René Fernandez, Charles J. Trefny, and Scott R. Thomas

Glenn Research Center, Cleveland, Ohio

Mel J. Bulman

Gencorp Aerojet Corporation, Sacramento, California

NASA/TM—2001-107181

November 2001

Parametric Data From a Wind Tunnel

Test on a Rocket-Based CombinedCycle Engine Inlet
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NASA/TM—2001-107181

November 2001

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

1995 Airbreathing Propulsion Subcommittee Meeting

sponsored by the Joint Army-Navy-NASA-Air Force Interagency Propulsion Committee

Tampa, Florida, December 5–9, 1995

René Fernandez, Charles J. Trefny, and Scott R. Thomas

Glenn Research Center, Cleveland, Ohio

Mel J. Bulman

Gencorp Aerojet Corporation, Sacramento, California

Parametric Data From a Wind Tunnel

Test on a Rocket-Based CombinedCycle Engine Inlet
Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

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Available electronically at http://gltrs.grc.nasa.gov/GLTRS
NASA/TM—2001-107181 1

PARAMETRIC DATA FROM A WIND TUNNEL TEST ON A

ROCKET- BASED COMBINED-CYCLE ENGINE INLET

René Fernandez, Charles J. Trefny, and Scott R. Thomas

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Mel J. Bulman

Gencorp Aerojet Corporation

Sacramento, California 95813–6000

SUMMARY

A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the

NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the Hypersonic Tunnel Facility (HTF) at NASA Glenn’s Plum Brook Station at Mach 5 and 6. This

engine will incorporate the configuration of this inlet model which achieved the best performance during the present

experiment. The inlet test was conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8° forebody compression plate, boundary layer diverter, and two compressive struts located within 2 parallel side-walls. These struts extend through the inlet, dividing the flowpath into three channels. Test parameters

investigated included strut geometry, boundary layer ingestion, and Reynolds number (Re). Inlet axial pressure

distributions and cross-sectional Pitot-pressure surveys at the base of the struts were measured at varying backpressures. Inlet performance and starting data are presented. The inlet chosen for the RBCC engine self-started at all

Mach numbers from 4 to 6. Pitot-pressure contours showed large flow nonuniformity on the body-side of the inlet.

The inlet provided adequate pressure recovery and flow quality for the RBCC cycle even with the flow separation.

NOMENCLATURE

Ad boundary layer diverter freestream flow area

Ai inlet capture area

Ao freestream captured stream tube area

M4 mass averaged Mach number at the end of the struts

Mo freestream Mach

ps model static pressure

pso freestream static pressure

pTO freestream total pressure

TTO freestream total temperature

ηpt mass averaged total pressure recovery

ηKE mass averaged kinetic energy efficiency

γ gamma, ratio of specific heats, 1.4

BACKGROUND AND INTRODUCTION

Rocket-based combined-cycle (RBCC) engines combine the high thrust-to-weight ratio of rockets with the

high specific impulse of ramjets in a single, integrated propulsion system that is capable of generating thrust from

sea-level-static to high Mach number conditions. The “strutjet” (ref. 1) is one example of this engine concept which

is being developed cooperatively by a government-industry team.

The strutjet is an ejector-ramjet engine in which small, fuel-rich rocket chambers are embedded into the

trailing edges of the inlet compression struts. The engine operates as an ejector-ramjet from take-off to about
NASA/TM—2001-107181 2

Mach 3. At low Mach number, entrained air is completely consumed by the fuel-rich rocket exhaust. As freestream

Mach number and airflow increase, additional fuel is introduced to maintain the stoichiometric combustion of all

available oxygen. At approximately Mach 3 the strut rockets are turned off. From Mach 3 to 6, the engine operates

as a thermally-choked ramjet, and then transitions to supersonic-combustion (scramjet) mode at about Mach 6. For

space-launch applications, the rockets are re-ignited at a Mach number beyond which air-breathing propulsion

becomes impractical.

A strutjet engine test article, designed for up to Mach 8 cruise on hydrocarbon fuel, is currently being

developed and will be tested initially at the Hypersonic Tunnel Facility (HTF) free-jet at the NASA Plum Brook

Station (ref. 2). To properly match Mach 6 and 7 freestream flight conditions and simulate vehicle forebody compression, Mach 5 and 6 nozzles will be used at Mach 6 and 7 enthalpy levels. Figure 1 shows a schematic of the

cross section of this free-jet engine design. An 8° compression ramp upstream of the inlet top surface leading edge is

not shown. The 8° compression ramp can be positioned for either ingestion or diversion of the boundary layer (BL).

