NASA Technical Memorandum 89044 

Measured Pressure Distributions Inside 

Nonaxisymmetric Nozzles With 

Partially Deployed Thrust Reversers 

Robert S. Green and George T. Carson, Jr. 

MARCH 1987 
NASA Technical Memorandum 89044 

Measured Pressure Distributions Inside 

Nonaxisymmetric Nozzles With 

Partially Deployed Thrust Reversers 

Robert S. Green 

Wright-Patterson Air Force Base 

Dayto#, Ohio 

George T. Carson, Jr. 

Langley Research Center 

Hampton, Virginia 

National Aeronautics 

and Space Administration 

Scientific and Technical 

Information Branch 

1987 
Summary 

An investigation was conducted in the Langley 

16-Foot Transonic Tunnel at static conditions to 

measure the pressure distributions inside a nonaxisymmetric nozzle with simultaneous partial thrust 

reversing (50-percent deployment) and thrust vectoring of the primary (forward-thrust) nozzle flow. 

Geometric forward-thrust-vector angles of 0" and 15" 

were tested. These nozzles were installed on a 

model of a generic twin-engine fighter aircraft. Highpressure air was used to simulate jet exhaust. Test 

data were obtained at static conditions while nozzle 

pressure ratio was varied from 2.0 to 4.0. 

The three-dimensional nature of internal exhaust 

flow in a nonaxisymmetric nozzle was indicated 

by variations in pressure between the nozzle flaps 

and the sidewall. Additionally, a complicated, asymmetric exhaust flow pattern was evident in the 

primary-flow exhaust duct of the 15" vectored nozzle. 

Introduction 

Mission requirements for the next generation 

of fighter aircraft may necessitate a vehicle with 

high maneuverability and short take-off and landing 

(STOL) capability. Such an aircraft will probably 

be equipped with multifunction nozzles which 

wiii have thrust-reversing and thrust-vectoring 

capabilities. Both axisymmetric and nonaxisymmetric nozzle concepts have been studied. The 

nonaxisymmetric, thrust-reversing, thrust-vectoring 

nozzle offers the possible advantage of a lower 

installed weight penalty when compared with an 

advanced axisymmetric nozzle having the same 

functions (refs. 1 to 3). 

In addition to an experimental data base which 

provides propulsive performance characteristics, 

studies have provided details of the nozzle internal 

flow fields (refs. 4 to 6). Still, there is much to be 

learned about the internal flow of nonaxisymmetric 

nozzles. A knowledge of the internal flow phenomena 

is needed for improved nozzle designs. Furthermore, 

the development of accurate numerical procedures 

which predict the flow inside a nonaxisymmetric 

nozzle requires an understanding of the physical 

characteristics of the flow field. 

The purpose of this investigation is, therefore, 

to provide detailed static-pressure measurements of 

the flow inside two nonaxisymmetric nozzle configurations with partially deployed thrust reversers. 

The two configurations were identical except the primary nozzle flow was unvectored for one configuration and was pitch vectored 15" for the other. 

Performance data for these nozzles are presented in 

reference 7. The investigation was conducted in the 

T,angley l6-Foot Transonic Tunel (ref. 8) at static 

conditions (M = 0) while nozzle pressure ratio was 

varied from 2.0 to 4.0. 

Symbols 

M 

NPR 

P 

Pa 

Pti 

W 

Wt 

X 

XI 

xt 

Y 

rl 

Vt 

Mach number 

nozzle pressure ratio, pt,j/p, 

local static pressure, psia 

ambient pressure, psia 

jet total pressure, psia 

vertical coordinate, in. (see fig. 5(a)) 

vertical distance from nozzle centerline 

to upper nozzle wall at nozzle 

entrance, 1.050 in. (see fig. 5(a)) 

axial coordinate, in. (see fig. 5(a)) 

surface coordinate starting at internal 

surface break points [see fig. 5(c)) 

axial distance from nozzle entrance to 

primary nozzle throat, 1.595 in. 

(see fig. 5(a)) 

lateral coordinate perpendicular to nozzle 

centerline, in. (see fig. 5(c)) 

perpendicular to line of pressure orifices 

in coordinate system obtained by 15" 

downward rotation of nozzle centerline, 

in. (see fig. 5(b)) 

perpendicular distance in coordinate 

system measured from 15" vectored 

nozzle centerline to nozzle flap, 0.221 in. 

(see fig. 5(h)) 

Abbreviations: 

BL buttock line, in. 

FS fuselage station, in. 

NS nozzle station, in. 

WL waterline, in. 

Apparatus and Procedure 

Wind Tunnel 

This investigation was conducted in the Lzng!ey 

l6-Foot Transonic Tunnel. The free-stream Mach 

number was zero and the wind tunnel was used only 

as a static test facility. 

Twin-Jet-Propulsion Simulation System 

The two nozzle configurations tested were installed on a portion of a wing-tip-supported model 
of a generic twin-engine fighter aircraft. A sketch 

of the aircraft model (without empennage surfaces) 

and the support system is shown in figure 1. A photograph of the test setup is shown in figure 2. The 

external forebody shell, the centerbody top shell, and 

the afterbody shell were removed for this investigation since external flow was not present. A more 

complete description of the wing-tip model support 

system is contained in references 7 and 8. 

