'=2.46

Davis Correlation

Slater Correlation

1.00

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

2.00

0.00 0.20 0.40 0.60 0.80 1.00 1.20 1.40

Qscaled

Pplen,scaled

Willis et al., M=1.27, Re'=1.35

Willis et al., M=1.58, Re'=1.34

Willis et al., M=1.98, Re'=1.77

Willis et al., M=2.46, Re'=1.75

Present,M=1.33, Re'=1.89

Present,M=1.62, Re'=1.89

Present,M=1.97, Re'=1.89

Present,M=2.46, Re'=1.89
NASA/TM—2012-217620 22 

Preliminary Computational Results 

The methods of computational fluid dynamics (CFD) were applied to simulate the bleed through the 

C01 (90) and the C02 (20) bleed hole configurations. The flow conditions simulated were those of 

FC003, which were for a nominal Mach number of 1.4 with Reynolds number at 2.46E+07/m. The WindUS code was used to solve the Reynolds-Average Navier-Stokes equations for the steady-state flow using 

upwind methods and the Menter SST turbulence model. Slater (Ref. 1) discusses the use of Wind-US for 

simulation of flows through bleed holes. 

The computational flow domain modeled portions of the tunnel test section, bleed hole, and plenum. 

Geometric symmetry about a streamwise vertical plane at the center of the hole allowed the flow domain 

to be reduced to include only half of the hole, tunnel, and plenum. The modeling also assumed a row of 

holes perpendicular to the direction of the core tunnel flow. Therefore, a vertical, streamwise symmetry 

plane was placed equidistant between adjacent holes. Thus, the flow domain only included a streamwise 

slice of the tunnel and plenum that included only one-half of one bleed hole. The sides of the flow domain 

were modeled with symmetry boundary conditions. The top of the flow domain was a horizontal plane 

located at 10.16 cm (4.0 in.) above the test section bottom wall and modeled with an inviscid wall 

boundary condition. The inflow boundary was placed upstream of the bleed hole at the reference station 

at which the boundary layer properties were measured in the experiment. The inflow boundary conditions 

were specified with a turbulent boundary layer profile obtained from a separate CFD simulation to match 

the boundary layer edge Mach number and thickness as measured in the experiment. An outflow 

boundary was placed 3.81 cm (1.5 in.) downstream of the bleed hole within the tunnel and specified with 

a supersonic outflow boundary condition. Another outflow boundary was specified at the bottom of the 

bleed plenum to extract the bleed flow. This boundary was specified with a subsonic outflow boundary 

condition with the static pressure specified. This outflow boundary condition was used to set the plenum 

static pressure, which controlled the rate of bleed flow. The surfaces of the tunnel test section floor, bleed 

hole, and plenum completed the flow domain and adiabatic, viscous wall boundary conditions were 

applied. The bleed plenum was sized large such that the bleed flow dissipated to near quiescent flow 

within the bleed plenum and the static pressure specified at the plenum outflow approximately equaled 

the static pressure in the plenum. The plenum was shaped to form a converging nozzle at the bottom of 

the plenum where the outflow boundary was placed. 

The computational grid for the flow domain was automatically generated and with specification of the 

bleed hole diameter, bleed hole inclination angle, and hole spacing within the row. Since the experiment 

only involved one bleed hole, a series of CFD simulations were performed to determine that the hole 

spacing within the row had to be 4.0 hole diameters to avoid interaction between bleed holes. Structured 

grids were generated using 13 blocks in which grid lines were contiguous across the block boundaries. 

The grid resolution normalized by the hole diameter was 0.03378 within the hole. The grids consisted of 

approximately 2.5106

 grid points. 

The CFD simulations were performed over a series of plenum pressures. Figure 19 shows the 

variation of the flow coefficients with comparison to the experimental data and the comparisons are very 

good. Some differences are apparent for the 20 hole at plenum pressure ratios about 0.34. The good 

agreement boosts confidence in both the experiment and CFD data. 

Figure 20 shows Mach number contours of the bleed flow on the hole symmetry plane for the supercritical operating point on the flow coefficient curve, which occurs for the lowest value of bleed plenum 

pressure. The 90 hole requires significant turning of the bleed flow, which manifests itself as a detached 

barrier shock at the top of the hole, a recirculation region at the front of the hole, and a supersonic jet 

toward the back of the hole. The jet extends into the plenum, but is dissipated. The flow for the 20 hole 

does not require a barrier shock; however, a supersonic jet forms that extends into the plenum before 

dissipating. 
NASA/TM—2012-217620 23 

Figure 19.—Comparison of the flow coefficients between 

CFD and experiment for the 90 and 20 bleed holes at 

Mblk = 1.4 and Re = 2.46. 

Figure 20.—Mach contours for the super-critical flow through the 90 (left) and 20 (right) bleed holes. 

Summary and Conclusions 

Flow coefficient data for 90 and 20 single bleed holes have been obtained and compared to multihole data under similar conditions. The results show that the 90 hole data differs most under supercritical 

operation and the 20 data differs most under subcritical operation. The preliminary CFD shows good 

agreement with the experimental results. The single-hole studies will be continued under Phase II of the 

FIBE project looking initially at intermediate hole inclination angles and length-to-diameter ratios. This 

will then be followed by boundary layer to hole diameter scaling and also entrance conditions (radius) to 

the hole. These data will be used to develop improved bleed models that can account for these effects. 

Finally, Phase III will look at bleed regions with and without shock waves to validate the performance of 

the resulting bleed models. 

