NASA Technical Reports Server (NTRS) 20040111313: A Parametric Investigation of Nozzle Planform and Internal/External Geometry at Transonic Speeds

23 Pages·1.8 MB·English

by  NASA Technical Reports Server (NTRS)

A Parametric Investigation of Nozzle Planform and Internal /
External Geometry at Transonic Speeds
Daniel L. Cler*
NASA Langley Research Center, Hampton, Virginia
ABSTRACT INTRODUCTION
An experimental investigation of Potential future military fighter aircraft
multidisciplinary (scarfed trailing edge) nozzle will be complex multi-role vehicles with special
divergent flap geometry was conducted at features designed to counter both enemy air
transonic speeds in the NASA Langley 16-Foot attack and ground defensive action (refs. 1-6).
Transonic Tunnel. The geometric parameters Because of multidisciplinary design issues,
investigated include nozzle planform, nozzle nozzle exhaust systems require special shaping
contouring location (internal and/or external), in addition to providing high performance (ref.
and nozzle area ratio (area ratio 1.2 and 2.0). 7). This special shaping usually takes the form
Data were acquired over a range of Mach of scarfing or angling of nozzle trailing edges.
Numbers from 0.6 to 1.2, angle-of-attack from However, scarfing of nozzle trailing edges can
0.0" to 9.6" and nozzle pressure ratios from 1. O affect integrated wind-on performance.
to 20.0. Results showed that increasing the Optimizing the nozzle shape to meet many
rate of change internal divergence angle across mission requirements, without compromising
the width of the nozzle or increasing internal thrust performance and nozzle boattail drag is
contouring will decrease static, aeropropulsive the objective of a good nozzle design. Several
and thrust removed drag performance investigations have looked at the internal
regardless of the speed regime. Also, performance of nozzles designed to address
increasing the rate of change in boattail angle multidisciplinary issues and have found
across the width of the nozzle or increasing minimal impact on performance (ref. 8-10), but
external contouring will provide the lowest only a few have looked at the nozzle boattail
thrust removed drag. Scarfing of the nozzle drag from a parametric view (ref. 11).
trailing edges reduces the aeropropulsive
performance for the most part and adversely Thirteen nozzle configurations were
affects the nozzle plume shape at higher nozzle tested on an isolated (no tails or wings) two-
pressure ratios thus increasing the thrust dimensional body in the NASA Langley 16-
removed drag. The effects of contouring were Foot Transonic Tunnel to determine the effects
primary in nature and the effects of planform on aeropropulsive and thrust removed drag
were secondary in nature. Larger losses occur performance of various geometric parameters.
supersonically than subsonically when scarfing The geometric parameters investigated include
of nozzle trailing edges occurs. The single nozzle planform, nozzle contouring location
sawtooth nozzle almost always provided lower (internal and/or external), and nozzle area ratio
thrust removed drag than the double sawtooth (area ratio 1.2 and 2.0). Nozzle pressure ratio
nozzles regardless the speed regime. If internal was varied from 1.0 to 20.0, Mach Number
contouring is required, the double sawtooth was varied from 0.6 to 1.2, and angle-of-attack
nozzle planform provides better static and was varied from 0" to 9.6"
aeropropulsive performance than the single
sawtooth nozzle and if no internal contouring is SYMBOLS
required the single sawtooth provides the
highest static and aeropropulsive performance. Ae nozzle exit area, in2
* 414 area ratio (see figure 3)
Aerospace Engineer, Research Facilities Branch
4 nozzle throat area, in 2
Copyright 0 1995 by the American Institute of Cd-t thrust removed drag coefficient of
Aeronautics and Astronautics, Inc. No copyright is
afterbody
asserted in the United States under Title 17, U.S. Code. The