For jet-propulsion simulation, the Langley 

16-Foot Transonic Tunnel is equipped with an external high-pressure air system which provides a 

continuous flow of clean, dry air at a nominal 

controlled temperature of 540'R ahead of the nozzle throat. Two remotely operated flow-regulating 

valves are used to balance the jet total pressure in 

the left and right nozzles. The airflow path can be 

traced from the sketches in figure 1. Compressed 

air flows through the two wing-tip support booms, 

then through each wing panel into two high-pressure 

plenums on each side of the fuselage. The air is 

then discharged, perpendicular to the model axis, 

through eight equally spaced sonic nozzles in each 

high-pressure plenum. Next, the airflow passes successively through a tailpipe, a circular-to-rectangular 

transition section, a choke plate, and an instrumentation section, and then expands through the nozzle 

configuration being tested. 

Nozzle Models 

The two nozzle configurations used in the current investigation were part of a test matrix of nozzle 

configurations previously evaluated for aeropropulsive performance (ref. 7). These configurations were 

designed to simulate various percentages of thrustreverser deployment of twin dry-power, nonaxisymmetric, convergent-divergent nozzles having a design 

NPR of approximately 3.5. One of the nozzle design values held constant during this investigation 

was total nozzle throat area; that is, the sum of the 

primary-flow throat area and the reverser-port throat 

area was held constant. (Total nozzle throat area was 

independent of reverser deployment .) In a full-scale 

operational nozzle of the current design, thrust reversal is achieved by rotation of the top and bottom convergent and divergent flaps such that primary flow is 

gradually blocked while the forward-facing reverser 

ports are opened. The convergent and divergent flaps 

are independently actuated such that the divergent 

flaps can be rotated for thrust vectoring of the remaining primary nozzle flow (for reversers deployed 

less than 100 percent). 

In the current investigation, two nozzle configurations simulating 50-percent reverser deployment were 

tested. Reverser deployment of 50 percent means 

that half of the original total throat area is diverted 

to the reverser ports and the remaining half is used 

for the primary (forward-thrust) exhaust flow. One 

nozzle configuration had unvectored primary flow 

(thrust-vector angle equal to O'), and the other configuration had vectored primary flow (thrust-vector 

angle equal to 15O). Sketches of these two configurations are shown in figure 3. Photographs showing 

the nozzle configuration with the reverser deployed 

50 percent and primary flow vectored 15' are shown 

in figure 4. 

Instrumentation 

Jet stagnation conditions were measured in an instrumentation section upstream of each nozzle. (See 

fig. 1.) A four-probe rake was used to measure jet total pressures, and the average of these measurements 

was used in the computation of the nozzle pressure 

ratio NPR and the nozzle static-pressure ratio p/pt,j. 

Jet total temperature was measured with a single 

thermocouple probe located in each instrumentation 

section. Static-pressure orifice locations on the internal surfaces of the nozzle flaps and sidewall are shown 

in the photographs of figure 4 and in the sketches of 

figure 5. Coordinates of the nozzle static-pressure 

orifice locations are given in table I. Fifty-six staticpressure orifices were located on the internal sidewall 

surface of the 0' vectored nozzle (see fig. 5(a)), and 

55 static-pressure orifices were located on the internal sidewall surface of the 15' vectored nozzle (see 

fig. 5(b)). The nozzle upper flaps for the Oo vectored (fig. 5(c)) and 15' vectored (fig. 5(d)) configurations each had 15 static-pressure orifices. Nine 

static-pressure orifices were located on the lower divergent flap of the 15' vectored nozzle (fig. 5(e)). As 

indicated in figure 5 and table I, some of the orifices on the upper flap were positioned slightly off of 

the flap centerline, as necessitated by installation of 

pressure instrumentation. Data obtained from these 

orifices are presented undistinguished from the data 

from the flap centerline since the flow is assumed 

to be essentially two dimensional near the nozzle 

centerline. 

Tests and Data Reduction 

This investigation was conducted in the Langley 

16-Foot Transonic Tunnel at static conditions. Nozzle pressure ratio was varied from 2.0 to approximately 4.0. Test data were recorded on magnetic 

tape with the Langley 16-Foot Transonic Tunnel data 

acquisition system. At each test point, 50 samples of 

data were taken over a 5-sec period. The average 

value of these samples was used for computation. 

2 
Presentation of Results 

Nozzle internal static-pressure distributions are 

presented in figure 6 for the 0" vectored nozzle and 

in figure 7 for the 15' vectored nozzle. Pressure 

data are shown as the ratio of local static pressure 

to jet total pressure p/pt,j as a function of the nozzle 

downstream distance nondimensionalized by the distance from the entrance of the nozzle to the primary 

nozzle throat x/q. (See fig. 5(a).) In figure 6(c), 

the ratio w/wt represents the nondimensionalized 

sidewall height. (See fig. 5(a).) In figure 7(d), 

the ratio q/qt represents the nondimensionalized 

primary-exhaust-duct height. (See fig. 5(b).) Staticpressure-ratio data are presented in table 11. For the 

0" vectored nozzle, these data are given at the locations defined by z/zt and w/wt. For the 15O vectored 

nozzle, these data are given at the locations defined 

Figure 8 presents sketches illustrating some of the 

internal flow phenomena which are indicated by the 

static-pressure distributions. No attempt was made 

to scale these phenomena (location or magnitude) on 

these sketches. 

by x/xt, WIWtl and dvt. 

Discussion 

0" Vectored Nozzle 

Figure 6 presents internal static-pressure data for 

the 0" vectored primary-flow nozzle configuration. 