Figure 2 is a photograph of the base regions of the two existing struts which have been extensively tested in a direct

connect experiment and will be incorporated into the RBCC engine. These struts each contain three separate rocket

chambers, and two hydrocarbon fuel (JP-10) injector wedges. The rocket elements operate on monomethyl hydrazine (MMH) and inhibited red fuming nitric acid (IRFNA). Liquid JP-10 fuel can also be injected through offices in

the struts further upstream. The upstream injectors are piloted using small MMH-IRFNA pilots to promote fuel

vaporization and ignition. The free-jet airflow is compressed by the 8° ramp, then by the strut leading edges which

reach a maximum thickness at the cowl leading edge. There is no internal contraction downstream of the cowl leading edge; although the cross sectional geometry varies, the net cross sectional area is constant. The convergence

between the cowl and body-side panels is compensated for by a reduction in strut thickness. This lack of internal

contraction enables the inlet to self-start at lower Mach numbers. The non-compressive sidewalls simulate symmetry

planes. The combustor and nozzle area variation downstream of the strut trailing edges can be varied manually using

the screw-jacks.

During development of the free-jet engine test article, uncertainty in inlet performance and operability

emerged as a significant risk to successful strutjet demonstration. A subscale, parametric inlet test was therefore

conducted to determine the optimum strut geometry, explore the effect of boundary layer diversion, and quantify the

performance and operability of the inlet at the HTF test Mach numbers of 5 and 6. Establishing the cross-sectional

flow distribution at the base of the struts was also desired in order to optimize the hydrocarbon fuel injector locations. This paper documents the test program conducted in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel

(1×1 SWT).

APPARATUS AND PROCEDURE

Sub-Scale Inlet Model

A photograph of the model with the cowl removed, mounted onto the tunnel sidewall is presented in

figure 3. The model is uncooled and constructed of carbon-steel. The flowpath is a 40-percent scale version of the

inlet for the RBCC freejet engine currently being developed. Figure 4 is a schematic drawing of the model showing

the major components. Coordinates of the body and cowl-side surfaces are listed in table I. The model was positioned such that the leading edge of the top surface was either flush with the precompression plate to ingest the

boundary layer, or dropped 0.4 in. vertically below the precompression plate for BL diversion. This was accomplished by using shims between the tunnel sidewall and the baseplate of the 8° precompression plate (forebody

simulator). With the diverter height set to 0.4 in., the projected capture area is 13.76 in. (ref. 2). The outer sidewalls

are flat on the inside surface to simulate symmetry planes. Two identical compression struts divide the flowpath into

three channels. The three sets of strut configurations tested are depicted in figure 5. Configurations “A” and “B”

reach a maximum thickness at the cowl lip (inlet entrance) plane, resulting in a constant geometric cross-sectional

area in all three channels from the cowl lip to the base of the struts. Configuration “A” is tapered from cowl-side to

body-side, configuration “B” is untapered. The constant thickness struts (configuration “C”) result in an internal

geometric contraction ratio of 1.55. All three strut geometries are 0.90 in. thick by 1.60 in. high, aft of the 8° turn on

the body-side (shoulder point). The central channel is 0.94 in. wide while that of the outer channels is 0.49 in. The

height of the channels at the inlet shoulder (normal to the body-side surface) is 2.48 in. and 1.60 in. at the strut trailing edge station. Aft of the struts, the flowpath is 2.64 in. wide and 1.60 in. high.
NASA/TM—2001-107181 3

The cowl, body, and compression ramp surfaces of the model are instrumented with a row of static pressure

taps along the centerline. The model sidewalls each have a row of static pressure taps at nominally 1 in. spacing on a

line halfway between the cowl and body. Table II gives the coordinates of the static pressure taps. A 12-tube pitot

pressure rake was used to survey the flowfield exiting all three channels at the base of the struts. The rake traversed

from body-side to cowl-side with measurements taken at 20 equally spaced intervals. Of the 12 tubes on the rakes,

6 were used in the center channel and 3 were used in each of the side channels. The rake was translated away from

the body-side by a set of thin bars which protruded through the surface directly behind the strut trailing edges, in

order to not introduce any additional blockage.

