A Mach 5 cruise aircraft was studied in a joint program effort. The propulsion system chosen for this aircraft was an over-under turbojet/ramjet system. The ramjet portion of the inlet is to be tested in NASA Lewis' 10 x 10 SWT. Goals of the test program are to obtain performance data and bleed requirements, and also to obtain analysis code validation data. Supporting analysis of the inlet using a three-dimensional Navier-Stokes code (PEPSIS) indicates that sidewall shock/boundary layer interactions cause large separated regions in the corners underneath the cowl. Such separations generally lead to inlet unstart, and are thus a major concern. As a result of the analysis, additional bleed regions were added to the inlet model sidewalls and cowl to control separations in the corners. A two-dimensional analysis incorporating bleed on the ramp is also presented. Supporting experiments for the Mach 5 programs were conducted in the Lewis' 1 x 1 SWT. A small-scale model representing the inlet geometry up to the ramp shoulder and cowl lip was tested to verify the accelerator plate test technique and to obtain data on flow migration in the ramp and sidewall boundary layers. Another study explored several ramp bleed configurations to control boundary layer separations in that region. Design of a two-dimensional Mach 5 cruise inlet represents several major challenges including multimode operation and dual flow, high temperatures, and three-dimensional airflow effects

Coltrin, Robert E.

English

NASA Technical Reports Server

HIGH-SPEED INLET RESEARCH PROGRAM AND SUPPORTING ANALYSES 
R o b e r t  E .  C o l t r i n  
ABSTRACT 
A Mach 5 cruise aircraft was studied in a joint program effort by NASA Lewis, NASA 
Langley, and Lockheed. The propulsion system chosen for this aircraft was an 
over-under turbojet/ramjet system. The ramjet portion of the inlet is to be tested 
in NASA Lewis' 10 X 10 SWT. To test the Mach 5.0 design in this facility, which has 
a maximum Mach number capability of 3.5, the 1/3-scale inlet model is to be mounted 
under a large plate at a negative angle of attack. 
freestream flow is expanded up to the desired speed on the inlet first ramp. Goals 
of the test program are to obtain performance data and bleed requirements, and also 
to obtain analysis code validation data. The inlet was designed by using a 
conventional method of characteristics approach with boundary layer correction. 
By this means, the Mach 3.5 
Supporting analysis of the inlet using a three-dimensional parabolized Navier- 
Stokes code (PEPSIS) indicates that sidewall shock/boundary layer interactions cause 
large separated regions in the corners underneath the cowl. Such separations 
generally lead to inlet unstart, and are thus a major concern. This sidewall 
glancing shock/boundary layer interaction has been previously documented with 
fundamental research in the Lewis 1 X 1 SWT. As a result of the analysis, 
additional bleed regions have been added t o  the inlet model sidewalls and cowl to 
control separations in the corners. A two-dimensional analysis incorporating bleed 
on the ramp is also presented. 
Supporting experiments for the Mach 5 program have been conducted in Lewis' 
1 X 1 SWT. A small-scale model representing the inlet geometry up to the ramp 
shoulder and cowl lip was tested to verify the accelerator plate test technique and 
to obtain data on flow migration in the ramp and sidewall boundary layers. Another 
study explored several ramp bleed configurations to control boundary layer 
separations in that region. 
Design of a two-dimensional Mach 5 cruise inlet represents several major challenges 
including multimode operation and dual flow, high temperatures, and 
three-dimensional airflow effects. 
6-79 
https://ntrs.nasa.gov/search.jsp?R=19880006429 2020-03-20T07:53:52+00:00Z
OMGINAt PAGE 1s 
QB EOOR QUALITY 
MACH 5 CRUISE AIRCRAFT STUDY 
In 1980, a joint NASA Lewis, NASA Langley, and Lockheed California (with PdW as 
subcontractor) program was initiated. 
aircraft capable of sustained high-speed cruise, and specifically, to define the 
propulsion system for this aircraft. The final configuration from this study is 
shown in the figure. The aircraft would employ four propulsion modules (two under 
each wing). 
turbojet plus ramjet system with a two-dimensional dual flow inlet and nozzle. 
The purpose of this study was to define an 
The propulsion system chosen for this aircraft is an over-under 
MACH 5 CRUISE AIRCRAFT STUDY 
CD-87-28920 
6-80 
OVER-UNDER TURBOJET/RAMJET 
The various modes of operation for the over-under turbojet plus ramjet are 
illustrated in the figure. At subsonic flight speeds, the turbojet powers the 
aircraft, with cold flow through the ramjet. Near Mach 1.0 the ramjet is ignited, 
and both systems are operating until the aircraft approaches Mach 3 . 0 .  