Figure 8(a) presents a sketch of the hypothesized internal flow phenomena based on these data. The 

upper flap centerline pressures on the 0" vectored 

nozzle (fig. 6(a)) show little variation with NPR. 

However, the sidewall centerline (w/wt = 0) pressures (fig. 6(b)) are slightly dependent on NPR for 

NPR < 3.02. These observations indicate the nozzle 

exhaust flow has certain three-dimensional qualities. 

Since the primary flow exhausts through a constantarea duct, a second choke point may exist for the 

primary flow at the exit because of boundary-layer 

growth in the duct. In this case, the primary-flow 

constant-area duct acts as a supersonic diffuser, and 

this would account for both the relative independence of flap static pressure with NPR and the rise 

in flap static pressure as the duct exit is approached 

(p/pt,j = 0.528 for M = 1.0). In any event, the design NPR for a nozzle having equal throat and exit 

areas is 1.89, and all the data from the current investigation were obtained at underexpanded nozzle 

operating conditions (NPR > 1.89). This fact by itself would cause the nozzle flap static pressures to be 

independent of NPR (ref. 9). 

A stagnation region (p % pt .) on the primaryflow convergent surface is shown in figure 6(a) from '? 

x/xt 2 0.75 to x/xt % 0.82. A strong expansion occurs on either side of the stagnation region as the 

flow approaches the reverser-port throat (decreasing x/q from stagnation point) and the throat of 

the primary-flow constant-area duct (increasing x/xt 

from stagnation point). As indicated by the staticpressure distributions of figure 6(a), expansion of the 

primary exhaust flow continues past the constantarea duct entrance (x/q = 1.00); the sonic line 

(p/ptj = 0.528 for M = 1.0) appears to be located at 

x/xt x 1.07. A separation bubble may be associated 

with the strong expansion around the sharp corner 

of the constant-area duct entrance. A short distance 

downstream of the sonic line (possibly at the separation reattachment point), a shock or strong compression results in a rapid increase in static pressure, 

which is followed by additional supersonic flow expansion. For z/q > 1.50, there is a gradual rise in 

static pressure, which is characteristic of supersonic 

flow in a convergent duct. Boundary-layer growth in 

the constant-area duct would account for this effect. 

As noted previously, the constant-area duct may be 

acting as a supersonic diffuser and a second sonic 

line may exist at the duct exit. Figure 8(a) presents 

a sketch illustrating some of the flow phenomena on 

the upper flap discussed above. It is assumed that, 

because of model symmetry, flow on the lower flap is 

identical to that on the upper flap. 

Static-pressure distributions along the sidewall 

centerline (w/wt = 0; see fig. 6(b)), unlike those 

on the upper flap, show no evidence of a shock or 

strong compression. (See fig. 8.) Examination of 

figure 6(c) shows that the sidewall static-pressure 

distribution varies as flow on the sidewall approaches 

the upper wall of the constant-area duct. At w/wt = 

0.095 (upper wall of constant-area duct located at 

w/wt = 0.210), a weak shock may be present near 

x/xt x 1.40. The above results again illustrate threedimensional flow effects in a nonaxisymmetric (twodimensional) nozzle. The sidewall static-pressureratio contours shown in figure 6(d) illustrate the 

stagnation region on the primary-flow convergent 

surface and the rapid expansion at the corner of 

the entrance to the primary-flow constant-area duct 

discussed previously. 

15" Vectored Nozzle 

Figure 7 presents internal static-pressure data for 

the nozzle configuration with the primary-flow pitch 

thrust vectored 15". Figure 8(b) presents a sketch 

of the hypothesized internal flow phenomena for this 

configuration. For the range of nozzle pressure ratios of this investigation, all the data presented for 

this configuration are at underexpanded nozzle operating conditions. Similar to the 0" vectored nozzle, 

3 
flap static pressures on the 15" vectored nozzle are 

essentially independent of nozzle pressure ratio (see 

figs. 7(a) and 7(b)). Sidewall centerline pressures of 

the 15" vectored nozzle (fig. 7(c)) show less sensitivity to NPR than those noted previously for the 0' 

vectored nozzle (fig. 6(b)). 

The static-pressure data obtained on the 15" 

vectored nozzle configuration indicate internal flow 

phenomena (fig. 8(b)) similar to that discussed 

previously for the 0" vectored nozzle (fig. 8(a)). 

The flow phenomena shown in figure 8(b) are hypothesized from the pressure data, and other flow 

phenomena may explain the same static-pressure distribution trends. For example, an extreme thickening of the boundary layer could result in the same 

trends as those associated with the separation bubble hypothesized in figure 8. In addition, the sharp 

rises in static pressure on the upper and lower walls 

of the primary-flow constant-area duct (figs. 7(a) 

and 7(b)) shown in figure 8(b) as individual shocks 

or compressions could also be caused by a single 

shock which originates on the lower wall just downstream of the sonic line (x/q = l.lO), reflects off 

the upper wall (x/q = 1.45), and again strikes the 

lower wall (x/q M 1.58). However, individual shocks 

or compressions are shown in figure 8(b) because 

the sidewall centerline static-pressure distributions 

(fig. 7(c)) do not indicate passage of a compression 

wave. The static-pressure distribution presented in 

figure 7(a) indicates the possibility of a shock wave 

at x/xt M 1.44 on the upper flap. Figure 7(b) indicates the possibility of shock waves at x/q M 1.58. 

Also, the sidewall static-pressure distribution closer 

to the lower flap (fig. 7(d), q/vt = -0.452) tends to 

confirm the presence of two shock waves on the lower 

flap. 

