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PERFORMLANCE CHARACTERISTICS FRO-M MACH 2.58 TO 1/98 OF A N AXISYMMETRIC MIXED-COMPRESSION INLET SYSTEM WITH GO-PERCENT INTERNAL CONTRACTION

by Robert W. Cztbbisoiz, Edwmd T. i t f elemoia, mad Dmid F. Johiason

Lewis Reseclrch Ceizter C'ieuekzizd, Ohio

N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N W A S H I N G T O N , D . C . FEBRUARY 1 9 6 9

https://ntrs.nasa.gov/search.jsp?R=19690009792 2020-04-05T08:37:04+00:00Z

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PERFORMANCE CHARACTERISTICS FROM MACH 2.58 T O 1.98 O F AN

AXISYMMETRIC MIXED-COMPRESSION INLET SYSTEM WITH

60 - PERCENT INTERNAL CONTRACTION

By Robert W. Cubbison, Edward T. Meleason, and David F. Johnson

Lewis Research Center Cleveland, Ohio

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

For s a l e by the Clearinghouse for Federa l Scientific and Technica l lnformotion Springfield, Virginio 22151 - CFSTI pr ice $3.00

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ABSTRACT

A study has been made in the Lewis 10- by 10-Foot Supersonic Wind Tunnel to determine the per formance character is t ics of an axisymmetr ic , mixed-compression inlet sys t em designed for Mach 2. 5. Inlet perforniance i s presented for Oo and maximum un- s t a r t angle of attack. Diffuser s t a t i c -p res su re distributions and compressor face total- p r e s s u r e profiles a r e shown for several operating conditions. fects of Mach number, Reynolds number, bypass flow, and vortex gene ra to r s on pe r - formance. presented.

Data presented show ef-

Per formance during r e s t a r t cycles and required a r e a r a t io s for r e s t a r t a r e

I .

i i

1 .

~ PERFORMANCE CHARACTERISTICS FROM MACH 2.58 TO 1.98 OF AN AXISYMMETRIC

MIXED-COMPRESSION INLET SYSTEM WITH 60-PERCENT INTERNAL CONTRACTION

by Robert W. Cubbison, Edward T. Meleason, a n d David F. J o h n s o n

Lewis Research Center

SUMMARY

A study had been made to determine the performance characteristics of an axisym-

The spike could be translated metric, mixed-compression inlet system designed for Mach 2.5. overall supersonic a r e a contraction occurred internally. to vary the internal contraction ratio. The inlet also included a high-response overboard bypass system for shock-position control and a low-speed valve to control secondary flow through the nacelle for cooling purposes.

Data were obtained over the nominal Mach range of 2.0 to 2.6 for angles of attack of 0' and the maximum value prior to an unstart. The Reynolds number varied from 1.24X10 to 3. 82x10 based on the inlet cowl-lip diameter.

A moderate stable operating range between design airflow conditions and inlet unstart was obtained at zero angle of attack for all Mach numbers without large sacrifices in performance. At Mach 2.5, this range was about 10.5 percent of design corrected airflow. At maximum angle of attack of about 2.7', this range w a s sizeably reduced. The overall pressure recovery level of the inlet increased, and the distortion decreased at all Mach numbers as bypass flow increased up to about 12.5 percent of the capture flow rate. Installing vortex generators on the centerbody prevented flow separation from the centerbody for bypass flows up to 40 percent of capture. Without generators, separation occurred at bypass flow ra t e s above about 12. 5 percent. For the airflow schedule assumed, the pressure recovery increased about 1 .6 percent as the Mach number w a s reduced from 2.5 to 2.3 and then decreased about 2.2 percent when the Mach number was reduced to 2.0. Above Mach 2.5, the recovery decreased rapidly. Reducing the

6 6 Reynolds number from 3 . 8 2 ~ 1 0 to 1.24X10 at Mach 2.5 reduced the maximum pressure recovery for maximum engine mass flow ratio by about 3 percent. The corresponding distortion level increased from 0.102 to 0.167.

Sixty percent of the

6 6

covery flow with low distortion to the engine. The necessary boundary-layer bleed pattern for high recovery was based on the resul ts of reference 1. With the engine installed, it is a l s o necessary to provide bypass flow capability for terminal-shock-position control

SYMBOLS

A flow a rea 2 capture area, 0.1758 m

flow area at cowl lip station bounded by model geometry AC

*cow1 b vortex generator height, 1.27 cm

2

compressor-face total-pressure profiles, and inlet restart maps are included. also shown on the effects of Mach number, angle of attack, bypass flow, vortex generators, and Reynolds number.

The test w a s conducted at Mach numbers of 1 .98 to 2.58 at zero angle of attack and at maximum angle of attack before inlet unstart. varied from 317 to 400 K and Reynolds number based on inlet capture diameter was var-

6 ied from 1.24 to 3. 82x10 .