2.5 and 3.0, the turbojet spools down and the upper duct of the inlet is closed 
off. From Mach 3.0 to cruise speed, the aircraft is powered by the ramjet only. 
Between Mach 
OVER-UN DER TURBOJET/RAMJET 
TURBOJET 
COLD-FLOW RAMJET 
MACH 0 TO 1 
TURBOJET & RAMJET 
MACH 2.9 
RAMJET ONLY 
MACH 3 
RAMJET ONLY 
MACH 5 CD-87-28921 
6-81 
DESIGN METHODOLOGY 
The ramjet (lower duct) portion of the inlet was designed by using a conventional 
method of characteristics (MOC) approach with boundary layer correction. One of the 
main driving factors in the design was length minimization, so as to reduce the 
weight as m c h  as possible. The cowl shock is to be cancelled at the inlet 
shoulder, and the design throat Mach number is 1.6 inviscidly, which is reduced to 
approximately 1.2 when boundary layer corrections are made. The compression split 
is about 85 percent external (with four ramps) and 15 percent internal. The inlet 
employs variable ramp geometry for off-design operation. 
t 
I 
I 
DESIGN METHODOLOGY 
DESIGNED BY METHOD OF CHARACTERISTICS WITH BOUNDARY LAYER CORRECTION. 
DESIGN MACH NUMBER = 5.0 
DESIGNED FOR MINIMUM LENGTH DUE TO WEIGHT CONSIDERATIONS. 
COWL OBLtQUE SHOCK CANCELLED AT RAMP SHOULDER. 
THROAT MACH NUMBER AT DESIGN CONDITIONS: INVISCID MT = 1.6 
WITH B.L. MT = 1.2 
COMPRESSION: EXTERNAL (85 PERCENT): 4 RAMPS 
INTERNAL (15 PERCENT): COWL SHOCK 
DISTRIBUTED ISENTROPIC COMPRESSION 
TERMINAL SHOCK 
VARIABLE RAMP GEOMETRY FOR OFF-DESIGN OPERATION 
* 
CD-87-28922 
6-82 
Mach 5 INLET AERODYNAMIC DESIGN CONFIGURATION 
, 
The aerodynamic design of the inlet is shown. X- and Y- dimensions are 
nondimensionalized to cowl lip height. Mach numbers in the various flow regions are 
shown for the cruise (Mach 5) condition. At cruise conditions, the initial wedge 
angle of the inlet is at 9". 
MACH 5 INLET 
AERODYNAMIC DESIGN CONFIGURATION 
2 . 6 1  1 . 6 7  
1 .o 
Ylhcl 
0 
0 1 .o 2.0 3.0 4.0 5.0 6.0 7.0 8.0 
Vhcl 
CD-87-28923 
6-83 
Mach 5 TWO-DIMENSIONAL INLET CHALLENGES 
At high supersonic/low hypersonic speeds, several challenges face the inlet 
designer. The problems of multiple modes of operation and dual flow must be 
solved. High temperatures must be considered. 
achieved; and the effects of three-dimensional flow, including corner flow and 
sidewall shock boundary layer interaction require complex bleed control systems. 
Inlet bleed and leakage cause serious performance penalties at the high speed 
conditions. Drag at off-design conditions is a problem, and weight reduction is a 
constant concern. 
Shock stability is not easily 
MACH 5 TWO-DIMENSIONAL INLET CHALLENGES 
ULTIPLE MODE/DUAL FLOW OPERATION 
IGH TEMPERATURE 
iOCK SYSTEM STABILITY 
LANCING SIDEWALL SHOCKIBOUNDARY LAYER INTERACTIONS 
ILET BLEED AND LEAKAGE PENALTIES 
'EIGHT REDUCTION 
DUNDARY LAYER CONTROL 
iREE-DIMENSIONAL CORNER FLOWS 
FF-DESIGN DRAG 
CD-87-28924 
6-84 
SIDEWALL BOUNDARY LAYER-GLANCING MOCK WAVE INTERACTION 
FLOW 
4 
I 
One problem with two-dimensional inlets is three-dimensional flow effects. Basic 
studies and analysis have shown that flow migration occurs in the boundary layer 
near regions where an oblique shock wave impinges on an adjacent sidewall. With 
thick boundary layers on an inlet sidewall, the higher pressure downstream of a 
sidewall-glancing shock can influence the upstream flow through the subsonic 
houndary layer. 
along the oblique shock rather than parallel to the downstream ramp surface. 
simple sketch showing the migration and analytical results are shown. 
migration has a cumulative effect, with large regions of low-energy flow sweeping up 
the sidewalls as the flow moves aft toward the cowl. 