The most striking similarity between the staticpressure distributions of the 15' vectored and 0" vectored configurations is the existence of a stagnation 

region on the primary-flow convergent-flap surface. 

(Compare figs. 6(a) and 6(d) with figs. 7(a) and 7(e).) 

The most striking dissimilarity between the two data 

sets is the asymmetry of the flow in the constant-area 

duct of the 15' vectored nozzle configuration. Staticpressure distributions on the upper flap of the 15" 

vectored nozzle (fig. 7(a)) show no indication of a 

separation bubble near the duct entrance, whereas a 

large separation bubble is hypothesized on the lower 

flap (fig. 7(b)). Although a compression (shock) wave 

could account for the sharp increase in static pressure 

on the lower flap from x/q = 1.10 to x/xt 1.35, 

the large range of x/xt over which this static-pressure 

rise is "smeared" more likely is explained by the 

normal supersonic flow deceleration which occurs in 

a converging duct. A converging duct in this case 

is caused by the large separation bubble (or thick 

boundary layer) on the lower flap. The separation 

bubble on the upper flap of the 0" vectored nozzle (fig. 8(a)) has been eliminated on the upper flap 

of the 15' vectored nozzle because less flow turning is required. Conversely, the separation bubble 

on the lower flap of the 15" vectored nozzle is much 

larger than that on the lower flap of the 0" vectored 

nozzle because more flow turning is required during 

vectored-thrust operation. Other signs of flow asymmetry in the vectored-thrust, constant-area duct are 

location of the sonic line (x/xt M 1.20 on the upper 

flap and x/xt < 1.10 on the lower flap) and location 

of the shock or compression waves (x/q M 1.44 on 

the upper flap and x/xt = 1.25 and xlxt = 1.58 on 

the lower flap). The existence of a second sonic line 

at the 15' vectored nozzle exit is not shown in figure 8(b) since, unlike the 0' vectored nozzle, static 

pressures on both the upper and lower flaps indicate 

additional expansion (downstream of last shock or 

compression) as the exit is approached. 

Concluding Remarks 

An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to measure staticpressure distributions inside a nonaxisymmetric 

nozzle with simultaneous partial thrust reversing 

(50-percent deployment) and thrust vectoring of 

the primary (forward-thrust) nozzle flow. Geometric forward-thrust-vector angles of 0" and 15" were 

tested. Test data were obtained at static conditions 

while nozzle pressure ratio was varied from 2.0 to 4.0. 

The three-dimensional nature of internal exhaust 

flow in a nonaxisymmetric nozzle was indicated by 

variations in pressure between the nozzle flaps and 

the sidewalls. Additionally, a complicated asymmetric exhaust flow pattern was evident in the primaryflow exhaust duct of the 15" vectored nozzle which 

was not evident for the 0" vectored nozzle. 

NASA Langley Research Center 

Hampton, Virginia 23665-5225 

December 31, 1986 

I 

4 
~ References 

1. 

2. 

3. 

4. 

5. 

1 

Capone, Francis J.: The Nonaxisymmetric Nozzle-It Is 

for Real. AIAA Paper 79-1810, Aug. 1979. 

Dusa, D. J.; Speir, D. W.; Rowe, R. K.; and Leavitt, 

L. D.: Advanced Technology Exhaust Nozzle Development. AIAA-83-1286, June 1983. 

Dusa, D. J.; Speir, D. W.; and Dunbar, D. K.: MultiFunctional Nozzle for Advanced Weapon Systems. SAE 

1989 Itansactions, Section 4, Volume 92, SOC. of Automotive Engineers, Inc., c.1984, pp. 4-54-4-64. (Available 

as SAE Paper 831426.) 

Mason, Mary L.; Putnam, Lawrence E.; and Re, 

Richard J.: The Eflect of Throat Contouring on TwoDimensional Converging-Diverging Nozzles at Static Conditions. NASA TP-1704, 1980. 

Putnam, Lawrence E.; and Strong, Edward G.: Internal Pressure Distributions for a Two-Dimensional 

Thrust-Reversing Nozzle Operating at a Free-Stream Mach 

Number of Zero. NASA TM-85655, 1983. 

6. Re, Richard J.; and Leavitt, Laurence D.: Static Internal 

Performance Including Thrwt Vectoring and Reversing of 

Two-Dimensional Convergent-Divergent Nozzles. NASA 

7. Carson, George T., Jr.; Capone, Francis J.; and 

Mason, Mary L.: Aeropropulsive Characteristics of 

Nonaxisymmetric-Nozzle Thrwt Reversers at Mach Numbers From 0 to 1.20. NASA TP-2306, 1984. 

8. Peddrew, Kathryn H., compiler: A User’s Guide to the 

Langley 16-Foot Itansonic ITLnnel. NASA TM-83186, 

1981. 

9. Shapiro, Ascher H.: The Dynamics and Thermodynamics 

of Compressible Fluid Flow. Volume I. Ronald Press Co., 

c.1953. 

TP-2253, 1984. 

5 
Table I. Internal Static-Pressure Orifice Locations 

(a) Sidewall surface for 0' vectored primary flow (see fig. 5(a)) 

6 
x, in. 