U.S. Government has a royalty-free license to exercise all CPJI External nozzle pressure coefficient,
rights under the copyright claimed herein for Governmental
(P-P-Wl
purposes. All other rights are reserved by the copyright
owner. D" nozzle drag, Ibf
1
F measured thrust along body axis,
maximum nozzle boattail angle, deg
Pmax
lbf
(see figure 3)
Fi ideal isentropic gross thrust, lbf,
minimum nozzle boattail angle, deg
Pmin
(see figure 3)
maximum nozzle internal
'max
divergence angle, deg
measured normal force, lbf
(see figure 3)
resultant gross thrust, lbf,
+ + m',i minimum nozzle internal divergence
J F ~F ; F;
angle, deg
measured side force, lbf (see figure 3)
acceleration due to gravity,
total change in boattail angle, deg
AP
32.174 ft/sec2
average nozzle exit height or height
Pmax-Pmin
at x = 8.25 in, in (see figure 3), A6 total change in internal divergence
(h$de 4- htip)/2 angle, deg
height of nozzle exit at nozzle
hside
sidewall, in (see figure 3) m' ax-m' in
h, nozzle internal flowpath height Y ratio of specific heats, 1.3997 for
at throat, h, = 0.861 in (see air
figure 3)
@/Jy rate of change in boattail angle
height of nozzle exit at nozzle
across width of the nozzle, deg/in
apex, in (see figure 3)
total length of nozzle, in (6 max-6mi(n2)w/(, )(# of sawteeth))
(see figure 3) %/dy rate of change in internal divergence
total length of nozzle from
angle across width of the nozzle,
attachment station to exit at
deg/in
sidewall, in (see figure 3)
Mach Mach Number (6 max-6mi(n2)w /(, ) (# of sawt eeth))
NPR nozzle pressure ratio, p,,j/p,
ABBREVIATIONS
NPR, design nozzle pressure ratio for
fully expanded flow at the nozzle
BL butt line
exit
ref. reference
local external static pressure, psia
2-D two-dimensional
atmospheric pressure, psia
C-D convergent-divergent
average jet total pressure, psia
FS fuselage station
freestream static pressure, psia R radius
tunnel freestream dynamic pressure, WL water line
psi
gas constant, 1716 ft2/sec2-"R APPARATUS AND PROCEDURES
jet total temperature, OR
nozzle internal flowpath width Wind Tunnel
at throat, w, = 4.972 in
X axial distance measured from the This test was conducted in the NASA-
nozzle connect (positive Langley 16-Foot Transonic Tunnel. A detailed
downstream), in (see figure 3) description of the facility is given in reference
Y lateral dimension measured 12. The tunnel is a single-return atmospheric
from nozzle centerline (positive wind tunnel with a slotted octagonal test section
toward left wing), in. and continuous air exchange. The wind tunnel
(see figure 3) has a continuously variable airspeed up to a
Z vertical dimension measured from Mach Number of 1.30. Test-section plenum
nozzle centerline (positive down), suction is used for speeds above Mach = 1.05.
in. (see figure 3)
a angle of attack, deg
2
Model and Support System
When the nozzle trailing edge angle
A sketch of the sting-strut-supported changes from being perpendicular to the nozzle
single-engine simulator with a typical nozzle centerline to some angle, there are several
(OO),
installed is shown in figure 1 and photographs methods available to fair the nozzle trailing
of a single and double sawtooth nozzle edge into the afterbody. One method of fairing
planforms installed in the wind-tunnel are is continuous curvature (streamwise and
shown in figure 2. The model is an isolated spanwise change in curvature) of the nozzle
(no vertical or horizontal tails) 2-D fuselage surface (see ref. 11). The method used in the
model with a forebody section, a centerbody investigation being discussed in this paper
and a nozzle. The forebody of the model is involved spanwise fairing of linear segments
non-metric (not on the balance) and the resulting in a “warped planar surface”. The
centerbody and nozzle are metric (on the degree of warping depended upon whether the
balance). The metric break is located at contouring would be accomplished externally,
fuselage station 27.000 and the nozzle internal internally, or both internally and externally.