To simulate the back-pressure caused by combustion, a wedge-shaped, hydraulically-actuated mass flow

plug was used. At the limit of its forward travel, the tip of the wedge extends to a point corresponding to the trailing

edge of the first movable panel in the free-jet engine.

1- by 1-Foot Supersonic Wind Tunnel

The 1×1 SWT (ref. 3) is a one-pass continuous flow facility that can test models from Mach 1.3 to 6.0. The

upstream high pressure air is provided by the lab-wide central air services, while the downstream flow is routed to

the lab’s altitude exhaust system. Figure 6 shows the tunnel circuit as well as facility details. The different Mach

numbers are achieved by changing two-dimensional nozzle “blocks”; for the present test program the Mach 4.0, 5.0,

5.5, and 6.0 blocks were used. The tunnel incorporates an electrical resistance heater to bring the air up to elevated

temperature. The air is heated to a high enough temperature to prevent condensation in or liquification of the air in

the flow. The test conditions properly match the required Mach numbers and Reynolds numbers. The facility stagnation conditions that were run for the various Mach numbers are listed in table III.

Inlet Test Parameters

For each specified set of compression struts, the inlet backpressure (simulated combustor pressure) was

varied by translating the mass flow plug. Inlet mass capture was computed independently from four different techniques: employing the mass flow plug data; integrating local mass flow rates determined from static and Pitot pressure measurements at the traversing rake plane; using model dimensions and static pressure measurements at the

capture station; and from CFD computations. Boundary layer ingestion or diversion was varied by positioning the

precompression plate using steel shims. The compression struts were manually interchanged between test series.

Reynolds number variations per Mach number were accomplished by varying the tunnel stagnation pressure and/or

the tunnel total temperature.

Inlet Operability Tests

The goal of this program was to select a fixed geometry configuration that consistently and easily self

started and provided an acceptable overall inlet performance. Establishing the inlet starting characteristics as a function of freestream Mach for the three different strut configurations was a primary objective. Determining the inlet

performance parameters and exit flow uniformity to assess overall flowfield quality at the exit plane was also an

important objective. Table IV summarizes the resultant Test Matrix used to satisfy the goal.

To quantify the quality of the flowfield that this inlet provided to the combustor, Pitot pressure, total pressure, Mach number, and mass flow contours were determined for the flowfield at the base of the struts. The Pitot

pressure surveys at the inlet exit that along with static pressure measurements at the plane at the base of the struts,

were used to compute Mach number contours. The static pressure taps were located 0.38 in. ahead of the traversing

rake station, on the centerline cowl and body side of each of the three ducts, and on the outboard sidewalls of each

of the outer ducts.

Reynolds numbers effects on inlet performance were recorded; these are aerodynamic effects that primarily

manifest themselves through the boundary layer. The Reynolds number was changed by varying either or both the

tunnel plenum pressure or the supply airflow line temperature.
NASA/TM—2001-107181 4

Inlet Backpressure Mapping

The mass flow plug was used to determine the maximum pressure that could be maintained in the combustor region before inlet unstart. Inlet operation was characterized from the supercritical condition (plug fully

retracted) incrementally through plug translation up to minimum stable operation then on to inlet unstart. As the

mass flow plug travel was incremented in the upstream direction, the model exit area was gradually reduced. As the

minimum stable point was approached, very fine increments were used to establish a high degree of accuracy when

determining the maximum combustor pressure for all configurations. The maximum combustor pressure was established at each Mach number as a function of strut configuration for both boundary layer ingestion and diversion.

RESULTS AND DISCUSSION

Mass Capture and Operability Data

The results of the parametric variations are summarized in table V. The remainder of this paper will focus

on configuration B1, the untapered struts with boundary layer diversion, at Mach numbers of Mach 5 and 6. This

configuration showed good starting characteristics, was stable, had acceptable flow uniformity, and the best overall

inlet performance. This configuration was selected for use in the RBCC engine. Data for all other configurations are

available and will be documented.

Inlet mass capture ratios for a started inlet versus freestream Mach are shown in figure 7. The experimental

results were calculated from integration of the inlet exit survey data. Theoretical predictions were based on uniform

flow and Ames Tables calculations. The NPARC Computational Fluid Dynamics (CFD) code (ref. 4) calculations

were made with the following boundary conditions: ideal gas, adiabatic wall, and turbulent flow (k-ε model).