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.0 0.1 0.2 0.3 0.4

Sonic Coefficient,

Q

pplen / pt,e

Data90‐deg

CFD  90‐deg

Data20‐deg

CFD  20‐deg
NASA/TM—2012-217620 24 

References 

1. Slater, J.W., “Improvements in Modeling 90° Bleed Holes for Supersonic Inlets,” AIAA Paper 

2009-0710, 47th AIAA Aerospace Sciences Meeting and Exhibit, Orlando, FL, January 5-8, 2009 

(also NASA/TM—2009-215597). 

2. Willis, B.P., Davis, D.O., and Hingst, W.R., “Flow Coefficient Behavior for Boundary-Layer Bleed 

Holes and Slots,” AIAA Paper 95-0031, 33rd AIAA Aerospace Sciences Meeting and Exhibit, 

Reno, NV, January 9-12, 1995 (also NASA TM-106846). 

3. Davis, D.O., Grimes, M., and Schoenenberger, M., “Effect of Flow Misalignment and Multi-Hole 

Interaction on Boundary-Layer Bleed Hole Flow Coefficient Behavior,” Proceedings of the ASME 

Fluids Engineering Division, FED-Vol. 242, American Soc. of Mech. Eng., New York, 1996, 

pp. 323-328 (also NASA TM-107480). 

4. Davis, D.O., “Fundamental Bleed Experiments at NASA Glenn Research Center 1995-1999,” 

1st Shock Wave/Boundary Layer Interaction Workshop, Cleveland, OH, April 15-16, 2008. 

5. Bodner, B.P., Davis, D.O., Hingst, W.R., and Greber, I., “Experimental Investigation of the Effect of 

a Single Bleed Hole on a Supersonic Turbulent Boundary Layer,” AIAA Paper 96-2797, 32nd AIAA 

Joint Propulsion Conference and Exhibit, Lake Buena Vista, FL, July 1-3, 1996. 

6. Bodner, J.P., “Experimental Investigation of the Effect of a Single Bleed Hole on a Turbulent 

Supersonic Boundary Layer,” M.S. Thesis, Dept. of Mechanical and Aerospace Engineering, Case 

Western Reserve University, Cleveland, OH, May 1996. 

7. Schoenenberger, M., Greber, I., and Davis, D. O., “Flow Measurements Downstream of a Single 

Bleed Hole in a Subsonic, Turbulent Boundary Layer Using a New 5-Hole Pressure Probe,” AIAA 

Paper 99-0293, 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, January 11-14, 

1999. 

8. Schoenenberger, M., “Experimental Investigation of a Bleed Hole in a Subsonic Turbulent Boundary 

Layer,” M.S. Thesis, Dept. of Mechanical and Aerospace Engineering, Case Western Reserve 

University, Cleveland, OH, May 1998. 

9. Davis, D.O., and Reichert, B.A., “Ethylene Trace-Gas Techniques for High-Speed Flows,” AIAA 

Paper 94-0733, 32nd AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, January 10-13, 

1994. 

10. Davis, D.O., Willis, B.P., Hunter, G.W., Liu, C-C., and Wu, Q., “Mass-flow Measurement Using a 

Hydrocarbon Trace-Gas Technique,” Proceedings of the ASME Fluids Engineering Division, FEDVol. 239, American Soc. of Mech. Eng., New York, 1996, pp. 465-469. 

11. van Driest, E.R., “Turbulent Boundary Layer in Compressible Fluids,” J. Aeronautical Sciences, 

Vol. 18, 1951, pp. 145-160. 
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1. REPORT DATE (DD-MM-YYYY)

01-08-2012 

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Technical Memorandum 

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4. TITLE AND SUBTITLE

Research on Supersonic Inlet Bleed 

5a. CONTRACT NUMBER

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5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Davis, David, O.; Vyas, Manan, A.; Slater, John, W. 

5d. PROJECT NUMBER

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5f. WORK UNIT NUMBER

WBS 984754.02.07.03.13.02 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

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E-18208 

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NASA 

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Available electronically at http://www.sti.nasa.gov 

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14. ABSTRACT

Phase I data results of the Fundamental Inlet Bleed Experiments project at NASA Glenn Research Center (GRC) are presented which

include flow coefficient results for two single-hole boundary-layer bleed configurations. The bleed configurations tested are round holes at 

inclination angles of 90° and 20° both having length-to-diameter ratios of 2.0. Results were obtained at freestream Mach numbers of 1.33, 

1.62, 1.98, 2.46, and 2.92 and unit Reynolds numbers of 0.984, 1.89, and 2.46×107/m. Approach boundary-layer data are presented for each 

flow condition and the flow coefficient results are compared to existing multi-hole data obtained under similar conditions. For the 90° hole, 

the single and multi-hole distributions agree fairly well with the exception that under supercritical operation, the multi-hole data chokes at 

higher flow coefficient levels. This behavior is also observed for the 20° hole but to a lesser extent. The 20° hole also shows a markedly 

different characteristic at subcritical operation. Also presented are preliminary results of a Computational Fluid Dynamics (CFD) analysis of 

both configurations at the Mach 1.33 and a unit Reynolds number of 2.46×107/m. Comparison of the results shows the agreement to be very 

good. 

15. SUBJECT TERMS

Flow coefficients; Inlet flow; Mach number; Static pressure; Supersonic flow; Supersonic inlets; Bleeding; Boundary layer control; 

Boundary layer flow; Boundary layer separation; Compressible flow 

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