round-to-square transition begins at fuselage The internal contoured configurations force all
station 53.000. The nozzle connect station is at warping to occur internally resulting in a
fuselage station 54.486. continuously varying area ratio form the nozzle
centerline to the sidewall. The external boattail
Single-Engine Propulsion System angle is constant. The external contoured
configurations affect only the external boattail,
An external high-pressure air system hence the local boattail angle changes with
provides a continuous flow of clean, dry air at changing butt line. The internal nozzle contour
a stagnation temperature of approximately is planar with a constant internal divergence
540”R at the nozzle. As shown in figure 1, the angle. For a given area ratio, the combined
pressurized air is transferred from the supply internallexternal contour case results in both the
source to the simulator by six air lines that run internal nozzle and external nozzle boattail
through the support strut and into a high- having warped planar contours resulting in
pressure plenum chamber. The air is then maximum values of local boattail and internal
discharged perpendicularly into the model low- divergence angles which are approximately half
pressure plenum through eight multi-holed of those realized for the other two contouring
sonic nozzles equally spaced about the high- methods.
pressure plenum. The high-pressure plenum is
separate from the balance system (nonmetric), Figure 3 shows sketches of the selected
but the low-pressure plenum is attached to the planform and contouring geometries with a
balance (metric). This system is designed to table of all pertinent dimensions and figure 4
minimize axial momentum forces generated by shows three-dimensional sketches of several of
the air as the air passed from the nonmetric the nozzle geometries. The baseline and single
high-pressure plenum to the metric low- sawtooth planforms were tested at two values
pressure plenum. Two flexible metal bellows of area ratio (1.2 and 2.0) and the double
seal the air system between the plenums and sawtooth planfonns were tested at only the area
compensated for pressurization forces. From ratio 1.2. The single sawtooth nozzles had a
the low-pressure plenum, the air passes single protruding sawtooth on the nozzle
through a circular choke plate / flow divergent flap with the apex of the sawtooth
straightener into an instrumentation section and being on the nozzle centerline. The double
then into the nozzle. sawtooth nozzles had two protruding sawteeth
with the apexes at 2y/w, = -0.50 and 0.50.
Nozzle Design Both planforms utilized a 45” scarf angle. In
addition, an externally contoured single
As discussed previously, a parametric sawtooth configurations with a modified
investigation was conducted to examine the “i~entropici’n~t ernal contour which provided a
impact of nozzle planform on the 0” divergence angle at the nozzle trailing edge
aeropropulsive characteristics of isolated 2-D was examined.
convergent-divergent nozzles. Two baseline
planforms (0’ trailing edge angle) and eleven Referring to figure 3, L is the overall
other nozzle configurations (with non-zero length of the nozzle and L, is the length of the
trailing edge angles) were investigated. nozzle at the sidewall. For all nozzles tested,
3
the median length of the scarf tip was kept independent of planform as dp/dy was.
constant at 8.25 inches, the baseline nozzle
length as shOwn in figure 3. Referring to
Instrumentation
figure 3 again, the external boattail angle and
internal divergence angle were held constant for
A six-component strain-gauge balance
the baseline nozzle or Pmin= , ,p and 6mi=n 6,, was used to measure forces and moments on
Pin the metric section of the model. The measured
where is the external boattail angle at the
weight-flow rate was obtained from a multiple-
,,p
contour peak, is the external boattail angle critical venturi located in the high pressure air
at the contour valley, 6,, is the internal system. Total pressure in the jet was measured
by a ten-probe rake in the instrumentation
divergence angle at the contour peak and 6maixs
section (see fig. 1). The nozzle total pressure
the internal divergence angle at the contour
was computed as the average of the individual
valley. For the externally contoured nozzles, total pressures. An iron-constantan
Pmin< ,,p, and 6,in = 6,,,. For the combined thermocouple positioned aft of the rake plane
contour nozzles, Pmin< ,,p and 6,, 6,,,. measured jet total temperature. Total-pressure
and venturi static pressure measurements were
pmin
For the internally contoured nozzles, = made with individual pressure transducers.
Nozzle static pressures were measured using
and <
Pmax 6max. electronic pressure scanners. Flow
visualization information using an oil, kerosene
Also shown in figure 3 are Ap, A6,
and dry powder paint mixture was acquired.
dp/dy and d6/dy for the various configurations
Data Reduction
tested. Ap is the total change in external
boattail angle across the width of the nozzle.
Fifty frames of data, acquired over a 5-
As the planform goes from single to double,
second sample interval, were averaged for each
A p decreases by approximately half, therefore measured data parameter at each data point.