0.595 

.645 

.795 

395 

.995 

1.195 

1.345 

1.395 

1.495 

0 

X 

X 

X 

X 

X 

X 

X1.t 

0.373 

.404 

.498 

.561 

.624 

.749 

343 

375 

.937 

0.095 0.219 0.343 0.467 0.590 0.714 0.838 0.933 

X X X 

X X X X X 

X X X X X X 

X X X X X X X 

X X X X X X X 

X X X X 

X X 

X 

Table I. Continued 

(b) Sidewall surface for 15' vectored primary flow (see fig. 5(b)) 

w, in. 

0 I 0.100 I 0.230 I 0.360 I 0.490 I 0.620 I 0.750 I 0.880 I 0.980 I 1.080 

1.692 

1.810 

1.837 

2.052 

2.078 

2.293 

2.320 

2.535 

2.561 

2.805 

1.061 

1.135 

1.152 

1.287 

1.303 

1.438 

1.455 

1.589 

1.606 

1.759 

1.029 

X 

X 

X 

7 
Table I. Continued 

x', in. 

0.120 

(c) Upper flap surface for 0' vectored primary flow; 

NS 1.095 to NS 1.143 (see fig. 5(c)) 

y, in. 

x, in. x/xt -0.125 

1.217 0.763 X 

y, in. 

x', in. x, in. 

0.693 

.360 1.126 .706 

1.291 

1.365 

1.438 

1.512 

(d) Upper flap surface for 0' vectored primary flow; 

NS 1.143 to NS 1.595 (see fig. 5(c)) 

.809 

.856 

.902 

.948 

.240 

.360 

.480 

.600 

X 

X 

X 

X 

(e) Upper flap surface for 0' vectored primary flow; 

NS 1.595 to NS 3.049 (see fig. 5(c)) 

x', in. 

0.120 

.720 

.840 

1.080 

1.200 

8 

x, in. 

1.715 

1.835 

1.955 

2.315 

2.435 

2.675 

2.795 

x/xt 

1.075 

1.150 

1.226 

1.451 

1.527 

1.677 

1.752 

y, in. 

X 
Table I. Continued 

0.120 

.360 

(f) Upper flap surface for 15' vectored primary flow; 

NS 1.095 to NS 1.143 (see fig. 5(d)) 

1.105 0.693 X 

1.126 .706 X 

I 21. in. I x. in. I xlxt 

XI, in. 

0.120 

"l 

x, in. x/xt -0.125 

1.217 0.763 X 

(g) Upper flap surface for 15O vectored primary flow; 

NS 1.143 to NS 1.595 (see fig. 5(d)) 

1.291 

1.365 

1.438 

1.512 

I I I I u. in. 

.809 

.856 

.902 

.948 

x, in. 

1.711 

1.827 

1.943 

2.175 

2.290 

2.406 

2 638 

2.754 

.240 

.360 

.480 

.600 

y, in. 

XIXt 0 

1.073 X 

1.145 X 

1.218 X 

1.364 X 

1.436 X 

1.508 X 

1.654 X 

1.727 X 

X 

X 

X 

X 

(11) Upper flap surface for 15O vectored primary flow; 

NS 1.595 to NS 3.049 (see fig. 5(d)) 

XI, in. 

0.120 

.240 

.360 

.600 

.720 

.840 

1,080 

1.200 

9 
Table I. Concluded 

(i) Lower flap surface for 15' vectored 

NS 1.595 to NS 3.049 (see fig. 

XI, in. 

0.120 

.240 

.360 

.480 

.720 

.840 

.960 

1.080 

1.200 

primary flow; 

5(e)) 

x, in. 