The result of this influence is a migration of the boundary layer 
A 
This flow 
SIDEWALL BOUNDARY LAYER-GLANCING 
SHOCK WAVE INTERACTION 
SECONDARY VELOCITY VECTORS 
\ 
BOUNDARY LAYER 
CD-87-28925 LTUNNEL WALL 
6-85 
ORIGINAL PAGE IS 
OF POOR QUALITY 
THREE-DIMENSIONAL SIDEWALL SHOCK INTERACTION 
EXPERIMENTAL SURFACE OIL FILM PATTERNS 
The figure shows a simple compression wedge installed on the flat wall of a 
supersonic wind tunnel. The oblique shock from this wedge interacts with the tunnel 
boundary layer which has a thickness of approximately one inch. The surface oil 
film shows flow patterns in the boundary layer. The oil flows indicate that the 
boundary layer on the wall ahead of the oblique shock and a large portion of the 
boundary layer flow aft of the shock is influenced and is turned in a direction 
along the oblique shock angle rather than parallel to the wedge surface. 
I 
THREE-DIMENSIONAL SIDEWALL SHOCK INTERACTION 
EXPERIMENTAL SURFACE OIL FILM PATTERNS 
, 
CD-87-28926 
6-86 
THREE-DIMENSIONAL SIDEWALL SHOCK INTERACTION 
ANALYTICAL SURFACE OIL FILM PATTERNS 
A three-dimensional parabolized Navier-Stokes (PEPSIS code) solution for the simple 
case shown in the previous figure is presented below. The velocity vectors show 
that the boundary layer flow tends to turn in the direction of the shock wave. 
THREE-DIMENSIONAL SIDEWALL SHOCK INTERACTION 
ANALYTICAL SURFACE OIL FILM PATTERNS 
r WEDGE SURFACE 
/ 
/ 
TUNNEL WALL --- 
- THREE-DIMENSIONAL 
/’ GLANCING SIDEWALL/ 
CORNER SHOCKWAVE 
CD-87-28927 
6-8 7 
LEWIS HIGH-SPEED INLET PROGRAM 
TOTAL PRESSURE DISTRIBUTION 
PEPSIS analyses have shown that sidewall shock boundary layer interactions will have 
a dominant effect on the Mach 5 inlet. The figure shows total pressure 
distributions on cross-planes at several stations in the inlet aft of the cowl lip. 
Only half-planes are shown, as flow is symmetrical. 
sidewalls is captured by the cowl, resulting in a large separation region in the 
cowl-sidewall corner at a location forward of the inlet ramp shoulder. Such large 
separations will cause the inlet to unstart (expel the terminal shock out of the 
inlet), and thus are a major concern. 
Low energy flow swept up the 
LEWIS HIGH-SPEED INLET PROGRAM 
TOTAL PRESSURE DI~TRI~UTION 
, 
6-88 
LEWIS HIGH-SPEED INLET PROGRAM 
TOTAL PRESSURE 
The figure shows the total pressure distribution for a cross-section of the inlet 
just aft of the cowl lip. 
distribution, showing recirculation in the corner separation region. The influence 
of the cowl oblique shock on the mid-stream flow can also be seen. 
Velocity vectors are superimposed on this pressure 
LEWIS HIGH-SPEED INLET PROGRAM 
COWL SURFACE \ 
RAMP SURFACE 
I 
CD-87-28929 
6-89 
TWO-DIMENSIONAL KUMAR CODE ANALYSIS: Mach NUMBER CONTOURS 
A two-dimensional analysis of the Mach 5 inlet, showing the benefits of boundary 
layer bleed on the ramp and cowl was carried out by W. Rose and E. Perkins of Rose 
Engineering and Research, consultants to Lockheed on the Mach 5 project. The 
analytical code was developed by A. Kumar of the NASA Langley Research Center. The 
Mach number distribution for the no-bleed case shown on the left represents the area 
encompassed by the large dashed box in the inlet sketch at the top. These contours 
show very large, separated regions on the ramp. 
started mode with these conditions, but would operate with the normal shock external 
to the cowl lip. The Mach contours on the right, with 17 percent bleed distributed 
through the inlet, represent the area in the inlet throat region on the sketch 
above. Bleed has prevented the large separations, and a normal shock has been 
stabilized in the inlet. 
The inlet will not operate in a 
TWO-DIMENSIQNAL KUMAR CODE ANALYSIS: MACH NUMBER CONTOURS 
Y Y 
CD-87-28930 
X 
6-90 
BENEFITS TO BE OBTAINED FROM INLET TEST PROGRAM 
A 1/3 scale model of the Mach 5 inlet is to be tested in the NASA Lewis 10 x 10 SWT. 