The averaged values were used in all
Ab is a measure in a sense of the change in
subsequent computations. Each balance
planform. A6 is the total change in internal component was corrected for model weight
divergence angle across the width of the tares, balance-component interactions, model-
nozzle. As the planform goes from single to installed balance interactions, and model base
double, A6 decreases by approximately half. and internal pressurization effects. The
procedure for correcting the balance
therefore A6 is another measure of the change
measurements is documented in references 12
in planform. dp/dy is the rate of change in and 13.
boattail angle across the width of the nozzle.
As the amount of external contouring decreases Performance results are presented as
(i.e. the amount of internal contouring is static thrust ratio, F/F,, thrust-minus-nozzle
drag ratio, (F-Dn)/Fi and thrust removed drag
increasing), @/dy is decreasing. Even though
for the single sawtooth external contoured coefficient, Cd-t.T he static thrust ratio, F/Fi, is
representative of the axial static thrust
nozzle Ap is twice as large as the double
efficiency. The ideal thrust, Fi, is calculated
sawtooth external contoured nozzle, dp/dy is using the nozzle total pressure, total
almost the same. Therefore @/dy is temperature, measured weight flow, and
assumes isentropic flow (see Symbols section).
independent of planform and only measures
The thrust-minus-nozzle drag ratio, (F-Dn)/Fi,
changes in contour location (amount internally
or externally). d6/dy is the rate of change in is representative of the total installed
aeropropulsive efficiency of the nozzle. The
internal divergence angle across the width of
thrust removed drag coefficient, Cd.t, is
the nozzle. As the amount of internal
representative of the external drag with the
contouring decreases (Le. the amount of
internal jet effects removed.
external contouring is increasing), d6/dy is
decreasing. Also, d6/dy is somewhat
4
External static pressure data are divergence angle, d6/dy, and rate of change in
presented as pressure coefficient, C,, = (p-
boattail angle, dp/dy, on static, aeropropulsive
pm)/q, and plotted as a function of x/L, a
and thrust removed drag performance for both
nondimensional term representative of orifice the single and double sawtooth nozzles (no
location. Flow visualization photographs are baseline nozzles) are presented in figure 10 for
also included to provide details of flow static conditions at NPR = 2.0, 3.9 & 10.0, in
behavior. Minimal internal static pressure data
figure 11 for Mach = 0.8 and a = 0.0" at NPR
were acquired, but are not presented herein.
= 2.0, 3.9 & 10.0 and in figure 12 for Mach =
Accuracy 1.2 and a = 0.0" at NPR = 8.0 & 10.0. The
effects of contouring location on external
The calibrated balance instrument pressure at the nozzle sidewall centerline,
accuracy (this is not an uncertainty) is estimated divergent flap edge and divergent flap
centerline are shown in figures 13 and 14 for
to be 0.26% of full scale (f2.08 lbs) in axial
Mach = 0.80 at NPR = 3.9 and Mach = 1.20 at
force and 0.44% of full scale (k2.65 lbs) in NPR = 10.0 respectively. Data are compared
normal force. Based on the model reference with the unscarfed baseline nozzle for the area
cross sectional area (not wing area, so numbers ratio 1.2 nozzle.
will be an order of magnitude larger than
typical airplane drag counts), Am = 42.396 Static Performance- In figure 5 it can
in2, and the balance accuracy only, the drag be seen that the internally contoured nozzles
have the lowest static performance. In figure
coefficient accuracy is f0.0170 at Mach =
10 it is shown that the static performance
0.60, f0.0114 at Mach = 0.80, k0.0099 at decreases with increasing rate of change in
Mach = 0.90, f0.0095 at Mach = 0.95, and internal divergence angle, d6/dy, regardless of
the nozzle pressure ratio. As discussed earlier,
kO.0081 at Mach = 1.20.
the rate of change in internal divergence angle
is nearly directly proportional to the amount of
The pressure measurement accuracy of
internal contouring. Also in figure 10 it can be
the 150 psi jet total pressure transducers was
seen that increasing the rate change in the
f0.1 50 psi. The pressure measurement
boattail angle, dp/dy, increases performance.