Results obtained using the mass flow plug are not shown because large flow non uniformity flowing downstream

from the isolator and combustor into the mass flowmeter rendered the mass flow results unreliable. Compared in

figure 7 are the theoretical total mass capture (the captured inlet flow (Ao) plus the captured diverter flow (Ad)), the

theoretical captured inlet flow (Ao), the experimentally derived captured inlet flow, and the CFD calculated captured

inlet flow. The theoretical total captured streamtube was determined from model geometry and dimensions, and

from the freestream Mach number dependent pre-compression plate shock angle. The experimentally derived captured inlet flow was determined by computing the required captured inlet area that would provide the massflow

computed by the traversing rake at the end of the struts. By subtracting the captured inlet area ratio from the theoretical total captured streamtube ratio, the captured diverter area ratio was determined to be 0.122 at Mach 4, 0.136

at Mach 5, and 0.153 at Mach 6. As is shown, the captured inlet area ratio did not vary significantly at the higher

Mach numbers.

Backpressure Data

Static and total pressure inlet data with increasing combustor pressure are shown in figures 8 to 11 for

Mach 5 and 6. Figures 8 and 10 show profiles of surface static pressure normalized by freestream static pressure for

the model body, cowl, and sidewalls. In order to keep the figures uncluttered, only six cases (I, II, III, IV, V, and VI)

are shown out of the large experimental data set. Figures 9 and 11 present corresponding Mach number contours at

the base of the struts. The Mach contours are shown are for four cases: a supercritical inlet case, two increasing back

pressure cases, and a minimum stable case.

Figure 8 outlines the inlet behavior with increasing back pressures from super critical inlet operation past

minimum stable and on to inlet unstart. In the center duct the shock system ahead of the base of the struts is not initially affected by the increasing back pressure; the static pressures in the combustor region rise, but upstream of the

combustor the profile has not changed. Subsequent profiles show pressure rises well forward into the isolator and

ahead of the inlet shoulder. The minimum stable condition shows the pressure rise near the cowl lip station (maximum thickness station of the struts). These plots show that as the back pressure was increased, the shock system

would stop, or stall, on certain sections of the model, such as the end of the struts or the inlet shoulder, before progressing forward again with increasing back pressures. The final, unstarted, profile shows that although the first

static pressure tap on the inlet body is at an elevated pressure, the rest of the profile drops. The left sidewall
NASA/TM—2001-107181 5

appeared to experience a classic hard unstart however the right sidewall taps exhibit a profile similar to the middle

duct. This indicates that once the left duct has unstarted, sufficient flow was spilled to prevent the other two ducts

from unstarting. The pressure dropped in the inlet isolator due to the drop in mass flow through the model resulting

from the left duct unstart.

Figure 9 highlights static pressure profile cases I, III, IV, and V, showing the corresponding combustor

entrance Mach number flowfield. The supercritical case shows a supersonic flowfield and symmetry between the

left and right ducts. Case III confirms what the static pressure profiles indicated in figure 8, that the effect of back

pressure is occuring only in the combustor and that the flow ahead of the rake is not being affected. Case IV shows

that when static pressure rise occurs ahead of the inlet shoulder the entire flowfield changes and regions of subsonic

flow appear on the body side of the duct. The minimum stable case, V, shows a highly distorted flowfield with

asymmetric outer ducts. Although most of the duct is subsonic flow, a region of supersonic flow remains near the

cowl side of the duct.

Figure 10 outlines the inlet behavior with increasing back pressures at Mach 6. These results are different

than those for Mach 5. The shock train moves forward from the combustor through the isolator and into the front of

the inlet isolator without getting stalling on sections of the inlet geometry (such as the end of the struts or the inlet

shoulder) as occurred at the Mach 5 test condition. Note that the pressure rise is observed forward of the combustor,

through the isolator, and near the entrance of the inlet for a stable started case. Another difference with the Mach 5

data is the higher pressure ratio achieved in the combustor (60 percent greater at minimum stable). The one similarity is in the unstarted profile. The unstarted profiles are similar for both Mach numbers because the unstarted left

duct causes a pressure rise on the leading edge of the inlet centerline; the static pressure profiles, however, equalize

near the shared downstream exit of the isolator.