Data are

Free-s t ream total temperature was

_ _ _ _

~~~

d

H

h

M

m/mo

N

N*

P

P

R C

Re

r

T

W & / 6

X

Ly

6

e

e Z

cowl l ip diameter, 0.4732 m

distance

annulus height at diffuser station 5

distance from centerbody surface

Mach number

mass-flow ratio

engine speed

rated engine speed, 16 500 rpm

total p ressure

static pressure

cowl-lip radius, 0.2366 m

Reynolds number

radius

total temperature

engine corrected airflow

axial distance from spike t ip

angle of attack

P/lO. 1 3 1 ~ 1 0 ~ N/m2

T/288.2 K

cowl-lip-position parameter, tan-' l/(x/Rc)

Subscripts:

av average

bl bleed

by bypass

capt capture

e j ejector

h local conditions along rake

I lip

max maximum

. I

3

min minimum

X local

0 free s t ream

5 diffuser exit

)I

, ~ ~-

APPARATUS AND PROCEDURE

The inlet used in this investigation w a s an axisymmetric mixed-compression type designed for Mach 2 . 5 with a translating centerbody for inlet start and off-design operation. At design Mach number, 40 percent of the supersonic flow area contraction was external and 60 percent w a s internal. The inlet was attached to a nacelle in which either a 5-85/13 turbojet engine or a choked-exit plug assembly could be installed. For this study, only the choked-plug assembly was used. Figure 1 is a photograph of the inlet- nacelle combination mounted from a vertical s t rut in the wind-tunnel test section. The bulge in the nacelle w a s required to house the engine accessory package.

a temperature of 390 K. number of 2 . 5 (ref. l ) , it was determined that the capture airflow would be distributed as follows: 88 .6 percent for the engine, 5. 5 percent for the throat bleed system to increase inlet performance, 4 .0 percent for engine cooling, and 1 . 9 percent for the overboard bypass to provide a margin for shock position control. The assumed engine corrected airflow schedule over the test Mach range is shown in figure 2.

External compression w a s accomplished with a 12.5' half-angle conical centerbody. Internally, the initial cowl-lip angle of 0' generated a shock that reflected twice before reaching the throat. In addition, isentropic compression w a s also included. The theoretical average supersonic throat Mach number w a s 1 .239 with a supersonic total-pressure recovery of 0 .988 . Internal contours and flow conditions in the supersonic diffuser were calculated with the aid of reference 3. The subsonic diffuser consisted of an initial throat region four hydraulic radii in length with a 1' equivalent conical expansion followed by the main diffuser that was designed as an 8' equivalent conical expansion. The subsonic diffuser length using these cr i ter ia would be 3 . 5 cowl-lip radii. However, additional length of essentially constant area was provided to accommodate the bypass system. The resulting overall design length from cone tip to compressor face was 7 . 7 2 cowl-lip radii. Fig- u re 3(b) shows the internal area distribution for several spike positions. A more complete discussion on the aerodynamic design of the inlet is contained in reference 1.

The engine match corrected airflow requirement was 1 5 . 8 3 kilograms per second at Based on an inlet pressure recovery of 0 . 9 0 at the design Mach

Some elements of the inlet aerodynamic design are presented in figure 3,

Provisions were also included for installing vortex generators on the cowl and cen-

4

terbody just aft of the throat region at 0.980 meter downstream of the design spike tip position. Details of the various generator configurations are shown in figure 4. The

the basic ordinates were used for the upper surface and the original airfoil mean-camber line formed the lower surface. The radius of the leading-edge circle was 0.0154 centi-

generator height (1.27 cm) was selected to be about equal to the local boundary-layer height. With a chosen aspect ratio of 0.5, the required chord length was 2.54 centime- ters. The angle of attack of successive generators relative to the local flow was alter- nately plus and minus 16'. This installation results in the converging (C) and diverging (D) pair orientation shown in figure 4. In general, two different vortex generator spacings were tested: a narrow space (N) of approximately three generator heights (fig. 4(a)) and a wide space (W) of about six heights (fig. 4(b)). The small difference in the spacings between cowl and centerbody generators indicated in the figure is a result of requir- ing an equal number of converging and diverging pairs on the respective surfaces. Two orientations of generator pairs (relative to the centerbody support struts) were studied for their effect on the quality of flow along the strut surfaces. Configurations NC and WC had a converging pair ahead of the struts, ND and WD a diverging pair.

An isometric view of the inlet is shown in figure 5(a). The bleed patterns (fig. 5(b)) were selected on the basis of the resul ts presented in reference 1. throat bleed) is identical to configuration A of that reference. This pattern w a s found to be the most efficient from a payload viewpoint. However, a small change in diffuser back pressure would cause the terminal shock to move completely through the throat region. To study the duct dynamics and inlet-engine interactions, it is desirable that a small disturbance produce a well-defined shock movement. Bleed in the throat region will reduce the extent the shock moves fo r a small change in back pressure. As a result, configuration I1 was derived from configuration I by adding the throat bleed indicated in figure 5(b). This bleed was positioned at about the location of the leading edge of the terminal shock t ra in during normal operation at design Mach number.

The bleed flows f rom all regions were discharged overboard. Both cowl bleed flows were discharged through exits shown in figure 5(c) which have a 15' discharge angle relative to the external surface. These exits were sufficiently large to ensure choking of the porous surface. The fore and aft centerbody bleed flows were combined into one plenum and ducted through the upper two centerbody support s t ru ts to 30' louvered exits on the cylindrical portion of the nacelle. The bottom s t ru t exit was sealed because the nacelle bulge (fig. 1) covered about two-thirds of the opening and the pressure r i s e due to this bulge may have unchoked this exit. The exit areas of the remaining two s t ruts were large enough to ensure choking of the bleed holes.

1- - - - - _ _ - _- - uaviC geiieiat~r shape F;~S derii;& from the XACA 0012 airfoil. T;; cne (fig. 4(aj)

iiieiei.. F ~ ~ - tfle otflei- conf;g~rz~;tion, tL,. --.--iAtA A T A P A nn in n:n *

L l l C b U l l A p l C L C I'IAbA V U l O C & L L f 6 1 1 w%s used. The

Details of the model configurations used in this investigation a r e shown in figure 5.

Configuration I (no

.

5

The inlet design included two remotely controlled flow bypass systems: a high re - 8

sponse overboard system for shock position control, and a low-speed valve to control secondary flow through the nacelle for engine and ejector cooling. Details of the bypass plenum entrance are shown in figure 5(c). The entrance slot to the plenum was sized to pass 94.5 percent of the capture mass flow at the design Mach number assuming a total p ressure at the slot entrance equal to the main duct (or compressor face) static pressure . Based on the air management schedule, the flow would then be discharged overboard through 45' louvered door units (fig. 5(c) and (d)) and also through the ejector valve openings into the nacelle (fig. 5(c) and (e)). The overboard bypass exit area was sized to remove 88 percent of the on-design capture mass flow assuming again that main- duct static pressure equals the total p ressure ahead of the choked bypass exit. The remaining flow (6.5 percent of capture) would be discharged through the ejector valve. However, the design capability of this valve was 10 percent of capture mass flow assuming main-duct static pressure at the entrance of the choked valve. trance and plenum were divided into six separate sections by the centerbody support s t ru ts and cowl ribs. per section) were symmetrically located around the periphery of the bypass plenum. An equal number of ejector valve openings were similarly located in the base of the bypass plenum. Besides the steady-state mode of operation used in the present study, the overboard bypass system could be operated sinusoidally at frequencies up to 140 hertz with a flat amplitude response of 0.254 centimeter peak-to-peak at frequencies up to 100 hertz.

The overall inlet total-pressure recovery just ahead of the compressor face w a s determined from rakes l to 6 (fig. 5(f)) which were area weighted. Additional measurements (rakes 7 to 10) were included in the distortion calculations and in plots of the total- p ressure profiles. Static-pressure distributions along the top centerline of both cowl and centerbody were also measured.

plug and an average of eight static pressures in the cold pipe located 4. 3 cowl-lip radii ahead of the plug exit. Cowl bleed flow rates at both stations were determined from a measured static and total pressure at their respective exits (see fig. 5(c)) and the measured exit areas. Flow from the combined centerbody bleed regions w a s calculated from the Pitot-static measurements that were taken by the rake in the centerbody cavity just ahead of the support struts. sonic flow relation using the measured bypass plenum total pressure (see fig. 5(c)), the measured valve opening, and experimentally determined flow coefficients. The overboard bypass mass-flow ratio at 0' angle of attack was then determined by subtracting the sum of the engine, ejector bypass, and the three bleed mass-flow ratios f rom the theoretical capture mass-flow ratio for the spike position and Mach number in question.

Both the bypass en-

To ensure uniform flow removal, six overboard door units (one

The diffuser exit or engine mass-flow ratio was calculated using a calibrated choked

Ejector bypass mass-flow ratio was calculated from the

. The test was conducted in the Lewis 10- by 10-Foot Supersonic Wind Tunnel f rom

6

, Mach 2.58 to 1.98 at zero angle of attack and at maximum angle before unstart. The *