1.711 

1.827 

1.943 

2.059 

2.291 

2.406 

2.522 

2.638 

2.754 

./.t 

1.073 

1.145 

1.218 

1.291 

1.436 

1.508 

1.581 

1.654 

1.727 

10 
ORIGINAL PAGE IS 

OE POOR QUALITY 

EtP n 

Table 11. Internal-Static Pressure Ratios 

(a) Sidewall surface for 0' vectored nozzle 

P/Ptj for X/Xt of - 

.404 -496 -624 -749 -075 ,937 1.122 1.216 1.373 1.530 1.687 1.843 

~ ~~ 

1.001 

2.002 

2.4'46 

3.G15 

3.500 

3.997 

.999 1.001 1.000 1.000 1.000 1.000 1,000 1.000 1.000 1.000 1.000 1.000 

-965 -964 e963 -954 e913 .I356 -305 e393 -420 -437 e456 e500 

-964 ,962 -963 e954 -913 e836 e300 e391 -416 e432 e441 e453 

,964 -963 e963 .954 e913 -856 -307 -385 -410 ,427 -430 -451 

-965 -965 e963 -954 -913 .85t e390 ,384 e409 e424 ,439 e453 

a965 e966 e964 e953 ,914 a856 ,391 ,382 e408 e425 e441 ,457 

P/Ptj for X/Xt of- - 

-__ 

NP R -561 -624 e749 .ti43 e937 1.063 1.157 1.313 1.470 1.627 

1.001 1.001 1.001 1.000 1.000 1.000 1.002 1.oco 1.000 1.000 1.000 

2.002 -960 -960 e930 -930 e850 e407 ,401 -402 -441 ,443 

2.496 -961 ,961 a930 ,930 .e50 ,490 e400 e397 e444 ,436 

3.GA'J e962 -962 ,930 ,930 e850 e490 e398 ,393 e443 e434 

3.997 ,965 e963 ,933 e933 -851 e491 e398 a393 e447 -433 

3.500 -964 .963 .932 .932 .850 .491 e399 .392 .445 .432 

WIWt = 0.219 

NP K 

1.001 

2.002 

2.496 

3.015 

3.500 

3.997 

P/Pt,j for Z/Zt of - 

,561 ,624 e749 -043 

1.001 1.001 1.000 1.001 

e968 ,964 -956 .933 

e967 a965 -957 e935 

a967 e965 ,959 e936 

e968 e967 -961 -936 

-968 -967 -963 .937 

11 
Table 11. Continued 

(a) Continued 

12 

I -- I 

hPR 1 e404 ,498 -561 :;:: ,749 1 

1.001 1.000 1.000 1.00~ -999 1.002 

2.002 ,963 e969 .e970 ,976 e969 

2.'+90 ,902 a969 a971 e976 e968 

3.015 -962 -960 my75 -976 e960 

3.500 e963 e969 e974 ,978 e969 

3.997 -964 -971 -976 ,970 

I I 1 

i .001 

2.002 

2.496 

3.015 

3.300 

3.997 

1.000 1.001 1.001 1.000 

a955 -963 -970 -974 

-954 -961 -969 e975 

,955 e962 e970 e976 

-957 -963 -971 e976 1 .950 .963 a971 ,977 
ORIGINAL PASE Erj 

OF POOR QUALITY 

EtP R 

1.001 

2.002 

2.496 

3.~15 

3.500 

3.997 

Table 11. Continued 

(a) Concluded 

PIPtj for x/xt of- 

,373 e404 e498 e624 

1.OC1 1.000 1.001 1.000 

.Y58 .934 .954 .955 

-924 .930 .951 -952 

.924 .930 .952 .951 

-923 -931 .952 .949 

.PL5 -931 -952 e949 

W/Wt = 0.714 

I P/Pt,j for ./.t of - 

-373 -404 e498 e561 e624 

1.001 1.000 1.000 1.001 1.000 

~949 -953 -963 -971 -968 

-949 -951 -962 e971 -967 

,949 e951 e963 e973 ,967 

.94V -952 -963 e974 e967 

.Y50 ,952 -964 ,975 -967 

1.001 

2 .ooz 

3.015 

3.500 

w/Wt = 0.838 

WIWt = 0.933 

I NPR I -498 -561 e624 I 

2 .OG2 

3.015 

3.997 

2 -496 

1.ou1 1.000 1.000 -1 

.932 .943 .V38 

.932 .943 .937 

W/Wt = 1.029 I I 1 

hP R -496 .56i e624 

3.505 

13 
Table 11. Continued 

(b) Sidewall surface for 15' vectored nozzle 

t 

1.061 1.151 1.302 1,454 1.605 1.758 

e999 1.000 1.000 1.003 e999 1.000 

e656 e463 a406 e309 e416 e443 

-656 e464 e404 e387 -414 -433 

e655 -465 e402 e385 -412 -434 

,655 e466 ,400 e383 ,413 e437 

e656 -467 e398 e383 a414 -440 

WIWt = 0 

NP K 

1.060 

2.006 

2.503 

3.001, 

3.514 

4.011 

P/Ptj for ./.t of - 

-404 ,498 e624 a749 e875 -937 

-999 1.001 1.000 1.003 1.000 1.000 

-969 e966 -967 e959 ,926 e862 

e967 ,966 ,966 e958 e926 e881 

-968 -967 -966 -958 e926 e881 

-968 -968 ,966 -958 -926 e861 

e966 -969 -967 e958 e926 e881 

hP 13 

1.000 

2.036 

2-50> 

3.006 

3.514 

4.011 

- 

-561 -624 -749 ,843 e937 

I 

.999 1.001 .999 .999 1.000 

e965 -966 e941 e941 e879 

e965 -966 -940 e940 e878 

-965 e966 e940 e940 -878 

e966 e967 e942 -942 e878 

,967 e966 e943 -943 ,879 

1 hPR 

1.000 

2.006 

2.505 

3.006 

3.514 

4.011 

e447 a380 a436 

-565 e379 e446 a381 -436 

e564 -379 e446 e300 e435 

e564 e386 e446 ,382 e435 

,561 -624 -749 -843 

1.001 .99b 1.000 1.000 

e972 -966 e962 e942 

,970 .968 .962 ,942 

,971 -969 -963 e943 

-970 e970 e965 e943 

-971 -971 e967 e944 

14 
Table 11. Continued 

(b) Continued 

I 1 

NP- i .404 .498 ::5t --.624 .7& 

l.000 .998 1,000 1.000 .997 .999 

2.006 .9tb -971 -974 e979 e974 

2.503 e966 -972 e974 -979 a972 