Figure 11 highlights static pressure profile cases I, III, IV, and V. These Mach number contours concur

with the static profiles that the pressure rise through the inlet occurs more gradually at Mach 6 than at Mach 5. The

supercritical case I exhibits a supersonic flowfield. Case III shows flow field gradients from subsonic on the body

side to supersonic on the cowl side for all three ducts. For Case IV with back pressure, most of the exit profile is

subsonic but there is a small sonic/supersonic region near the cowl side for all three ducts. The minimum stable

Case V shows a more symmetric flowfield across all three ducts with predominantly subsonic flow and a small sonic

core near the cowl side of the center duct.

The overall performance of the B1 configuration at freestream Mach 5 and Mach 6 is summarized in

figure 12 which presents kinetic energy efficiency, ηKE, versus the Mach number ratio, M4/Mo at the end of the

isolator. The kinetic energy efficiency is given by:

η

η

γ

γ

γ

KE

pt

Mo

=

−

( ) − +

( ) −

1 1

1

2

1

1

2

where M4 is the mass averaged Mach number at the end of the isolator as determined by the traversing rake. Note

that as the inlet is back pressured ηKE decreases until the inlet unstarts before M4/Mo reaches the sonic value for the

Mach 5 case, and M4/Mo goes under the sonic value for the Mach 6 case.

CONCLUSION

A comprehensive parametric study was conducted to establish the operability and performance of a

40-percent scale inlet for a rocket based combined cycle engine. This study served to demonstrate that the best performance and operability were achieved using a configuration that incorporated nontapered struts, no net internal

geometric contraction, and had no boundary layer ingestion (BL diversion). This configuration was therefore

selected for use in the RBCC engine. The details of the performance data for this configuration were presented in

this paper. This inlet data and flowfield details will help to interpret engine test results and potentially obtain optimized engine performance by providing guidance with fuel injector placement. This study generated an additional

data base for a set of modular inlet geometries at Mach 4 to 6. This program served to verify acceptable operability
NASA/TM—2001-107181 6

and performance of a candidate inlet geometry for use with the RBCC engine. This, therefore, eliminated a significant element of risk associated with the upcoming engine test program by providing a very cost effective means of

obtaining critical inlet information prior to building the RBCC engine model. This will assure that the inlet will start

and achieve good performance during the HTF test, and will provide flowpath performance data to help interpret

and streamline the engine test matrix.

REFERENCES

1. Bulman, M.J.; and Siebenhaar, A.: The Strutjet Engine: Exploding The Myths Surrounding High Speed

Airbreathing Propulsion. Aerojet Propulsion Division, Sacramento, California, AIAA–95–2475, July 1995.

2. Thomas, S.R.; Woike, M.R.; and Pack, W.D.: Mach-6 Integrated Systems Tests of the NASA Lewis

Research Center Hypersonic Tunnel Facility. JANNAF, December 1995 (NASA TM–107083).

3. Skebe, S.A.: Experimental Investigation of Two-Dimensional Shock Boundary Layer Interactions. Ph.D.

Thesis, Case Western Reserve University, Cleveland, Ohio, 1983.

4. Cooper, K.: NPARC 2.0-Features and Capabilities. AIAA–95–2609, July 1995.

TABLE II.—AXIAL COORDINATES OF RBBCC

INSTRUMENTATION

Compression plate,

in.

Body side,

in.

Cowl side,

in.

Sidewalls,

in.

0.93

4.00

7.07

10.14

12.12

13.36

14.59

15.83

17.07

18.31

19.55

20.78

22.02

23.15

24.40

26.65

26.90

27.59

31.21

33.46

16.96

18.21

19.46

20.71

21.96

23.21

24.46

25.71

26.96

27.68

31.30

16.00

17.25

18.49

19.74

20.98

22.22

23.44

24.69

25.97

27.20

27.89

31.55

TABLE III.—FACILITY

STAGNATION CONDITIONS

M0 PT0,

psia

TT0,

R

4.0

4.0

5.0

5.5

6.0

6.0

40

40

80

150

160

160

530

615

686

670

610

765

TABLE I.—RBCC INLET COORDINATES

Description X,

in.

YBL Divert,

in.

YBL Ingest,

in.