The main goals of this test plan are to: (1) provide data for code validation and 
for the development of inlet design codes, and ( 2 )  determine overall inlet perfor- 
mance and bleed requirements. 
BENEFITS TO BE OBTAINED FROM INLET TEST PROGRAM 
PROVIDE FUNDAMENTAL AND DESIGN DETERMINE OVERALL INLET 
DATA FOR CODE VALIDATION PERFORMANCE 
GLANCING SIDEWALL SHOCK/ DESIGN AND OFF-DESIGN 
BOUNDARY LAYER INTERACTION AERODYNAMIC CHARACTERISTICS 
THICK BOUNDARY LAYERlMULTIPLE 
AND OBLIQUE SHOCK INTERACTION 
THICK BOUNDARY LAYERlNORMAL 
BLEED REQUIREMENTS 
UNSTARTlRESTART CHARACTERISTICS 
CONTROL DATA SIGNALS SHOCK INTERACTION 
CORNER FLOW 
INLET DESIGN CODE DEVELOPMENT 
VERIFICATION OF INLET DESIGN 
TECHNIQUE 
DATA VERIFICATION OF ACCELERATOR 
PLATE TEST TECHNIQUE 
CD-07-20931 
6-91 
INLET RESEARCH 
The Lewis 10 x 10 SWT has a maximum Mach number capability of 3.5. Therefore, to 
test the inlet at design conditions, the Mach 5.0 design freestream flow which is 
compressed to Mach 4.1 on the first ramp, is to be simulated by expanding the Mach 
3.5 tunnel flow to Mach 4.1 underneath a plate at negative angle of attack. This 
"accelerator plate" test technique duplicates the actual inlet flow conditions with 
the exception of the initial oblique compression shock. The data will be corrected 
for the total pressure loss for this initial compression. 
INLET RESEARCH 
. TWO-DIMENSIONAL INLET INSTALLED IN 10 x 10 SWT 
CD-87-28932 
6-92 
MACH 5 INLET MODEL 
. 
An isometric sketch of the model to be tested in the NASA Lewis 10 x 10 SWT is 
shown. The model incorporates remotely variable ramp geometry, main duct mass-flow 
control, and bleed exit plugs. The model is extensively instrumented with static 
pressure taps, total pressure rakes, translating flow angularity probes, and dynamic 
pressure transducers. It is a very large model, with the accelerator plate being 
100 in. wide and an overall model length of about 20 ft. The cowl lip height is 
16 in. with a capture width of 16 in. 
MACH 5 INLET MODEL 
INSTRUMENTATION 
STATIC PRESSURE 
TOTAL PRESSURE 
-FIXED RAKES 
-TRANSLATING PROBES 
-STAT1 C 
-TOTAL 
DYNAMIC PRESSURE TRANSDUCERS 
REMOTE CONTROL: 
VARIABLE GEOMETRY RAMP 
MAIN FLOW CONTROL 
VARIABLE BLEED EXITS (15) 
MACH 2.7-5.0 
DIMENSIONS 
PLATE WIDTH = 100 in. 
COWL LIP HEIGHT = 16 in. 
LENGTH = 20 ft. 
CAPTURE WIDTH = 16 in. 
CD-87-28933 
6-93 
ORlGlNAL PAGE IS 
OF POOR QU- 
MACH 5 INLET TEST MODEL 
Two views of the Mach 5 inlet model are shown. The model is made of stainless 
steel, except for the accelerator plate, which is aluminum. On the right, a side 
view of the inlet is shown with the sidewall removed to show variable ramp 
mechanisms. A single, large pair of actuators raises and lowers all moveable 
sections of the ramp simultaneously. The inlet duct is entirely two-dimensional, 
from leading edge to mass-flow control plug. 
prevent recirculation. 
ramp bleed through the ramp plenum. 
All bleed regions are compartmented to 
Collapsible bellows are used to duct the compartmentalized 
MACH 5 INLET TEST MODEL 
, 
6-94 c -9. 
IMPACT OF "PEPSIS" ANALYSIS ON MODEL INSTRUMENTATION AND BLEED 
As a result of the computational analysis indicating separations due to boundary 
layer migration from sidewall shock boundary layer interactions, modifications were 
made to the original model design. Additional instrumentation was added on the ramp 
and sidewall to map the flow migration phenomena and, also, to provide some code 
validation data. Additional bleed regions were added on the inlet sidewalls and on 
the cowl in the corners just aft of the cowl lip. 
control and/or eliminate corner separations. 