3.U06 e960 -971 a975 .Y80 e972 

3.514 ,907 -971 e976 e981 e973 

4.011 e967 e974 .Ye1 .973 

15 
Table 11. Continued 

(b) Concluded 

1.000 

2.UUb 

3.Ubb 

3.5i4 

4.011 

2.?bs 

16 

-999 1.001 -999 

-914 -933 e941 

-906 -930 -936 

.YO8 .929 .936 

.YO6 .C28 -936 

.909 .V28 .937 

-999 -9YY -998 1.000 -999 

,953 -525 e966 -976 e970 

.954 ,937 -967 .97Q -972 

.Y>3 .955 .96b ,977 .97C 

.953 .955 ,966 .977 .V70 

.9>3 .Y55 .9b7 .978 .YlO 

1.000 1.000 1.000 1.300 

-933 -941 -960 .9!19 

-930 .937 .956 .956 

e929 -936 e956 e955 

-950 .V36 -956 -953 

-530 -936 .V36 .953 

I I 1 

I 4 
OR1GINAL PACE IS 

OF POOR QUALITY 

NY R 

1.000 

2.504 

3.026 

3.511 

2.327 

4.OU;l 

Table 11. Continued 

.6V4 -700 e705 -762 -809 e854 ,903 -948 

1.004 -998 -999 e998 -999 1.001 1.000 1.001 

-950 -973 e987 1.006 -999 -985 -971 e935 

,949 -972 -987 1.CO7 e999 e987 -971 e935 

-947 ,974 -989 1.009 1.002 e990 ,973 e937 

.$52 ,974 -986 1.005 .Y96 .985 .C69 .934 

-948 ,973 -990 1.039 1.003 -991 -973 .937 

(c) Upper flap surface for 0' vectored nozzle 

hPR 

I I I 

PlPt,j for XIXt of - 

1.076 1.149 1.224 1.452 1.527 1.678 1.753 

1.GOO 

2.027 

2.>04 

3.d26 

3.511 

4.003 

~~~ ~~ ~ ~ 

1.006 1.000 1.001 1.000 1.001 -999 1.000 

a359 ,454 e377 e362 -346 -366 .42G 

a356 -456 e382 ,366 -337 e367 ,420 

,321 -437 -381 -367 -331 ,367 ,418 

-351 ,459 .383 e370 e331 e368 -417 

,349 e461 e385 ,372 e332 a370 -415 

17 
Table 11. Concluded 

(d) Upper flap surface for 15O vectored nozzle 

1.000 1.001 1.001 1.L131 1.031 1.033 1.001 1.000 1.000 

2.001 ,969 -989 1.004 1.003 4992 e976 -946 a710 

2.516 e966 ,987 i.035 1.002 .YY1 ,975 -945 e710 

3.011 e964 .9d5 1.005 l.r)Ul -992 ,973 -944 e710 

3.500 eY66 e986 1.007 1.002 e993 -977 e945 e711 

3.99'4 e966 -986 A.LO7 1.002 e433 e978 -944 e710 

$ 

1 I plp, 4 for xlxt of - 7 

P/Pt$ for x/xt of - 

NPR 1.145 1.219 1.366 1.437 1.513 1.653 1.727 

1,000 1.000 1,000 1.000 .Y99 1.053 1.001 1.000 

2.001 ,611 .4?0 ,234 -241 -415 ,456 -362 

L.5lb e609 ,489 e291 e232 e413 ,393 e363 

3.011 e610 ,491 a297 -236 e415 e394 e365 

3.999 ,612 a493 ,299 ,239 ,419 e395 e367 

4 

3.533 -612 .492 .L9Y .~39 .+Ab .395 .367 

, p/pt,j for x/xt of - 

NPR 1.072 1.145 1.229 1.296 1.437 1.510 1.564 1.653 1.727 

I 

1.030 1.000 1.coo 1.013 1.000 1.033 1.0uo 1.000 .999 1.003 

2.331 e418 .42R e323 a324 e306 ,448 ,444 e354 e240 

2.516 e417 -427 e307 -320 ,364 e448 e441 -345 .Z~L 

3*L111 -420 *431 e307 e323 ,367 -449 .442 e345 e239 

3.999 a423 -435 e304 ,323 ,369 a453 ,444 ,343 ,236 

3.53P e421 e434 ,304 e325 -334 .432 ,443 e344 .239 

18 
ORIGINAL PAX IS 

OF POOR QUALITY 

Y 

I Y 

I 

i 

ir 

f 

t 

~ 

! 

t 

1 i 

I 

1 

+, 

.7 

W 

M 

C 

0 

.A 

3 

3 

3 

N 

C 

19 
20 
NS 0.000 NS 3.323 

1.443 ' 7- 

1.050- 

II 

G 

f 

L ,073 

7 4 1.143 

41.095 + 

7 *. 900 

.700 k- 

/ 

/ 

L/ -375 radius \ 

/- Upper 

JAq I 

L- .221 

- Lower flap 

(a) 0' vectored primary flow. 

Figure 3. Geometric details of nozzle with reverser deployed 50 percent. All linear dimensions are in inches. 

21 
NS 0.000 NS 3.323 F 3.323 .L_____Jc 

3.049 w 

! 

.375 radius 

Upper flap 

/- 

Sidewall /- 

15. OOo kq L. 221 

\ Lower flap 

(b) 15' vectored primary flow. 

Figure 3. Concluded. 
0 

m 

3 

a 

0 

2 

23 
ORlGiNAL Pfi2.E IS 

Of POOR QUALITY 

24 
Reverser-f 1 ow 

converging surface 

converging surface 

constant-area duct 

NS 0.000 

Orifice (typ.) 

constant-area duct 

(a) 0' vectored primary flow. 

Figure 5. Internal static-pressure orifices of nozzle with reverser depioyed 50 percent. Orifice iocations are 

given in table I. All linear dimensions are in inches. 

25 
Reverser-f 1 ow 

converging surface 

NS 0.000- 

converging surface 

constant-area duct 

0000 0 

primary-flow 

constant-area duct 

(b) 15' vectored primary flow. 

Figure 5. Continued. 

26 
I NS 1.095 

IO 

0 

0 

4'NS 1.143 

0 

0 

0 

I 

B 

NS 1.143 

NS 1.595 

NS 1.595 

NS 3.049 

(c) Internal surface of 0' vectored upper flap (see fig. 3(a)). 

Figure 5. Continued. 

27 
NS 1.095 

X' 

X' TI Y 

I L 

7 

0 

01 

0 

0 

0 

I 

I 

NS 1.143 

NS 1.595 

NS 1.595 