Compression plate

Inlet leading edge

Strut leading edge

Cowl lip

Inlet shoulder

Strut trailing edge

Inlet trailing edge

0

11.2

12.04

16.47

22.89

27.96

34.26

0

2.37

2.49

5.22

3.62

3.62

3.62

0

1.97

2.09

4.82

3.22

3.22

3.22
NASA/TM—2001-107181 7

87.22

Screw jack

Rocket extension

Isolator

Strut rocket

Untapered struts

Figure 1.—Strutjet engine test article.

29.22 16.00 42.00

TABLE IV.—AS TESTED TEST MATRIX

Configuration Strut/windscreen Pre-compression plate M0 Comments

A1

A2

B1

B2

C2

D1

Tapered

Tapered

Untapered

Untapered

Constant thickness

No strut

B.L. Diverted

B.L. Ingested

B.L. Diverted

B.L. Ingested

B.L. Ingested

B.L. Diverted

4,5,5.5,6

5,6

4,5,6

5,6

5,6

5

Acceptable performance

Negligible effect of B.L.

Best performance

Negligible effect of B.L.

Self start intermittent

Inlet did not start

TABLE V.—RBCC INLET PERFORMANCE SUMMARIZED

Configuration Mach Starting Superscript

ηpt, ηKE

Max

ps/pso

Stable

A1

A1

A1

A1

A2

A2

B1

B1

B1

B2

B2

C1

C2

C2

D1

4.0

5.0

5.5

6.0

5.0

6.0

4.0

5.0

6.0

5.0

6.0

6.0

5.0

6.0

6.0

Yes

Yes

Yes

Yes

Yes

Yes

Yes

Yes

Yes

Yes

Yes

No

Difficult

Yes

No

0.38,0.900

0.31,0.921

0.28,0.927

0.26,0.935

0.32,0.923

0.24,0.930

0.53,0.938

0.41,0.942

0.30,0.943

0.41,0.942

0.33,0.948

N/A

0.40,0.940

0.31,0.945

N/A

19

33

68

73

53

51

22

35

100

46

78

N/A

34

55

N/A

Yes

Yes

Yes

Yes

Yes

Yes

Yes

Yes

Yes

Yes

Yes

N/A

No

No

N/A
NASA/TM—2001-107181 8

Figure 2.—Rocket chambers and injector wedges on the base region 

 of the two struts. 

Figure 3.—Forty-percent RBCC inlet model mounted on tunnel sidewall.
NASA/TM—2001-107181 9

0.0 in.

0.4 in.

Pre-compression

plate 

11.2 in.

16.47 in. 1.6 in.

27.96 in.

22.89 in.

34.26 in.

Interchangeable

struts

Mass

flow

plug 

2.64 in.

Sidewall removed

Cowl removed

Figure 4.—Sub-scale inlet model.

Configuration A – Tapered strut 

Configuration B – Untapered strut 

Configuration C – Constant thickness strut 

Figure 5.—Three RBCC strut configurations.
NASA/TM—2001-107181 10

Heater

Valve 'H'

Expansion joint

Hydraulic lift

Surge tank

Air inlet – 40 psi and 125 psi

Valve 'J'

Model supply pressure – 450 psi

Valve 'F'

Nozzle blocks

in storage

48 in. altitude exhaust

Nozzle

section

Test section

Bleed housing chamber

Expansion joint

Cooling water spray

Diffuser section

Bleed system

Valve 'E'

Figure 6.—1- by 1-Foot Supersonic Wind Tunnel, Mach number 1.3 through 6.0, maximum total

 temperature 650 °R.

0.50

0.60

0.70

0.80

0.90

1.00

3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0

Theoretical Ao + Ad

Theoretical Ao

Experimental Ao (rake)

NPARC CFD Ao

Mo

Ao/Ai

Figure 7.—RBCC inlet captured massflow ratio, 

 configuration B1.
NASA/TM—2001-107181 11

10

20

30

40

50

60

70

80

90

100

110

0 5 10 15 20 25 30 35

Body side, centerline, in.

ps/pso

I

II

III

IV

V

VI

10

20

30

40

50

60

70

80

90

100

110

0 5 10 15 20 25 30 35

Cowl side, centerline, in.

ps/pso

10

20

30

40

50

60

70

80

90

100

110

0 5 10 15 20 25 30 35

Left sidewall, centerline, in.