6 Reynolds number based on inlet cowl-lip diameter varied from 1.24 to 3.82x10 . Free- s t ream total temperature w a s varied f rom 317 to 400 K. Temperature levels above the normal tunnel operating levels were obtained using the natural gas fired heater installed in the tunnel bellmouth upstream of the supersonic nozzle. On the aerodynamic cycle (normally used for isolated inlet testing without heat addition) the normal temperature levels are 317 K for Mach 2 .0 to 2 .5 and 372 K for Mach 2. 5 to 2.6. The 55 K step change at Mach 2.5 resul ts from the startup of the second tunnel compressor, which is required for tunnel operation above Mach 2 . 5 . For the propulsion cycle (used for engine tests, tunnel heater operation, or both) the normal unheated levels at all Mach numbers a r e reduced about 25 K. A potential disadvantage of using a gas-fired heater installed in the s t ream is the possibility of flow nonunifor mities resulting from condensation of the water vapor that is introduced into the flow. It was found in reference 4 that in the 10- by 10-Foot Supersonic Wind Tunnel the tes t section Mach number and total pressure were decreased by this effect. For example, at Mach 2.5 the temperature rise from 317 to 373 K resul ts in a Mach number reduction from 2.50 to 2.47. At Mach 2. 3, the temperature increment of 68 K reduced Mach number by 0.063, and at Mach 2.0, the Mach number decrement was 0.032 from a temperature increase of 30 K. A s discussed in reference 4, these changes in flow conditions were repeatable and were accurately calibrated. Of pr imary interest in the present study is this Mach reduction that could influence the inlet performance. The optimum inlet geometry (spike position) was determined for the unheated test condition. This geometry was held fixed for both the heated and unheated conditions. At the nominal Mach 2.3 test condition, the optimum inlet configuration corresponded to a near-maximum contraction ratio and, as a result, the Mach number reduction due to heater operation was sufficient to unstart the inlet. quently, the tunnel wall was positioned at the next higher nominal setting (M = 2.4) prior to the addition of heat. The resulting free-stream Mach number was 2.34.

Conse-

RESULTS AND DISCUSSION

Throughout the following discussion, the nominal value of Mach number will generally be used. The actual free-stream values a r e indicated in the figures and when used in the text will be referred to as the free-s t ream Mach number. *

The general presentation of the basic performance data of the inlet system is as fol-

figures 6 and 7; (2) cowl and centerbody static-pressure distributions in figures 8 and 9, respectively; (3) compressor-face total-pressure profiles for various operating conditions in figures 10 to 13; (4) total bleed flows in figure 14; (5) ejector bypass operating

lows: (1) mass-flow ratio, total-pressure recovery, and distortion characteristics in 4

7

I .characteristics in figure 15; (6) performance characterist ics during restarts with the corresponding diffuser static-pressure distributions and compressor-face total-pressure profiles in figures 16 to 19. Following this general presentation of the basic data, com- parisons w i l l be presented showing the effects of (a) vortex generators on total-pressure profiles during bypass operation; (b) bypass flow on peak recovery and distortion; and (c) Mach number and Reynolds number changes. In addition, the experimental and predicted restar t area ratios for this inlet configuration will be compared.

Performance D u r i n g Normal Operation

The performance characteristics of configurations I and IIND' without bypass flow and that of configuration IIND' with varying amounts of bypass flow are shown in figures 6 and 7, respectively, for the various Mach numbers, angles of attack, and free-s t ream temperature conditions. In figure 6, both configurations indicated a high performance capability at Mach 2.5. However, it was noted for configuration I that a small change in plug position caused the terminal shock to move rapidly through the throat region. With the throat bleed of configuration IIND', a small change in plug position produced a slower shock movement that w a s desired for a la ter study of shock-position dynamics. Conse- quently, the IIND' configuration was selected for the future phases of the overall inlet- engine investigation. Thus, the majority of the data presented a r e for configuration IIND' and only limited data a r e available for configuration 1.

variation without bypass flow (fig. 6) were obtained by simulating various engine speeds with the choked exit plug. With the bypass system functioning (fig. 7), the inlet character is t ics were obtained both by maintaining constant bypass position and varying plug position and by maintaining constant engine corrected airflow and varying bypass flow. A comparison of figures 6 and 7 shows an increase in engine face pressure recovery when the bypass system was operating. This is primarily due to the removal of the cowl side boundary layer by the bypass. The intersections of the corrected airflow lines and the engine speed simulation curves were of particular significance for subsequent engine tests since they simultaneously matched the engine airflow demand and positioned the terminal shock such that it had a slight influence on the downstream row of centerbody bleed holes. Tn this manner, the shock w a s about the same distance downstream of the geometric throat at all Mach numbers. This terminal shock position for normal operation was based on the shock movement desired in the duct dynamics program. The simulated engine speed line was established by varying the plug position with the overboard and ejector bypass openings constant at the match point conditions. The bypass line w a s

The mass-flow, pressure-recovery characteristics and the corresponding distortion

8

determined by holding the plug position and ejector bypass settings constant and vmying .

the overboard bypass setting, As evident f rom figures 6 and 7, the inlet exhibited a moderate mass-flow range be-

tween the engine match point and inlet unstart at all Mach numbers for an angle cf &.ac'ti

of 0'. change in corrected airflow between the performance levels with the terminal shock in the normal operating position (m5/m0 = 0.944 and P5/Po = 0.895) and the conditions just before inlet unstart (m5/mo = 0. 889 and P5/Po = 0.928). With bypass flow (fig. 7) this range was increased to about 10.5 percent. This increase w a s a result of more flow being removed through a given bypass opening when the pressure recovery increased as the terminal shock was moved forward in the inlet. At maximum angle of attack (about 2.7'), the stable range w a s considerably reduced for both configurations studied.

operating conditions. Reducing the pressure recovery from peak inlet operation increased the unstart angle of attack, a,,, until a maximum value w a s reached. decrease in pressure recovery had no effect on the maximum unstart angle of attack. As Mach number was reduced from 2.6 to 2.3, both the aUn at peak pressure recovery and maximum flun decreased. At Mach 2.6 with the centerbody translated forward to keep the tip shock ahead of the cowl lip, the contraction ratio w a s reduced well below the maximum allowable. Data at Mach 2. 3 and control-room observations at Mach 2 .0 indicated that the tolerance to angle of attack or yaw w a s reduced to l e s s than 0.72'. This occurred at all operating conditions with the spike retracted to a position creating near maximum contraction ratio in the inlet at the Mach number in question. The loss was probably due to a combination of this contraction ratio and a mismatch in the cowl and centerbody contours at off-design spike positions. As a result, no data were obtained for normal operating spike positions at both Mach 2. 3 and 2.0, and only the operating maps at maximum aun at Mach 2 .5 and 2.6 a r e shown in figure 7 . It w a s also observed that translating the centerbody forward would increase tolerance to angle of attack effects but with a sacrifice in performance. This performance loss may be acceptable during short-duration maneuvers when relatively large angles of attack and/or yaw may be en- countered.

Also evident, at Mach 2.5 and 2.6 (fig. 7(a) and (b)), w a s a discontinuity in the mass-flow, pressure-recovery characteristic near a mass-flow ratio of 0. 87. This w a s due to a centerbody boundary-layer separation that occurred when the terminal shock was positioned on the porous surface ahead of the geometric throat. This occurrence is discussed in more detail in reference 1.

the Mach range studied and for a = 0' are presented in figures 8 and 9, respectively. These distributions were obtained for the normal mode of operation; that is, operating

For example, at Mach 2.5, ~.Jithoiii bypass (fig. 6) i t provided about a 9-percent

The angle of attack at which the inlet unstarted is shown in figure 6 for different inlet

Further

The cowl and centerbody static-pressure distributions along the top centerline over

'.

9

along the match corrected airflow line by varying the overboard bypass flow. The data obtained with and without tunnel heat were quite similar; therefore, for comparison purposes only those for elevated temperatures at Mach 2.5 a r e shown. The static-pressure distributions were used to provide signals for a shock-position control using the high- response bypass doors in subsequent tests.

to the flagged points in figure 7 . Only those most closely corresponding to the engine match operating conditions a r e shown for Mach 2.6, 2. 3, and 2.0. At Mach 2.5, several operating conditions a r e shown. For comparable pressure recoveries resulting from either bypass or plug position variations, figures l l ( c ) and (f) or figures l l (d ) and (g) indicate very little difference in the general shape of the profiles. Generally, the differences between hot and cold tunnel data can be attributed to the difference in Mach number due to heater operation discussed in the apparatus and procedure section. In addition, the spike position used with heater operation was optimum for the cold tunnel condition and, hence, was not necessarily optimum for the Mach number resulting from heater operation. In general, the distortion is primarily in the radial direction with only a small circumferential variation.

figure 14 for the simulated engine speed variations shown in figure 7 . Details of the flow ra tes i n each of the bleed passages a r e given in reference 1. The pressure recovery of the inlet is a function of the terminal shock position which, in turn, directly affects the driving pressure on these choked bleed patterns. Consequently, the flow rates would be the same for similar shock positions regardless of whether the shock was positioned by bypass or plug position variations.

shown in figure 15. These data can be used to determine the ejector flow ra te at any inlet operating condition. flows for any operating condition from the capture flow will give the overboard bypass flow at that condition.

The compressor-face total-pressure profiles shown in figures 10 to 13 correspond

The total bleed flow variation with overall inlet pressure recovery is presented in

The operating characteristics of the ejector bypass used in this inlet system a r e

Subtracting the sum of the ejector, bleed, and engine mass

Performance D u r i n g Restart Cycle