NS 3.049 

(d) Internal surface of 15' vectored upper flap (see fig. 3(b)). 

Figure 5. Continued. 
1 I 

I 

NS 1.595 

NS 3.049 

(e) Internal surface of 15O vectored lower flap (see fig. 3(b)). 

Figure 5. Concluded. 

29 
,-Pressure orifice 

1.0 

.8 

.7 

.6 

.4 

.3 

.2 

.1 

0 4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 

x4 

(2) TvTpper flap cenierline pressures. 

Figure 6. Nozzle pressure distribution for 0' vectored configuration. 

30 
P 

pt, j 

1.0 

.9 

.8 

.7 

.6 

.5 

.4 

.3 

.2 

.1 

0 

4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 

xfit 

(b) Sidewall centerline pressures; W/W~ = 0. 

Figure 6. Continued. 

31 
1.0 

.9 

.8 

.7 

.6 

.5 

.4 

.3 

.2 

.1 

0 

(c) Variation of pressure across nozzle sidewall at NPR = 4.0. 

Figure 6. Continued. 
.95 .9 .8 .7 .6.5 

(d) Contours of p/ptj on Oo vectored nozzle sidewall for NPR = 4.0. 

Figure 6. Concluded. 

33 
34 

1.0 

.9 

.8 

.7 

.6 

.5 

.4 

.3 

.2 

.1 

0 

.6 .8 1.0 1.2 1.4 1.6 1.8 2.0 .4 

x4 

(a) Upper flap centerline pressures. 

Figure 7. Nozzle pressure distribution for 15" vectored configuration. 
1.0 

.9 

.8 

.7 

.6 

.5 

.4 

.3 

.2 

.1 

0 

.6 .8 1.0 1.2 1.4 1.6 1.8 2.0 .4 I 

~t 

(b) Lower flap centerline pressures. 

Figure 7. Continued. 

35 
36 

1.0 

.9 

.8 

.7 

.6 

.5 

.4 

.3 

.2 

.1 

n 

". 4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 

(c) Sidewall centerline pressures; w/wt = 0 for z/zt 5 1.0; q/qt = 0 for z/zt > 1.0. 

Figure 7. Continued. 
1.0 

.9 

.a 

.7 

.6 

.5 

.4 

.3 

.2 

.1 

0 

4 .6 .a 1.0 1.2 1.4 1.6 1. a 2.0 

x4 

(d) Variation of pressure across nozzle sidewall at NPR = 4.0. 

Figure 7. Continued. 

37 
38 

(e) Contours of p/pt,j on 15' vectored nozzle sidewall for NPR = 4.0. 

Figure 7. Concluded. 
Sonic line 

Separation bubble 

Shock wave or 

strong compression 

Boundary layer 

- Possible second 

sonic line 

(a) 0' vectored primary flow. 

Figure 8. Internal flow phenomena for nozzle with reverser deployed 50 percent. 

39 
Shock wave or 

--+ Subsonic flow 

t, 

Note: 

Supersonic flow in converging duct: velocity decreases, static pressure increases. 

Supersonic flow in diverging duct: velocity increases, static pressure decreases. 

(b) 15' vectored primary flow. 

Figure 8. Concluded. 

40 
Standard Bibliographic Page 

, Report No. 

NASA TM-89044 

2. Government Accession No. 

. Author(s) 

Robert S. Green and George T. Carson, Jr. 

.9. Security Classif.(of this report) 20. Security Classif.(of this page) 21. No. of Pages 

Unclassified Unclassified 41 

. Performing Organization Name and Address 

NASA Langley Research Center 

Hampton, VA 23665-5225 

22. Price 

A03 

2. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

3. R.ecipient's Catalog No. 

5. Report Date 

March 1987 

6. Performing Organization Code 

505-62-91-01 

8. Performing Organization Report No. 

L-16210 

10. Work Unit No. 

11. Contract or Grant No. 

13. Type of Report and Period Covered 

Technical Memorandum 

14. Sponsoring Agency Code 

5. Supplementary Notes 

Robert S. Green: Wright-Patterson Air Force Base, Dayton, Ohio. 

George T. Carson, Jr.: Langley Research Center, Hampton, Virginia. 

6. Abstract 

An investigation was conducted in the Langley 16-Foot Transonic Tunnel at static conditions to measure 

the pressure distributions inside a nonaxisymmetric nozzle with simultaneous partial thrust reversing 

(50-percent deployment) and thrust vectoring of the primary (forward-thrust) nozzle flow. Geometric 

forward-thrust-vector angles of 0' and 15' were tested. Test data were obtained at static conditions while 

nozzle pressure ratio was varied from 2.0 to 4.0. Results indicate that, unlike the Oo vector angle nozzle, a 

complicated, asymmetric exhaust flow pattern exists in the primary-flow exhaust duct of the 15' vectored 

nozzle. 

-7. Key Words (Suggested by Authors(s)) 

Nonaxisymmetric nozzles 

Thrust reversing 

Thrust vectoring 

18. Distribution Statement 

Unclassified-Unlimited 

For sale by the National Technical Information Service, Springfield, Virginia 22161 

NASA-Langley, 1