ps/pso

10

20

30

40

50

60

70

80

90

100

110

0 5 10 15 20 25 30 35

Right sidewall, centerline, in.

ps/pso

Figure 8.—Model surface static pressure ratios for Mach 5, configuration B1. 
NASA/TM

—2001-107181 12 I – ps/pso = 13.44 (supercritical) IV – ps/pso = 51.25 (isolator backpressured)

III – ps/pso = 26.85 (combustor backpressured) V – ps/pso = 56.59 (minimum stable)

Contour

label

Contour

value

ABCDEFGHIJKLM

0.000

.250.500.750

1.0001.2501.5001.7502.0002.2502.5002.7503.000

Figure 9.—Traversing rake plane Mach number contours. Mach 5, configuration B1.
NASA/TM—2001-107181 13

10

20

30

40

50

60

70

80

90

100

110

0 5 10 15 20 25 30 35

Body side, centerline, in.

ps/pso

I

II

III

IV

V

VI

10

20

30

40

50

60

70

80

90

100

110

0 5 10 15 20 25 30 35

Cowl side, centerline, in.

ps/pso

10

20

30

40

50

60

70

80

90

100

110

0 5 10 15 20 25 30 35

Left sidewall, centerline, in.

ps/pso

10

20

30

40

50

60

70

80

90

100

110

0 5 10 15 20 25 30 35

Right sidewall, centerline, in.

ps/pso

Figure 10.—Model surface static pressure ratios for Mach 6, configuration B1. 
NASA/TM

—2001-107181 14 I – ps/pso = 13.64 (supercritical) IV – ps/pso = 84.79 (isolator backpressured)

III – ps/pso = 72.14 (combustor backpressured) V – ps/pso = 100.45 (minimum stable)

Contour

label

Contour

value

ABCDEFGHIJKLM

0.000

.250.500.750

1.0001.2501.5001.7502.0002.2502.5002.7503.000

Figure 11.—Traversing rake plane Mach number contours. Mach 6, configuration B1.
NASA/TM—2001-107181 15

0.80

0.82

0.84

0.86

0.88

0.90

0.92

0.94

0.96

0.98

1.00

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

Mach 5

Mach 6

M4/Mo

KE

Figure 12.—RBCC inlet performance, configuration B1.
This publication is available from the NASA Center for AeroSpace Information, 1–301–621–0390.

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November 2001

NASA TM—2001-107181

E–10144

WU–708–48–13–00

21

Supersonic inlet; Hypersonic inlet; Ramjet; Scramjet; Rocket

Unclassified -Unlimited

Subject Categories: 07, 02, and 34 Distribution: Nonstandard

Prepared for the 1995 Airbreathing Propulsion Subcommittee Meeting sponsored by the Joint Army-Navy-NASA-Air

Force Interagency Propulsion Committee, Tampa, Florida, December 5–9, 1995. René Fernandez, Charles J. Trefny,

and Scott R. Thomas, NASA Glenn Research Center; Mel J. Bulman, Gencorp Aerojet Corporation, Sacramento,

California 95813–6000. Responsible person, René Fernandez, organization code 2780, 1–216–433–6281.

René Fernandez, Charles J. Trefny, Scott R. Thomas, and Mel J. Bulman

Parametric Data From a Wind Tunnel Test on a Rocket-Based Combined-Cycle

Engine Inlet

A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the NASA Glenn

Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the

Hypersonic Tunnel Facility (HTF) at NASA Glenn’s Plum Brook Station at Mach 5 and 6. This engine will incorporate the

configuration of this inlet model which achieved the best performance during the present experiment. The inlet test was

conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8° forebody compression

plate, boundary layer diverter, and two compressive struts located within 2 parallel sidewalls. These struts extend through

the inlet, dividing the flowpath into three channels. Test parameters investigated included strut geometry, boundary layer

ingestion, and Reynolds number (Re). Inlet axial pressure distributions and cross-sectional Pitot-pressure surveys at the

base of the struts were measured at varying back-pressures. Inlet performance and starting data are presented. The inlet

chosen for the RBCC engine self-started at all Mach numbers from 4 to 6. Pitot-pressure contours showed large flow

nonuniformity on the body-side of the inlet. The inlet provided adequate pressure recovery and flow quality for the RBCC

cycle even with the flow separation.