Overall inlet pressure recovery and distortion during the restart cycles for the Mach range of 2.6 to 2.0 are shown in figure 16. static-pressure distributions for the Mach 2.5 res ta r t cycle are shown in figures 17 and 18, respectively, with the engine-face total-pressure profiles presented in figure 19. The data of figure 16 are plotted against the combined engine and overboard bypass mass flow. These data (other than started conditions at design e,) were obtained by varying

Corresponding cowl and centerbody

10

the bypass with the exit plug set for engine match airflow. An unstart at the design Mach number (figs. 16(b) and (c)) reduced the pressiii*e recovery by about 50 percent and approximately doubled the distortion. In general, the unstarted inlet distortion level decreased as the centerbody was translated forward to res tar t . After res tar t , the distortion level was about the same as fo r the csrmai vperating spike position. The reversal in this trend for unstarted operation at O 1 = 22. 87' is a result of a separation region in the diffuseF. As shown by the total-pressure profiles of figures 19(c) and (d), these data plus other unpublished data for this inlet have shown a random circumferential location of this separation region. cated in the lower left (rake 3) in figure 19(c), and a tendency for separation is indicated on top (rakes 5 to 7) in figure 19(d). When this separation region was in the top of the inlet, evidence of it can be seen from the static-pressure distributions by comparing figure 17(c) with 17(d) and figure 18(r) xitL 18(d). This flow separation may be associated with an unsymmetrical shock structure that was noted from schlieren observation of the conical section of the centerbody at some spike extensions with the inlet unstarted. In general, a well-defined secondary shock can also be seen in the static-pressure distributions of figures 17 and 18. These data, along with centerbody position, can provide the required signal for bypass door position control during a res ta r t cycle. With such a control , it should be possible to minimize the distortion and maximize the pressure recov- e ry during a res ta r t cycle. portion of the cycle (fig. 19(a) to (d)) show that the distortion is primarily circumferential. Just before res ta r t (figs. 19(d) and (e)), and after res ta r t (figs. 19(f) and (g)), the distortion appears to be primarily radial as previously shown in figures 10 to 13.

For the two se ts of data shown, a definite separation is lo-

The engine-face total-pressure profiles during the unstarted

Effect of Vortex Generators

In other recent inlet studies (e.g. , refs. 5 to 7), vortex generators have been used in short subsonic diffusers to reduce distortion levels. However, in the present inlet, their main function was to maintain attached centerbody flow at the engine face during bypass operation at moderate flow rates. In all the configurations studied (fig. 4), the generators were equally spaced around the periphery and oriented to create counter- rotating vortices. The effect of the counter rotation was to create, in the mixing zone, regions of high pressure behind the diverging (D) generator pairs. These regions, because of the mixing action, dissipate with distance. Consequently, the generators must be located f a r enough ahead of the compressor face for this mixing of the flow to provide the desired pressure profile f ree of discrete vortices. Reference 5 reports that the vortex patterns were greatly dissipated after a distance corresponding to eight pair spacings (distance between centers of D pairs). In the present investigation, the gen-

11

I I .erators were spaced to make the distance from the generator station to the compressor

face correspond to 4.9 and 9 .8 divergent generator pair spacings for the W and N configurations, respectively. A comparison of the compressor-face total-pressure measurements at peak inlet operation behind the various vortex generator orientations with C and D generator pairs ahead of the s t ruts is shown in figure 20. The differences in the profiles for the wide-spaced configurations (WC and WD, fig. 20(a)) indicate the pres- ence of discrete vortices in the flow at the measuring station. This is primarily shown by rakes 7 to 9. It is also evident near the centerbody surface for rakes 5 and 6 where the low-pressure region behind a C pair of generators and the high-pressure region behind a D pair are quite obvious. The apparent lack of discrete vortices on the cowl side ahead of rakes 5 and 6 is a result of the angular location of the generators relative to the rakes (fig. 5(b)) rather than sufficient duct length for complete mixing. Due to the counter-rotating action, the centerlines of the vortices behind a D generator pair move toward each other. Consequently, they move apart behind a C pair of generators. As a result, the low-pressure region should be wider at the rake station than the high-

C pair was apparently sufficient to influence the rake measurements in the manner ex-

I

I

I l

I I

l

I

I pressure region. For configuration WD the spread of the low-pressure region behind a 1

pected behind a C pair. Therefore, to determine whether discrete vortices are still I present, the comparison should be made of measurements from a rake centered behind

1 I

I

first one type of pair (C o r D) and then the other. As shown in figure 20(b) the profiles 1 with the close-spaced generator configurations (fig. 4(a)) installed, indicate that the flow is quite well mixed and very little evidence of discrete vortices remain. How much of the difference in the profile was due to the vortex generators or attributable to the difference in bleed configuration is not apparent. From the data presented in figure 20, it is obvious that when generators are installed, the area weighted average total-pressure recovery from stationary rakes can be greatly influenced by their location relative to the generator pairs. If sufficient mixing length is provided, the effect can be eliminated. In the present inlet system, sufficient length was provided with the narrow spacing but not with the wide spacing.

,

I I

In the present configuration, a desirable compressor-face total-pressure profile was obtained with only the centerbody generators installed since the bypass uniformly re -

stalled, unpublished data indicated the total-pressure losses were about 2.7 percent. With just the centerbody generators about half of this loss would be eliminated. A comparison of typical total-pressure profiles with and without centerbody vortex generators is shown in figure 21 for several bypass flow rates. Without vortex generators, the profiles indicate the centerbody flow was near separation at a bypass flow ra te of 12. 3 percent of capture mass flow. From control-room observations, it w a s noted that separation occurred above 12.5 percent of capture mass flow. With vortex generator config-

I moved the cowl side boundary layer. When both cowl and centerbody generators were in- I I

l

I

I I

12

uration ND' (fig. 4(a)) installed, the centerbody flow remained attached with bypass flows to about 40 percent of the maximum c.a-pt??re flow rate. At. the same time the distortion ievel was reduced to approximately half the level without generators. The effect of some additional bleed is also included in these results. However, it is felt that a high percentage of the increased bypass capability w d rcchiceii distortion w a s due to the vortex generator action and only a small amount due to the additional bleed.

,

Effect of Bypass Flow

The compressor-face total-pressure recovery (with or without vortex generators) was affected when the cowl side boundary layer was removed by the bypass system. The effect of bypass flow ra te on the peak performance of configuration IIND' is shown in fig- u re 22. The inlet would not be normally operated at the peak condition because of unstart problems. However, the comparison is made at this operating point since, for the various bypass openings, the te rmina l shock is easily positioned in the same location relative to the bleed system. In this manner, the resul ts shown are due to bypass alone and a r e not influenced by a trade-off between bleed and bypass. At all Mach numbers tested, the overall pressure recovery increased as the bypass flow was increased from 0 to about 12.5 percent of capture. For example, at Mach 2.5 the increase was from 0.92 to 0.955. Increasing the amount of bypass beyond 12.5 percent resulted in a recovery decrease. The distortion level decreased for about the same initial bypass flow range. At Mach 2 .5 the decrease was from 0.08 to 0.042. The decrease at all Mach numbers was, i n general, followed by an increase to about the original level or higher for bypass flow rates up to that causing centerbody flow separation (approximately 0.40 bypass mass-flow ratio). After flow separation occurs, the distortion level decreases because the separation region encompasses an increasingly larger portion of the compressor face. The one exception is at the design Mach number where the distortion level remains essentially constant f rom about 12.5-percent bypass up to maximum bypass; however, the characteristic shape of the curve of the other Mach numbers is apparent.

Effect of Mach Number

The effect of Mach number on the performance parameters of the inlet system is shown in figure 23. The bypass exit a reas (when the bypasses were operating) were set at the match point values required at the Mach number in question except for Mach 2.0 without tunnel heat. As the Mach number was increased above design, the overall recovery decreased rapidly. Reducing the Mach number from design (2.5) to 2.3 increased

-the recovery at engine match and at peak conditions by approximately 0.02 with the bypasses operating. Mach number reduction to 2.0 resulted in a 0.020 recovery loss at match conditions and 0.010 loss at peak conditions. Similar trends are shown for the configurations without bypass with the maximum recovery occurring at Mach 2.45. erally, these data follow the theoretically predicted throat-exit pressure-recovery variation with Mach number. The scatter at the lower Mach numbers was probably a result

The maximum avail-

I I

Gen- I

I of the differences in the overboard bypass setting at Mach 2 . 0 and of using the lower I

f ree-s t ream Mach number inlet contraction settings at Mach 2. 3.

able mass-flow ratio shown in figure 23 (capture minus minimum bleed flow f rom fig. 14) is that corresponding to the spike positions shown in figure 7. These data are for the by-