This bleed was added to help 
IMPACT OF "PEPSIS" ANALYSIS ON MODEL 
INSTRUMENTATION AND BLEED 
0 FIXED RAKES 
TRANSLATING FLOW 
ANGULARITY PROBES 
CODE VALIDATION INSTRUMENTATION 
-STATIC TAPS 
-FIXED RAKES 
-FLOW ANGULARITY PROBES 
-COWL 
-SIDEWALL 
ADDITIONAL BLEED 
\ /I' 
K- ADDITIONAL 
BLEED 
CD-87-28935 
6-95 
ORIGINAL PAGE Is 
OF POOR Q u m  
SMALL-SCALE PLATE/RAMP MODEL: 1 X 1 SWT 
A simple, small-scale model of the high-speed inlet was tested in the NASA Lewis 
1 x 1 SWT to provide some calibration data for comparison with analytical 
predictions. This model had capture dimensions of 1.6 in. by 1.6 in. and duplicated 
the inlet geometry to the cowl lip and ramp shoulder. Aft of these stations, the 
inlet was opened-up to allow inlet starting. The Schlieren photo on the lower left 
is for design flow conditions, and oblique shock waves from the second, third, and 
fourth ramps, as well as the cowl shock can be seen. The surface oil film 
photograph at the right is for an off-design Mach number (Mach 3.0 on the first 
ramp). 
indicated by the ramp flow near the cowl lip station. However, upstream of this 
location the sidewall boundary layer flow migration that results from the boundary 
layer-glancing shock interaction can be seen. 
For the condition shown in the photograph, the inlet was unstarted, as 
SMALL-SCALE PLATE/RAMP MODEL: 1 x 1 SWT 
TURBOJET AND RAMJET 
SCHLIEREN CD-87-28936 OIL FLOW VISUALIZATION 
6-96 
BLEED CONTROL STUDIES: 1 X 1 SWT 
An aerodynamic design approach that may be used to reduce the weight of an inlet is 
to decrease the length over which the distributed cowl compression intersects the 
ramp surface. This is accomplished by increased curvature of the cowl and results 
in a large pressure rise over a short distance on the ramp. However, these large 
pressure gradients with large approach boundary layers can result in separation. A 
simple experimental program was conducted to study ramp bleed configurations to 
control the interaction of the pressure gradient-boundary layer in this region. For 
this test, the cowl was simulated by a contoured compression plate and the ramp by 
the tunnel wall (photo on the upper right in the figure). 
incorporated a bleed plate in which various bleed patterns could be studied. A 
translating probe was used to survey the flow field. Surveys for a distributed 
porous bleed configuration are shown. Note that a separation occurs in the 
left-hand figure, due to recirculation in the bleed plenum. When the bleed is 
compartmentalized (as shown on the right) boundary layer is successfully controlled. 
The tunnel wall 
BLEED CONTROL STUDIES: 1 x 1 SWT 
COMPRESSION FIELD 
DISTRIBUTED POROUS BLEED DISTRIBUTED POROUS BLEED 
(COMPARTMENTALIZED) 
C D-07-20937 
6-97 
QpYJGE.Jj?gL. 1;'4,.5F T3 
QE E)ooR ~ J ~ ~ ~ - J ~ ~ ~  HIGH-SPEED CRUISE INLETS 
The figure shows a plot of flight Mach number compared to altitude. Wind tunnel 
models of four representative inlets for different Mach numbers and altitudes are 
shown in the photographs. In the low supersonic speed range, inlets tend to be 
simple in geometry, and often have entirely external compression. The external 
compression HiMAT inlet is shown on the left. For aircraft operating in the Mach 2 
to 4 range, inlets are usually pod-mounted and have mixed compression. Such an 
inlet, with axisymmetric geometry, is shown in the second photo from the left. This 
inlet incorporates a collapsing variable-diameter centerbody. 
range, inlets tend to be more integrated into the airframe. A two-dimensional Mach 
5 inlet model is shown in the third photo from the left. Inlets for hypersonic 
(Mach 6+) aircraft are fully integrated into the airframe, and are normally 
two-dimensional rather than axisymmetric. A scram inlet model with sidewall 
compression is shown in the photograph at the right. 
In the Mach 4 to 6 
100K 
ALTITUDE 
FT 
0 
HIGH-SPEED CRUISE INLETS 
COMPRESSION 
COMPRESSION 2D 
SIMPLE AX1 AX1 
EXTERNAL 2D 
COMPRESSION 
FULLY 
INTEGRATED 
MIXED 
COMPRESSION 
2D 
2 4 6 
, 
MACH NUMBER 
CD-87-29935 
6-98 


High-speed inlet research program and supporting analyses

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