The difference between the supply and the engine demand is the required total bypass flow. As shown, the supply exceeds the demand by about 0.06 over the Mach range. The change in mass-flow ratio between engine match and peak pressure recovery was a combination of increased bleed due to terminal shock movement and more bypass due to higher overall inlet recovery. The decrease in maximum available flow at less than design Mach number w a s due to oblique shock spillage caused by a combination of centerbody translation and of increases in conical shock angle. Above design, the increase was due to a decrease in bleed flow caused by a combination of the movement of the internal shock structure relative to the bleeds and the reduced static pressure level in the throat bleed region. The distortion levels at match conditions (fig. 23) indicate a maximum at Mach 2.5 of about 0.11, slightly less at Mach 2.6, and decreasing to 0.08 at Mach 2.0. At peak recovery, the distortion levels indicate different trends with Mach number depending on whether the tunnel heater is used. Without tunnel heat, the distortion minimizes at about 0.04 near the design Mach number. At Mach 2.6 and 2.0, the values were 0.08 and 0.066, respectively. This trend might be anticipated since the diffuser contour i n this translating centerbody inlet is on design at Mach 2. 5. With tunnel heat, the distortion level is generally higher (about 0.07) and tends to be the highest around the design Mach number and lower on both ends of the Mach range studied. With the exception of the distortion levels just discussed, the variation in the other two pri- mary inlet parameters (capture mass flow and overall recovery) was consistent with Mach number changes with o r without tunnel heaters.

pass opening corresponding to the match point position for the Mach number in question. 1 I ,

,

Effect of Reynolds Number

6 6 The results of changing Reynolds numbers f rom 3.82X10 to 1.24X10 (based on cowl-lip diameter) are presented in figure 24. To determine the difference in inlet performance at the different Reynolds numbers, the effect of Reynolds number on the tunnel

14

f ree-s t ream Mach number must be considered. For the two cases in question, the free-’ s t ream Mach nuiiiber was 2.50 z(! 2 . 4 5 h r the high and lower Reynolds number, respectively. Based on the trends with Mach number already discussed, a change in both mass- flow ratio and overall pressure recovery would be expected. The data shown i n f i p re 23 indicate that a Mach decrease cf about 0.04 should produce about a 0.018 decrease in the super critical diffuser-exit mass-flow ratio. Based on these observations, the data of

6 6 figure 24 show that a reduction in Reynolds number f rom 3 . 8 2 ~ 1 0 to 1 . 2 4 ~ 1 0 did not affect the inlet mass-flow ratio. Similarly, accounting for the effect of Mach number on pressure recovery from figures 23 and 6, a difference of about 0.015 at maximum recov- e ry for maximum engine mass-flow ratio would be expected. Applying these resul ts to the level at this condition shown in figure 24, it can be concluded that this Reynolds number reduction resulted in about a 3-percent reduction in overall total-pressure recovery. The changes in distortion level reflect the effect of Reynolds number. At the same operating condition, the distortion increased from 0.102 to 0.167 as the Reynolds number

6 6 was reduced from 3.82xlO to 1.24X10 . This loss of performance at the lower Reynolds number may be a result of an increased shock boundary-layer interaction at the shock impingement points because of increased boundary-layer thickness and a delay in transi- tion, An indication of an unfavorable change in the interactions at these points was observed from unpublished centerbody and cowl static-pressure distributions. However, the exact nature of these changes could not be determined from these distributions.

Experimental and Predicted Restart Area Ratios

In a variable-geometry, mixed- compression inlet, the initial start or any res ta r t s of supersonic flow in the inlet is usually accomplished by increasing the ratio of the throat a r e a to the entering flow area. Mixed-compression inlets have been found to r e - start with less throat a r ea than would be predicted from a one-dimensional flow analysis assuming uniform flow in front of a normal shock at the cowl lip. Experimentally determined internal a r e a ratios required to restart the present mixed- compression inlet are presented in figure 25. As shown, this inlet also res ta r t s at area ratios well below the normal shock values. Similar results were obtained in reference 8, which presents an analysis of restart data on a Mach 3.0 mixed-compression inlet. The reference also suggests a prediction technique based on measurements of the unstarted flow field at the point of restart. Reference 8 states that the restar t flow a r e a ratio (ratio of minimum geometric a r e a to the measured effective flow area at the cowl lip) remained constant at a value of 0 .81 for all spike extensions. This value w a s 0.94 of the theoretical contraction ratio for normal shock flow conditions at the cowl lip and about 1.03 t imes the theoretical contraction ratio for the miltiple-shock flow field resulting from the oblique

15

'shock induced by the centerbody boundary-layer separation. To estimate r e s t a r t area ratios for the present inlet over the Mach number range, it was assumed that the restart flow a reas maintained the same relation to the theoretical multiple shock and normal shock contraction ratios of 1 . 0 3 and 0 . 9 4 , respectively. Reference 9 was used to pre- dict the separation wedge angle required to determine the local flow conditions in the multishock region. A s shown in figure 25, the predicted values a r e 2 percent low at the design Mach number of 2 . 5 , 1 . 3 percent low at Mach 2 . 3 , and identical to measured values at Mach 2 . 0 . The results indicate that this technique, as suggested by reference 8 along with the assumptions used in the present analysis, accurately predicted the required restar t a r e a ratios for the present inlet over the test Mach number range. The experimental results also compare quite favorably with a no-bleed, all-internal- compression inlet designed for Mach 3 . 0 . The area ratios shown for that inlet were obtained from reference 8. With the spike translated forward to the res ta r t position, the present Mach-2.5 inlet can also be considered a no-bleed configuration because the cowl bleeds a r e well downstream of the throat and the forward centerbody bleed removes enough flow so that the actual res ta r t minimum area, Amin, occurs at the leading edge of the forward centerbody bleed.

SUMMARY OF RESULTS

A cold-pipe investigation was conducted to determine the performance characteris- t i cs of an axisymmetric, mixed-compression inlet system designed for Mach 2 . 5 and to match the 5-85/13 turbojet engine. Sixty percent of the overall supersonic a r e a contraction occurred internally. The spike could be translated to vary internal contraction ratio. The inlet also included two operating bypass systems: a high-response overboard system for shock-position control and a low-speed valve to control secondary flow through the nacelle for engine and ejector cooling.

varied from 1 . 2 4 ~ 1 0 to 3 . 8 2 ~ 1 0 based on inlet cowl-lip diameter. Both zero angle of attack and the maximum angle before an inlet unstart occurred were studied. The following results were obtained:

1. At zero angle of attack, a moderate stable operating range between match conditions and inlet unstart was obtained at all Mach numbers without large sacrifices in performance. At Mach 2 . 5 , this range amounted to about 1 0 . 5 percent of design engine corrected airflow. At maximum unstart angle of attack of about 2. 'lo, this range was sizeably reduced.

s u r e recovery and reduced the distortion at all Mach numbers studied.

Data were obtained over the Mach range of 1 . 9 8 to 2 . 5 8 . The Reynolds number 6 6

2 . Flow bypass of up to 1 2 . 5 percent of capture mass flow increased the inlet pres-

16

I I 3. Use of vortex generators on the centerbody prevented flow separation from the '

centerbody for bypass flows up to about 4G percent of capture. 4 . As the Mach number was decreased from 2 . 5 to 2 . 3 (with bypasses open to match

point exit areas) the match pressure recovery increased by about 1 . 6 percent. Red~cing the Mach number f rom 2 . 3 to 2. Q resulted in a 2 . 2 percent reduction in the match- condition pressure recovery. Above Mach 2 . 5 the pressure recovery decreased rapidly.

slightly less at Mach 2 . 6 , and decreased to 0 . 0 8 at Mach 2 . 0 .

2 . 5 was to reduce the overall pressure recovery by approximately 3 percent at maximum recovery for maximum engine mass-flow ratio. The corresponding distortion levels increased from 0 . 1 0 2 to 0.167. No change in diffuser exit mass-flow ratio was indicated.

quired to pass the flow through a simple normal shock. A technique suggested in reference 8 predicted restart a rea ratios that were within 98 percent of the experimentally determined values.

I I ,

5. At engine match conditions, the distortion was a maximum of 0 . 1 1 at Mach 2 . 5 ,

6 6 6. The effect of reducing the Reynolds number f rom 3.82X10 to 1.24X10 at Mach

7. The inlet res tar ted at all test Mach numbers with throat a r eas smaller than re-

Lewis Research Center, National Aeronautics and Space Administration,

Cleveland, Ohio, October 2, 1968, 126-15-02-11-22.

REFERENCES

1. Cubbison, Robert W.; Meleason, Edward T.; and Johnson, David F. : Effect of Po- rous Bleed in a High-Performance Axisymmetric, Mixed- Compression Inlet at Mach 2 .50 . NASA TM X-1692, 1968.

2. Sanders, Bobby W. ; and Cubbison, Robert W. : Effect of Bleed-System Back Pres- su re and Porous Area on the Performance of an Axisymmetric, Mixed-Compression Inlet at Mach 2 . 5 0 . NASA TM X-1710, 1968.

3. Sorensen, Virginia L. : Computer Program for Calculating Flow Fields in Supersonic Inlets. NASA TN D-2897, 1965.

4. Cubbison, Robert W. ; and Meleason, Edward T. : Water Condensation Effects of Heated Vitiated Air on Flow in a Large Supersonic Wind Tunnel. NASA TM X-1636, 1968.

17

5. Mitchell, Glenn A. ; and Davis, Ronald W. : Performance of Centerbody Vortex Gen- e ra tors in an Axisymmetric Mixed-Compression Inlet at Mach Numbers F rom 2.0

t to 3.0. NASA T N D-4675, 1968.

6. Sorensen, Norman E. ; and Smeltzer, Donald B. : Investigation of a Large-Scale Mixed Compression Axisymmetric Inlet System Capable of High Performance at Mach Numbers 0.6 to 3.0. NASA TM X-1507, 1968.

7. Sorensen, Norman E.; Cubbison, Robert W.; and Smeltzer, Donald B. : Study of a Family of Supersonic Inlet Systems. Paper No. 68-580, A M , June 1968.

8. Mitchell, Glenn A. ; and Cubbison, Robert W. : An Experimental Investigation of the Restart Area Ratio of a Mach 3.0 Axisymmetric Mixed Compression Inlet. NASA TM X-1547, 1968.

9. Nussdorfer, T. J. : Some Observations of Shock-Induced Turbulent Separation on Supersonic Diffusers. NACA RM E51L26, 1954.

-- - --'~-'-----' -------

Figure L - Model installed in 10- by la-foot Supersonic Wind Tunnel test secti on_

u ~ ~ ~-~

'" "C

.!!l u ~ ~

0 u

'" c en c ....

19

18

I I I .1 1 I I I ~O = 322 K; (N/N"ve) x 100 = 94.7

I I

I'--- I I

17

......... '"'-.. ,rIO = 360 K;

'-c '""'" (N/N* ve) x 100 = 89.9

16 ~ TO = 390 K; ). ~ ~

(N/N"ve) x 100 = 86.9-' ~~

15 2. a 2. 1 2.2 2.3 2.4 2.5 2.6

Free-stream Mach number

Figure 2. - Assumed engine airflow schedule selected for inlet-engine interact ion studi es.

19

3 al

oundary-layer blee

Cowl surface condit ions

d In le t contour

?2 0 20 3 2:

3

c

.- ;E 0 c 1

2 0 . 5 1.0 1 .5 2.0 2 .5 3 . 0 3 . 5 4 .0 4 . 5 5 . 0 5 . 5 6.0 6 . 5 7.0 7 . 5 8 .0

Axial distance, dRC, in le t rad i i

Centerbody surface condit ions

(a) In le t dimensions and theoretical flow conditions,

Free-stream Cowl-l ip Mach position

number, parameter, OZ,

deg MO

A 2.02 20.6 B 2.20 23.87 C 2.50 26.6 D 2.58 26.23

Diffuser station, x/Dc, cowl- l ip diameter

(b) Diffuser area variation.

Figure 3. - Aerodynamic details.

20

D C

Looking downstream Narrow-converging, NC Narrow-diverging, ND

(cowl and centerbody) (cowl and centerbody) Narrow diverging, ND' (centerbody only)

Configurations

Surface coordinates from /' NACA 0012 air fo i l data

-\-

0.0154 c m rad

b = 1.27 cm

Generator details

Converging pair (C) Diverging pair (D)

Generator orientation

(a) Narrow spaced vortex generators for modified NACA 0012 air fo i l .

Figure 4. - Vortex generator design.

2 1

W ide-converging Looking downstream

W ide-diverging

Configurations

-7- r~~~~ 0012 a i r fo i l I <a- 0.3048 cm

b = 1.27 c m

Generator details

5.92 b on cowl side 6.22 b on centerbody

Flow Flow - - 16%

/ - Converging pair ( C ) Diverging pair (D)

Generator orientat ion

(b) Wide spaced vortex generators for NACA 0012 air fo i l .

Figure 4. - Concluded.

22

V I .-

23

Aft cowl m m

I1 0 0 Geometric a m Distance from spike t ip, xlRc <3.23 3.25, throat ~

7 Experi- / \ 1 mental 1 shock

1 ments Flow- / ,/

Row

I1 o o a m a Open

Configuration r-----, I nom m 8

Forward Aft a Half open centerbody centerbody m Closed

(b) Bleed configuration. Bleed hole diameter, 0.3175 centimeter.

Compressor Station 5 face station

Overboard 1.7498 m 1.8255 m Station 0

0

Bypass opening r1 = 1270 cm r2 = 0.635 cm r 3 = 22.88 cm f3 =ao y = 45"

I /'

Cowl bleed-exit ', ,' static pressure>' Centerbody bleed s t ru t

louvered exits (2 struts used) J

CD-9063 (c) Pressure instrumentation and bypass entrance dimensions (all stations referenced to Mo . 2.498 centerbody location.).

Figure 5. - Continued.

24

Hyd rau lic actuator -,

Pos ition sensor ""' "

Electric drive motor~" , , ,

, , ,

(d) Overboard bypass door assembly and actuator.

(e) Ejector bypass valve and actuator.


CD-9062-01

~ I I

! / I I

:' / , / / : L Valve opening (one

of six equally spaced around ring)

CD-9061-Dl

I I

25 I

J

S t r u t /

Rake 1

Stat ic- p ressu re

Downstream view

( f ) Total- and s tat ic -pressure i n s t r u m e n t a t i o n at

F igure 5. - Concluded.

d i f f use r s ta t ion 5.

26

ln 3

m a = E c . .- a

x F ' a

c-l

I= 0

c L 0 c Ln

.-

._ n

Free-stream Capture mass- Mach number, flow ratio,

MlMo 0 2.58 0.998 0 2.50 .998 0 2.30 .914 A 2.02 .767

. -. . - .. . Region of shock instabi l i ty

1.0

. 9

. 8

. 7 . 8 . 9 1.0 . 8 . 9 1.0 .7 .8 . 9 .6 . 7 . 8

Engine mass flow ratio. m51m0

(a-1) Free-stream (a-2) Free-stream la-3) Free-stream (a-4) Free-stream temperature, temperature, temperature, temperature, 373 K; cowl-lip- 317 K; cowl- l ip- 317 K; cowl- l ip- 317 K, cowl- l ip- position param- position param- position param- position parameter, 26.23"; eter, 26.6"; eter, 23.87"; eter, 20.6"; Reynolds n u m - Reynolds nu g~- Reynolds nu r- Reynolds n u m - ber, 3 . 4 5 ~ 1 0 ~ . ber. 3 .82~10 . ber, 3 .82~10 . k r , 3 . 8 2 ~ 1 0 ~ .

(a) Configuration I IND' ; angle of attack, 0.

Figure 6. - Overal l in le t performance wi thout bypass flow.

27

Angle of attack,

a, deg

Y 1-

Capture mass- flow ratio,

m/mO

0 0 2.6

0.998

0

9

8

7 . 8 .9 1.0

Angle of Capture mass- attack, flow ratio,

a. M l M O deg

0 0 0.993 2.50

P -2.50 _ _ _ _ _ _ Region of shock instabi l i ty

8 . 9 1.0 Engine mass flow ratio, m 5 l m o

(b) Conf igurat ion I; free-stream temperature, 317 K; free-stream Mach n u m - ber, 2.50; Reynolds number, 3 . 8 2 ~ 1 0 ~ ; cowl-l ip-position parameter, 26.6".

(c) Conf igurat ion I; free-stream temperature, 400 K; free-stream Mach n u m - ber, 2.46; Reynolds number, 3 . 2 4 ~ 1 0 ~ ; cowl-l ip-position parameter, 26.6.


28

y\

T

7=-

E

m a c .- a

x m E a

w

Capture mass- Reynolds Free-stream Angle of Free-stream flow ratio, number. Mach number, attack, temperature,

0 0.998 3 . 4 5 ~ 1 0 ~ 2.58 0 373 0 3.45 2.58 4.80 373 A .993 3.24 2.55 0 39 1 0 3.24 2.55 4.80 39 1

____.._.__ Region of shock instability

Tailed symbols denote total-pressure profiles (see fig. 10) Open symbol: denote engine airflow variat ion us ing choked exit p lug Solid symbols denote performance at constant engine corrected airf low

by varying overboard bypass exit area

. 9

. 8

.7 .75 .85 .95 .75 .85 .95 .75 .85 .95

Engine mass-flow ratio. m5/m0 75 .85 .95

(a) Nominal free-stream Mach number, 2.6; cowl-lip-position parameter, 26.23"; ejector area to capture area ratio, 0.0075; engine corrected airflow, 15.89 ki lo- grams per second.

Figure 7. - Overall performance of in le t configuration I I N D wi th bypass flow.

4 29

Capture mass- Reynolds Free-stream Angle of Free-stream flow ratio, number, Mach number, attack, temperature,

a, TO, deg K

0 0.998 3 . 8 2 ~ 1 0 ~ 2.50 0 317 0 3.82 2.50 2.7 317 a .993 3.24 2.47 0 390 0 3.24 2.47 2 .5 390

Im/mo) capt Re MO

_._____ Region of shock instabi l i ty Tailed symbols denote total-pressure profiles (see fig. 11) Open symbols denote engine airf low variat ion us ing choked exit plug Solid symbols denote performance at constant engine corrected airf low by

varying overboard bypass exit area

75 .85 .95 .75 .85 .95 Engine mass-flow ratio, m5lm0

(b) Nominal free-stream Mach number, 2.5; cowl-lip-position parameter. 26.6"; ejector area to capture area rat io, 0.0081; engine corrected airf low, 15.83 ki lo- grams per second.


30

Lr-

Capture mass- Reynolds Free-stream Free-stream flow ratio, number, Mach number temperature,

I m/ mol Re MO capt

Y)

T

2

E

m a

c .- a

o 0.914 3 . 8 2 ~ 1 0 ~ 2.30 317 a .921 3.37 2.34 362

Tailed symbols denote total-pressure profiles (see fig. 12) Open symbols denote engine airf low variat ion us ing choked

Solid symbols denote performance at constant engine exit p lug

corrected airf low by varying bypass exit area

Engine mass-flow ratio. m5/m0

(c) Nominal free-stream Mach number, 2.3; cowl-lip-position parameter, 23.87'; angle of attack, 0"; ejector area to capture area ratio, 0.0076; engine corrected airflow, 16.75 k i l q r a m s per second.

Capture mass- Reynolds Free-stream Free-stream Ejector area flow ratio, number. Mach number, temperature, to capture

TO, area rat io, K

( ml mol capt R e MO

0 0.767 3 . 8 2 ~ 1 0 ~ 2.02 317 0.02% a .762 3.48 1.98 324 .0248

Tailed symbols denote total-pressure profiles (see fig. 13) Open symbols denote engine airf low variat ion us ing choked exit p lug Solid symbols denote constant engine corrected airf low by varying

overboard bypass exit area

5 .6 .7 . 5 Engine mass-flow ra t io m5/m0

(d) Nominal free-stream Mach number, 2.0; cowl-lip-position parameter, 20.6", angle of attack, 0"

Figure 7 - Concluded.

31

Total- Engine pressure mass- recovery, flow

P51P0 ratio, m5'm0

0 0.875 0.936 0 .870 .932 0 .870 .932

.853 .914 a V .847 .909 n .742 .776 A .760 .763

.793 .842 V 7 ,813 .863 D .817 ,905

---- Theoretical Measured

parameter, 26.23"


P51P0 ratio, m5'm0

0 0.926 0.912 0 ,930 .918 0 .931 .921 n .934 .923 V .934 .924 n ,799 .780

.791 .769 A V .779 .743 7 .849 .836 D .873 .a57 a .887 .869 n .902 .887 0 ,912 .897 0 .915 .901 0 .917 .905 b .825 .911

Theoretical Measured

----

(b) Free-stream Mach number, 2.50; free-stream temperature, 317 K; cowl-l ip-position parameter, 26.6".

Figure 8. - In te rna l cowl surface static-pressure distr ibut ions for var ious bypass door settings and engine match corrected airflolu. Conf igurat ion IIND'; angle of attack, 0".

32

0

X

a

cz - d L 3 VI VI al L a. V c m I VI

.-


P5/P0 ratio, m5'm0

0 0.928 0.904 0 ,935 ,905 0 ,937 ,913 n ,919 ,893 v ,913 ,887 n .896 ,868 A .878 .845 V ,845 .812 \ ,829 .793 D ,804 ,772 a ,797 .758 n ,793 .742

---- Theoretical Measured

i.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 Distance from spike t ip, x lRc, dimensionless

a3

L 3 (c) Free-stream Mach number, 2.47; free-stream temperature 390 K; cowl-lip-position parameter, 26.6" VI m a L a

0 0 0 a D n A V P

----

Total- pressure recovery,

'5"O

0.815 .861 .891 ,913 .929 .929 .943 .938 .933


Distance from spike tip, x lRc , dimensionless

id) Free-stream Mach number, 2.30; free-stream temperature, 317 K; cowl-t ip-posit ion parameter, 23.87'.


Engine mass- flow

ratio, m5'm0 0.699

,736 .766 .788 .799 .804 .823 .819 .811

33

a, a,

0- c x L

V I L

t c n

a = 3 a,-

z 2 k z .u k

m z

O E .e z

c c .- m u v) - m - - * E L

c lv E

0

x n

n -...

2- 3 VI VI a, L

V

I m c VI

n .-

5 a, L c VI

a, a, L

0 - c

E 3 VI VI a, L

V

c m c VI

m u 0

0 0

m E

n ._

- - - .- c


8 p51p0

7 0 0.926 6 0 .925

0 .909 5 A .896

V .893 0 .853 4

V .802 T .864


3 A .a0 2

1

0 2.75 3.25 3.75 4.25 4.75 5.25 5.75 6.25 6.75 7.25 7.75 8.25

Distance from spike tip, x/Rc, dimensionless

----

le) Free-stream Mach number, 2.02; free-stream temperature, 317 K; cowl-lip-position parameter, 20.6".


Engine mass- f l w

ratio, m5'm0 0.665

,665 .651 ,639 ,633 .602 .591 .558 .612


P5lPo rat io, m5'm0

0 0.875 0.936 0 .870 .932 0 .870 .932 A .853 .914 v .847 .909

.760 ,763 n A .742 .776 V .793 . a 2 P .813 .863 D .817 .905

Th eoret ica I Measured Extrapolated

---- ---

Distance from spike tip, dRC, dimensionless

(a) Free-stream Mach number, 2.58; free-stream temperature. 373 K; cowl-lip-position parameter, 26.23".

Figure 9. - Centerbody surface static pressure distr ibutions for various bypass door settings and engine match corrected airflow. Configuration I IND' ; angle of attack, 0".

34

16

15

14

13

12

11

10

9

8

- 2 7

" x 6

3 5

2 4

G)- L

VI VI

CI

E m

2 1 m

0 0 0 n v n A V T D

0 0

a

O n n

____

Total- Engine pressure mass- recovery, flw

P5lPo ratio, m5'm0

0.926 0.912 .930 ,918 ,931 ,921 ,934 ,923 ,934 ,924 ,199 ,780 .791 ,769 ,179 ,743 ,849 ,836

,887 ,869 .902 .887 .912 ,897 ,915 ,901 ,917 .905 ,825 .911

.a73 ,857

Theoretical Measured Extrapolated

a, a, (b) Free-stream Mach number, 2.50; free-stream temperature. 317 K; cowl-lip-position parameter. 26.6". L

0 - c G) L

3 VI m G) L a

L 0 0

m E

.- c 0

0 0 A v 0 A V 7 D

a a


P5lPo ratio, m51m0

0.928 0.904 .935 ,905 .931 .913 .919 .893

,896 ,868 .878 ,845 ,845 ,812 ,829 ,793 ,804 .772 .797 ,158 .793 ,742

,913 ,887

---- Theoretical Measured Extrapolated

parameter 26.6".

35

- 0 0

m E

.- c

12 11 Total-

pressure 10 recovery, 9 '5"0

0 0.815 8 7 0 .861

0 ,891 - 3 6 n .913 V .929 h .929

- 5

2 4 A .943 aJ V .93a

n .933 b 3

Theoretical x 2 ---_ Measured Extrapolated 2 t ; o

& 2.25 2.75 3.25 3.75 4.25 4.75 5.25 5.75 6.25 6.75 7.25 7.75 - L 2 VI Tota 1 -

x

d L

VI

U .- c VI

5 1 ---

aa

0 Distance from spike tip, xlRc, dimensionless

c (d) Free-stream Mach number, 2.30; free-stream temperature, 317 K; cowl-l ip-position parameter, 23.87".

3 VI

a L

U

m VI

pressure recovery,

'5 ' '0 8

0 0.926 0 .925 7 0 .909 a .896

6

5 V .893 n .853 A .&to

4

3 V .802 n .864

Theoretical 2

Meas u red Extrapolated

1

0 2.75 3.25 3.75 4.25 4.75 5.25 5.75 6.25 6.75 7.25 7.75 8.25


n .- c c

_--_ ---

(e) Free-stream Mach number, 2.02; free-stream temperature, 317 K; cowl-l ip-position parameter, 20.6'.


Engine mass- flow

ratio, m5'm0 0.699

.736

.766

.788

.799

.804

.823

.819

.871

Engine mass- flow

ratio, m5'm0 0.665

.665

.651

.639

.633

.602 ,591 .558 .612

36

1.0

.8

. 6

. 4

.2

0

i o i a i - Engine Free- Free- ' l o ' 7 pressure mass- stream stream 1 I ' recovery, flow temper- Mach 1- { 1 P5lPo ratio, ature, number,

m5imo To, Mo

LLl 0 0.859 0.917 373 2.58 0 .864 ,914 391 2.55

Solid symbols denote wall static pressure

I- k?; K

Reynolds Distort ion number,

Re

3 . 4 5 ~ 1 0 ~ 0.086 3.24 .OM

.7 . 9 .7 . 9 .7 . 9 . 7 . 9 . 7 . 9 .7 . 9 .7 .9 Local t d a l pressure recovery. Fg h i p 0

Figure 10. - Comparison of compressor-face total-pressure profiles at engine match condit ions wi th and without t u n n e l heat. Configuration IIND'; nominal free-stream Mach number, 2.6; angle of attack, 0"; ejector area to capture area ratio, 0.0075.

.4 Total- Engine Free- Free- Reynolds Distort ion pressure mass- stream stream number, recovery, flow temper- Mach Re

m5lm0 To, Mo K

. 2 P5lPo ratio, ature, number,

0 o 0.920 0.905 317 2.50 3 . 8 2 ~ 1 0 ~ 0.105 o .929 ,894 390 2.47 3.24 ,100

Solid symbols denote wall static pressure 1.0

. 8

. 6

. 4

. 2

0 .8 1.0 . 8 1.0 .8 1.0 .8 1.0 .8 1.0 . 8 1.0 .8 1.0

Local total-pressure recovery, P5 h/PO

(a) Engine match corrected airflow; bypass set at match opening.

Figure 11. - Comparison of compressor-face total-pressure profiles at various operating condit ions w i th and wi thout t unne l heat. Configuration I IND' ; nominal free-stream Mach number, 2.5; angle of attack, 0"; ejector area to capture area ratio, 0.0081.

37

1.0

.8

. 6

. 4

.2

0 .8 1.0

1.0

.8

. 6

. 4

. 2

0 .7 . 9

.8 1.0

_ - - Total- Engine Free- Free- 110 I p ressure mass- stream stream r 1 - I recovery, flow temper- Mach _ - I - J P5lPo ratio, ature, number,

m5Jm0 To, MO K - mz

1 1 1 L ~~ 0 0.9@ 0.833 317 2.50

0 ,951 .833 390 2.47


8 1 . 0 . 8 1 . 0 . 8 1 . 0 . 8 1 . 0 .


Re

3 . 8 2 ~ 1 0 ~ 0.047 3. 24 .071

8 1.0 Local total-pressure recovery, P5,h/P0

(b) 90 Percent of engine match corrected airflow; bypass set at match opening.

~1-1 Total- Engine Free- Free- pressure mass- stream stream r-[!i recovery, flow temper- Mach

P5lPo ratio, ature, number, m5 lm0 Tb Mo

0 0.846 0.913 311 2.50 .847 ,896 390 2.47

Solid svmbols denote wall static pressure

K

Reynolds Distort i o n number,

Re

3 . 8 2 ~ 1 0 ~ 0.120 3.24 .155

.7 . 9 .7 . 9 . 7 . 9 .7 . 9 .7 . 9 Local total-pressure recovery, P5,h/P0

. 7 . 9

(cl 110 Percent of engine match corrected airflow; bypass set at match opening.

Figure 11. - Continued

- I .4, - 1 Total- Engine Free- Free- Reynolds Distort ion Rake8 19 ' i o 1 ' pressure mass- stream stream number,

a n 2 2 +.- h g +?-- '--- ' ' ' recovery, flow temper- Mach Re u l a c -c :: gz L I I c

c Io OdApJ l -1-J 0 0.783 0.914 317 2.50 3 . 8 2 ~ 1 0 ~ 0.234 0 ,785 .899 390 2.47 3.24 .246

. 6 i.0 . 6 1.0


.6 1 . 0 . 6 1 . 0 . 6 1 . 0 . 6 1.0 Local total-pressure recovery, P5,h/P0

. 6 1.0

(d) 119 Percent of engine match corrected airflow; bypass set at match opening.

. 4 r - r - 1 r - - I Total- Engine Free- Free- Reynolds Distort ion / 9 ~ , l o 1 I pressure mass- stream stream number,

I recovery. flow temper- Mach Re . 2

0

1.0

. 8

. 6

.4

. 2

0 . 8 1.0

E

. 8 1.0

P5lPo- ratio, atu're, number, m5lm0 To, Mo

K 0 0.934 0.924 317 2.50 3 . 8 2 ~ 1 0 ~ 0.108

.937 ,913 390 2.47 3.24 .lo5 Solid symbols denote wall static pressure

rtl

. 8 1.0 . 8 1.0 . 8 1.0 . 8 1.0 . 8 1.0 Local total-pressure recovery, P5,h /Po

(e) 101.5 Percent of match pressure recovery set wi th bypass; (w&/d), = match.


39

Total- Engine Free- Free-

recovery, flow temper- Mach


Re

3 . 8 2 ~ 1 0 ~ 0.088 3.24 . 106 .845 ,812 390 2.47


Local total-pressure recovery, P5,h/P0

( f ) 91.6 Percent of match pressure recovery set wi th bypass, ( w 4 / 6 ) = match. 5

7 . 9

, rr- , Total- Engine Free- Free- Reynolds Distort ion

P5/Po rat io, a ture. number,

pressure mass- stream stream number, recovery, flow temper- Mach Re

m5/m0 '$8 MO [fi 1'3 o 0.791 0.769 317 2.50 3 . 8 2 ~ 1 0 ~ 0.194 .193 .804 ,772 390 2.47 3.24

S

. 6 1 . 0 . 6 1 . 0 . 6 1 . 0 . 6 1 . 0 . 6 1 . 0 . 6 1 . 0 . 6 1.0 Local total-pressure recovery, P5,h /Po

(g) 86.3 Percent af match pressure recovery set wi th bypass, (~%6/6)~ = match.


40

. 4 - - iora l - Engine Free- Free- Reynolds Distort ion l R d k e 8 9 '10 1 I pressure mass- stream stream number,

I I t 434

I 1 recovery, flow temper- Mach Re

m5/m0 TO*

- - I

0 ~ - -1 1 I 1 0 0.933 0.811 3y7 2.30 3 . 8 2 ~ 1 0 ~ 0.093 0 .918 ,839 362 2.34 3.37 ,106


. 8 1.0 .8 1.0 . 8 1.0 . 8 1.0 L a a l tda l -pressure ieco iery F5,h iF0

Figure 12. - Comparison of compressor-face total-pressure profiles at engine match condit ions wi th and wi thout t u n n e l heat. Configuration IIND', nominal free-stream Mach number. 2.3, angle of attack, O", ejector area to capture area ratio, 0.0076,

. 4 r Total- Engine Free- Free- Reynolds Distort ion

5 2 P5lPo ratio, ature, number,

0 -11 0 0.908 0.672 317 2.02 3 . 8 2 ~ 1 0 ~ 0.089 .927 .665 324 1.98 3.48 .085

pressure mass- stream stream number, recovery, flow temper- Mach Re

Rake8 '

I 1 m5lm0 TO. Mo * K


. 6 1. Local total-pressure recovery, P5,h /Po

Figure 13. - Comparison of compressor-face total-pressure profi les at engine match condit ions wi th t u n n e l heat. Configuration IIND'; nominal free-stream Mach number, 2.0; angle of attack, 0".

Ejector area to capture area rat io,

Ae j ' Ac

0.0236 .02@

and wi thout

41

Reynolds Free- Free- number, stream stream

R e Mach temper- number, ature,

.10

.08

Mo To. K

0 3 . 4 5 ~ 1 0 ~ 2.58 373 0 0 3.45 2.58 373

A 3.24 2.55 391 0 3.24 2.54 390

E .M

.06

E

E

0-

- Region of shock instabi l i ty ------ .- - m

L .02 - s (a) Nominal free-stream Mach number, 2.6; cowl- l ip-posit ion parameter, 26.23". - VI VI

Reynolds Free- Free- a, a, number, stream stream R Re Mach temper-

number, ature,

.12 n -

.10 + 0

I- Mo T;.

.08 0 3 . 8 2 ~ 1 0 ~ 2.50 317 0 3.82 2.50 317 A 3.24 2.47 390 0 3.24 2.47 390 .06

- - - - - - - Region of shock instabi l i ty .04

.70 .74 ,78 .82 .86 .90 .94 .98 Total-pressure recovery, P5lPO

(b) Nominal free-stream Mach number, 2.5; cowl-lip-position parameter. 26.6".

Figure 14. - Variat ion of bleed flow of in let configuration IIND' with bypass flow.

Angle of

attack. a.

de9

0 4.8 0 4.8

Angle of

attack, a,

de3

0 2.7 0 2.5

Reynolds Free- Free- number, stream stream

Re Mach temper- .08 number, ature,

. 10

Mo Tot K

0 3 . 8 2 ~ 1 0 ~ 2.30 317 fl 3.37 2.34 362

.06 - - r)

E a- .M

s f .02

._ I

m L

m g E

Reynolds Free- Free- : .08

number, stream stream 0 Re Mach temper-

.06 number, ature.

(c) Nominal free-stream Mach number, 2.3; cowl-lip-position parameter, 23.87";

m - n - m .,-.

Mo T;.

0 3 . 8 2 ~ 1 0 ~ 2.02 317 fl 3.48 1.98 324

.M

.02 .78 .82 .86 .90 .94 .98

Total-pressure recovery, P51Po

(d) Nominal free-stream Mach number, 2.0; cowl-lip-position parameter, 20.6"; angle of attack, 0".


4 3

v) v) (a) Nominal free-stream Mach number, 2.6; free- stream temperature, 373 to 391 K. m a.

)r a

I 0 c V G)

W .-

1.0

.oa

.06

.04

.02

0 .01 .02 .03 .M .05 .06

(b) Nominal free-stream Mach number, 2.5; free stream temperature, 317 to 390 K.

0 .01 .02 .03 .M .05 .06 Ejector bypass area ratio, Ae,/Ac

(c) Nominal free-stream Mach number, 2.3; free- ( d ) Nominal free-stream Mach number, 2.0; free- stream temperature, 317 to 362 K. stream temperature, 317 to 324 K.

Figure 15. - Ejector bypass operating characterist ics.

Cow I-lip-position parameter,

deg 0 26.23 0 24.8 0 22.9 I 2i. 43 a 21.33

9 1.

Open symbols denote unstarted condit ion Half open symbols denote started condit ion Solid symbols denote unstable (buzz) condit ion Tailed symbols denote m i n i m u m stable o r peak

condit ion Double tailed symbols denote overboard bypass

set for match point operation + Average unstable condit ions immediately

after unstart; engine corrected airf low, 15.89 kglsec

Engine plus overboard bypass mass-flow rat io, (m5 + mby)/mo

(a) Free-stream Mach number, 2.58; free-stream temperature, 373 K; Reynolds number, 3 . 4 5 ~ 1 0 ~ ; ejector area to i n le t capture area rat io, 0.0075.

Figure 16. - Overall in le t performance du r ing restart sequence of configuration I IND' ; angle of attack, 0".

45

--," m a -2-

E c .- a

x m E a

L=_I

c 0

0 VI

0

.- c

c .-

0

L n

a

a \

r; al > 0 V al L

a, L

3 VI VI al L a - m 0 + I

Cowl-lip-position parameter,

012 de9

0 26.6 0 24.8 0 22.87 a 21. 7 n 21.15 n 21.02 0 24.13 n 25.55

Open symbols denote unstarted condit ion Half open symbols denote started condit ion Solid symbols denote unstable condit ion Tailed symbols denote m i n i m u m stable

o r peak condit ion Double tai led symbols denote overboard

bypass set for match point operation 4 Average unstable condit ions immediately

after unstart; engine corrected airflow, 15.83 kglsec

Engine plus overboard bypass mass-flow rat io, (m5 + mbyl/mo

(b) Free-stream Mach number, 2.50; free-stream temperature, 317 K; Reynolds number, 3 . 8 2 ~ 1 0 ~ ; ejector area to in let capture area rat io, 0.0081.


46

Cowl-lip-position parameter,

OZf de9

26.6 24.77 22.87 21. 7 21.37

0

0

0 a n a 21.17


o r peak condit ion Double tailed symbols denote overboard

bypass set for match point operation Triple tai led symbols denote overboard bypass

closed a Stable conditions immediately after

unstart; engine corrected airflow, 15.83 kglsec

. . . 5 .6 . 7 . 8 . 9 1.0

Engine plus overboard bypass mass-flow ratio, (m5 + mby)/mo

(c) Free-stream Mach number, 2.47; free-stream temperature, 390 K; Reynolds


number, 3 . 2 4 ~ 1 0 ~ : ejector area to in let capture area rat io, 0.0081.

47

0 a 1

a"

5 F E E

E

0 U

3 v) v)

n.

m 0 c

- -

1.0 -

&-- 4 -& Tailed ivmbols denote m i n i m u m stable

0 21.5 A 20.43 a 20.25

Open symbols denote unstarted condit ion Half open symbols denote started condit ion Solid svmbols denote unstable condit ion

o r peak condit ion , Double tai led symbols denote overboard -- bypass set for match point operation * Average unstable condit ions immediately

after unstar t , engine corrected airf low, 16.75 kglsec

. 6 . 7 .8 . 9 Engine p lus overboard bypass

mass-flow ratio, (m5t mby)/mO

Reynolds number, 3.82x106, ejector area to in le t capture area ratio, 0.0076.

Id) Free-stream Mach number, 2.30; free-stream temperature, 317 K:

Cowl - I ip-pos it ion parameter,

81. de9

0 20.6 0 19.95 0 19.4 a 19.38


o r peak condit ion Double tailed symbols denote overboard

bypass set for match point operation Triple tailed symbols denote overboard bypass

closed o Stable condit ions immediately after

unstar t ; engine corrected airf low, 17.78 kglsec

.6 .7 .8 Engine plus overboard

rat io, ( m 5 + mb Y pi".o bypass mass-fl

(e) Free-stream Mach number, 2.02; free-stream temperature, 317 K; Reynolds number, 3 . 8 2 ~ 1 0 ~ ; ejector area to i n le t capture area rat io, 0.0236.


48

Total- Engine plus pressure overboard bypass recovery, mass-flow r t io,

P5/Po (m5 + mby)/'mo o 0.485 0.484 Unstable 0 .488 .507 Unstable 0 .482 .544 Stable

!, 317 K. Total-

pressure recovery,

'5"O 0.563

.541

.520

Engine plus overboard bypass mass-flow r t io, ( m 5 t mby)/'mO

.659

.659

0.661 Min imum stable

317 K. Total-

pressure recovery,

'5"O 0.617

,619 .598

Engine plus overboard bypass mass-flow r t io, ( m 5 t mby)/'mO

.755

.759

0.740 M in imum stable

(b) Cowl-lip-position parameter, 24.8"; unstarted; free-stream tempera . ure, 10

0 0 0 6

4 I I I I I I I I Y I I I I I I I B y p a s s 1 1 I

'1.5 I 310 I 3!5 I 4.0 I 4.15 ' ! O ' 5!5 ' 610 ' d;l ' 7:O '7!5 ' 8!0 opening


(c) Cowl-lip-position parameter, 22.87"; unstarted; free-stream temperature, 317 K.

Figure 17. - In ternal cowl-surface static-pressure distr ibutions du r ing inlet restart sequence. Configuration I IND' ; nominal free-stream Mach number, 2.5; Reynolds number, 3 .82~10~ ; angle of attack, 0".

49

10

8

6

4

I 2

2 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 1.0 7.5 8.0

(d) Cowl-lip-position parameter, 22.87"; unstarted; free-stream temper


'5"O 0 0.630 0 .606 0 .586

.ahre, 390 K. Total-

pressure recovery,

'5lP0 0 0.680 0 ,680 0 .670 A .657

Engine plus overboard bypass mass-flow r tio, "ll5' mby)fmO

.755 ,766

0.733 M i n i m u m

Engine plus overboard bypass

0.732 M i n i m u m .755 .116 ,194

stable

stable

5 c 2.75 3.25 3.15 4.25 4.75 5.25 5.75 6.25 6.75 7.25 1.15 8.25

.- 0

0

( e ) Cowl-lip-position oarameter. 21.15'; unstarted; free-stream temperature, 317 K. 12 E

10

8

6

4

2 2.75 3.25 3.75 4.25 4.75 5.25 5.75 6.25 6.75 7.25 7.75 8.25


i a l - Engine plus pressure overboard bypass recovery, mass-flow r tio,

P5/Po (m5 + mby)finO 0 0.730 0.804 Peak 0 .723 ,810 0 ,706 .821 A .686 .826

( f ) Cowl-lip-position parameter, 21.02"; just started; free-stream temperature, 311 K.


50


F

E

c v1

W - 0 c U .-

Total- Engine plus pressure overboard bypass recovery, mass-flow ratio,

~ 5 1 ~ 0 (m5 + mby)/mo 0 0.863 0.891 Peak 0 .857 .891 0 . a 2 ,905 A .819 ,919

317 K.

Total- Engine plus pressure overboard bypass recovery, mass-flow ratio,

~ 5 1 ~ 0 (m5 + mby)/mo 0 0.485 0.484 Unstable 0 .488 .507 Unstable 0 .482 .544 Stable

(a) Cowl-lip-position parameter, 26.6'; unstarted; free-stream temperature, 317 K.

10

8

6

4

2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 Distance from spike tip, xlRc, dimensionless


'5lP0 0 0.563 0 .541 0 .520

Engine plus overboard bypass

0.661 M i n i m u m stable .659 .659

(b) Cowl-lip-position parameter, 24.8"; unstarted: free-stream temperature, 317 K.

Figure 18. - Centerbody surface static-pressure distr ibutions du r ing inlet restart sequence. Configuration IIND'; nominal free-stream Mach number, 2.5; Reynolds number, 3.82~106; angle of attack, 0".

51


'5"O 0 0.617 0 .619 0 .598

0 (c) Cowl-lip-position parameter, 22.87"; unstarted; free-stream temperature, 317 K.


p5"0 0.630 .606 .586

n

!, 390 K.


'5"O 0.680 .680 .670 .657

(e) Cowl-lip-position parameter, 21.15"; unstarted; free-stream temperature, 317 K.


Engine plus overboard bypass mass-flow r tio, 'm5t mbyl!mO

.755

.759

0.740 M i n i m u m

Engine plus overboard bypass mass-flow r tio, (m5 + mby)!mO

.755

.766

0.733 M i n i m u m

Engine plus overboard bypass mass-flow r tio, (m5t mby)!mO

.755

.776

.794

0.732 M i n i m u m

stable

stable

stable

52

12

10

8

6

pressure overboard bypass 8 4 .... x CL

d 3 2 L

VI VI 0) L

u n .- c , o c VI

I f ) Cowl-lip-position parameter, 21.02"; just started; free-stream temperature, 317 K. ". 5

(gl Cowl-lip-position parameter, 24.13"; started; free-stream temperature, 317 K ,


53

. 4 r ~' 7

l Rake 8 , Total- Engine plus pressure overboard bypass recovery, mass-flow r tio,

p g l p o (m5 + m b j m O 0 0.485 0.484 Unstable 0 ,488 .507 Unstable 0 .482 .544 Stable


(a) Cowl-lip-position parameter, 26.6"; unstarted; free-stream temperature, 317 K.

4- 1 Rake8 1

. 2

0

.4 . 6

r - - -

Total- Engine plus pressure overboard bypass

110 1

0 0.563 0.661 M i n i m u m stable .659

0 .520 .659 Solid symbols denote wall static pressure

. 4 . 6 . 4 . 6 . 4 . 6 4 . 6 . 4 . 6 . 4 . 6 Local total-pressure recovery, P5,h/P0

(b l Cowl-lip-position parameter, 24.8"; unstarted; free-stream temperature, 317 K.

Figure 19. - Comparison of compressor face total-pressure profiles du r ing in le t restart sequence. Configuration I I N D ' ; nominal free-stream Mach number, 2.50: Reynolds number, 332x106; angle of attack, 0".

54

I 4---- Total- Engine plus

;:e;;~;e O V ~ I b i d uypass recovery, mass-flow rgt io,

P ~ / P O (m5 + mby)/mo 0 0.617 0.740 M i n i m u m stable

.619 .755 0 .598 .759


(c ) Cowl-lip-position parameter, 22.87'; unstarted; free-stream temperature, 317 K.

1.0

. 8

. 6

. 4

.2

0 . 5 .7

Total- Engine plus pressure overboard bypass recovery, mass-flow r tio,

P5/Po (m5 + mbJmO 0 0.630 0.733 M i n i m u m stable 0 .606 .755 o .586 .766


5 .7 . 5 . 7 . Local total-pressure recovery, P5,h /Po

(d) Cowl-lip-position parameter, 22.87-; unstarted; free-stream temperature, 390 K.


55

Total- Engine p lus 19- CT-- pressure overboard bypass

recovery, mass-flow r tio, p 5 l p o ( m 5 + mb$fmo

J ‘-1- 2 L--]

0 0.680 0.732 M i n i m u m stable 0 .680 .755

- J L L A 0 .670 .776 A .657 .794

Solid symbols denote wall static pressure 1 0

.8

6

4

. 2

0 55 75 55 .75 . 55 75 . 55 75 . 55 75 55 .75 .55 .75

Local total-pressure recovery, P5,h/P0

(e) Cowl-lip-position parameter, 21.15”; unstarted; free-stream temperature, 317 K.

Total- Engine p lus p r e s s w e overboard bypass recovery, mass-flow r tio,

P5/Po (m5 + mby)fmO 0 0.730 0.804 Peak

.723 .810 0 .706 .821

.686 .826 Solid symbols denote wall static pressure

1.0

\ . 8

.6

.4

. 2

0 . 6 . 8

( f ) Cowl-

. 6 .8 . 6 . 8 .6 . 8 . 6 . 8 Local total-pressure recovery, P5 h /PO

-lip-position parameter 21.02; just started; free-stream


.6 .8

temperatut

.6 .8

.e, 311 K.

56

Total- Engine plus pressure overboard bypass recovery, mass-flow r tio,

P5/Po (m5 + mbJ'mo 0 0.863 0.891 Peak 0 .857 .891 0 .&I2 .905 A .819 .919

Solid symbols denote wall s!a!ic pressure

. 7 . 9 .

lg) Cowl

7 .9 . 7 . 9 . 7 .9 . 7 . 9 . 7 . 9 . 7 . 9

-l ip-position parameter, 24.13"; started; free-stream temperature, 317 K.

Local total-pressure recovery, P5 hlPO

Figure 19. -Concluded.

57

Config- Total- Engine mass- Distort ion urat ion pressure flow ratio,

recovery. m5lm0 '5l'O

A IWD 0.905 0.929 0.087 IWC .912 .922 .120

Solid symbols denote wall static pressure r

m 3 . 4

2 3 5 .- = ? E ...- . 2 0 53

" t: s -- o m

x - 3 E =

- ._

5 0

Config- Total- Engine mass- Distort ion urat ion pressure flow ratio,

recovery, m5lm0 '51p0

0 IND 0.900 0.931 0.126 0 I I N C ,901 .923 .108


1.0

. 8

. 6

. 4

. 2

n .8 .9 1.0 . 8 . 9 1.0 . 8 . 9 1.0

(a1 Configurations IWC and IWD.

Local total-pressure recovery, P5,h I Po

(b) Configurations I I N C and IND.

Figure 20. - Effect of vortex generator orientation on engine-face total-pressure measurements. Free-stream Mach number, 2.50; angle of attack, 0"; no bypass; free-stream temperature, 317 K; Reynolds number, 3 . 8 2 ~ 1 0 ~ .

. 8 . 9 1.0 . 8 . 9 1.0 . 8 . 9 1.0

58

o n o n 0 0 o n 0 0 o n Configuration I I IND ' I I INO' I I IND' I I I N D ' I I I N D ' I I I N D ' Bypass mass-flow 0 0 0.041 0.041 0.079 0.077 0.123 0.128 0.123 0.247 0.123 0.445

Total-pressure 0.924 0.928 0.925 0.944 0.937 0.950 0.940 0.955 0.940 0.950 0.940 0.939

Distort ion 0.115 0.079 0.106 0.055 0.099 0.063 0.101 0.058 0.101 0.046 0.101 0.052

ratio, mby/mO

recovery, P5lPo


1.0

.8

. 6

. 4

.2

0 8 1.0 .8 1.0 .8 1.0 . 8 1.0 . 8 1.0 .8 1.0

Local total-pressure recovery, P5,hlPO

Figure 21. - Effect of centerbody vortex generators on engine face total-pressure profiles of rake 7 with bypass operation. Peak in le t operation; free-stream Mach number, 2.50; free-stream temperature, 317 K; angle of attack, 0'; Reynolds number, 3 .82~10~ .

59

0

In a

a --.

>; W

0 u L a, L 3 VI VI

a m

E - - - W U m

L 0 VI m m L CL

-

E V

Free- Reynolds Cowl-lip- Free- stream number, position stream Mach Re parameter, temper-

number, O z , ature, deg T o

K

0 2.58 3 . 4 5 ~ 1 0 ~ 26.23 373 0 2.50 3.82 26.6 317 0 2.30 3.82 23.87 317 A 2.02 3.82 20.6 317

MO

Solid symbols denote overboard bypass f u l l y open .98

.96

.94

.92

.90

.88

.86

I n

T m a g"; . ._ ._

c L E o a V I , c .- n X . z a

L=J

16

12

08

04 0 .1 . 2 . 3 .4 .5 . 6 .7 .8

Bypass mass-flow ratio, mb Y O I m

Figure 22. - Effect of bypass flow on peak i n l e t performance. Configuration I IND'; angle of attack, 0".

60

Bypass Operating condit ion

0 No Peak 0 Yes Peak 0 Yes Match

---- Theoretical throat recovery --- Maximum throat-exit mass-flow

ratio, (m/mo)capt - (mbl/mo)

Pla in symbols denote configuration I I N D ' Tailed symbols denote configuration I Open symbols denote n o t u n n e l heat Solid symbols denote t u n n e l heat

1.0

.8

. 6

I. 0

. 9

.8

20

10

0 1.9 2.0 2 .1 2.2 2.3 2.4 2.5 2.6

Free-stream Mach number, Mo

Figure 23. - Effect of Mach number on in le t performance parameters. Angle of attack, 0".

61

I Free- Reynolds

I stream number, Mach Re

number,

- In

r-7 m a 6- 21 I o c L E o a E l

.- .- c

MO ~~ I z I 2.46 1.24

i - 1 0 2.50 3 . 8 2 ~ 1 0 ~ x . 1 -

a 0

~ 1 J 0

- - I J

I T T T I ,7 A L 1 -

. 8 . 9 1.0 Engine mass-flow

ratio, m5/m0

Figure 24. - Effect of Reynold's number on in le t performance wi thout bypass. Cen- terbody vortex generators installed, con- f igurat ion I I N D ' ; free-stream temperature, 317 K. angle of attack, 0"; cowl- lip-position parameter, 26.6".

62

I I 1.0

Free- Reynolds stream number, Mach Re

nu m ber , . 9 - s - a" MO

.F . 8 + 3.0 3 . 6 ~ 1 0 ~ (ref. 6) < E 0 2.50 3.82 +-. 0 2.02 3.82

0 2.30 3.82 ;; . 7 2

w . 6

L m

Extrapolated Cur ren t data

-___ I 0

V L

L

Tailed symbol denotes data predicted by ._ 3

.o .5

method based on ref. 6 u a, L

# m L

m a, L

2 . 4 m c L a

c c - . 3

. 2 1.0 1.4 1.8 2.2 2.6 3.0

Local cowl Mach number at restart, McOwI

Figure 25. - Comparison of actual to predicted restart area ratio.

NASA-Langley, 1969 - 1 E-4642 63

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