accuracy of the 2000 psi multiple critical
Since external contouring does not directly
venturi pressure transducers used to calculate
affect static performance, the effect of
mass flow was f2.0 psi. The pressure
increasing performance with dp/dy is simply
measurement accuracy of the 5 psi metric break
due to the fact that increasing external
pressure transducers used to correct balance
contouring decreases internal contouring
data was f0.005 psi. The electronically
required for a given nozzle configuration. The
scanned pressure modules used to measure the
nozzle with the highest rate of change in
external surface static pressure had an accuracy
boattail angle has the lowest rate of change in
of f0.015 psi. internal divergence angle and hence the highest
static performance. It can be concluded that
RESULTS AND DISCUSSION increasing the rate of change in internal
divergence angle or increasing internal
Effect of Contouring Location contouring will decrease static performance.
Results showing the effects of Wind-On Performance- In figure 6 for
contouring location on the static performance Mach = 0.8 it can be seen that increasing the
are shown in figure 5. The aeropropulsive amount of internal contouring decreases
performance results are presented in figures 6 aeropropulsive performance for the single and
and 7 for Mach = 0.80 and Mach = 1.20, double sawtooth nozzles. The results are
respectively and the thrust removed drag similar in figure 7 for the supersonic condition
performance is presented in figures 8 and 9 for at Mach = 1.2. The baseline nozzle had the
Mach = 0.80 and Mach = 1.20, respectively. highest overall performance, other than the
The effects of the rate of change in internal single sawtooth externally contoured nozzle in
the subsonic regime, thus indicating the
5
scarfing of nozzle trailing edges reduces 10.0, respectively. The nozzle pressure ratios
aeropropulsive performance for the most part. chosen were meant to be somewhat indicative
In figures 8 and 9 it is shown that increasing of actual throttle settings at the given Mach
the amount of internal contouring decreases the number. The subsonic data (figure 13) all
thrust removed drag performance. The show a large expansion region just downstream
baseline nozzle had substantially lower thrust of the nozzle connect station (A= 0) resulting
removed drag at the higher nozzle pressure from flow turning over the shoulder. Negative
ratios, thus indicating that nozzle scarfing values of C of course result in drag if acting
increases external drag. For NPR < 4, the on areas wit% aft facing projected area, (such as
single sawtooth nozzles had lower drag than is the case here). Flow then begins to
the baseline nozzle subsonically however. The recompress just downstream of the nozzle
design nozzle pressure ratio is NPR, = 3.86 connect station (and shoulder) and by x/L=0.4
for the AJ4 =1.2. In addition, the externally to 0.5 actually becomes positive, hence has a
contoured single sawtooth nozzle had lower favorable effect on drag. It is interesting to
thrust removed drag than any of the scarfed note that all scarfed nozzle configurations
nozzles, but not lower than the baseline. This provided higher static pressures than the
is probably the reason that the single sawtooth baseline nozzle configuration. Based on this
externally contoured nozzle had higher observation, one would expect external drag to
aeropropulsive performance than the baseline be highest on the baseline at Mach = 0.8 and
nozzle because the more dominate internal NPR = 3.9. In figure 8 it is seen that at Mach
losses for the externally contoured single = 0.8 and NPR = 3.9 the baseline nozzle had
sawtooth nozzle were less than the baseline the highest external drag. However for NPR >
nozzle as indicated in figure 5, thus making up 4, the baseline nozzle has lower drag at Mach
for the increase in drag of the single sawtooth 0.8 and has lower drag at all NPR's
externally contoured nozzle. Also it can be supersonically as seen in figure 9. This would
seen in figures 9 and 10 that larger losses occur indicate that the pressure trends presented are
supersonically than subsonically when nozzle not necessarily typical at all NPR's and Mach
scarfing is done. Furthermore it can be seen Numbers. On the single sawtooth
that the thrust removed drag decreases much configuration where boattail angles are largest,
faster for the baseline nozzles than the scarfed it is apparent that contouring had an impact on
nozzles with increasing NPR thus indicating the recovered static pressure on the aft portion
that the plume shape for the scarfed nozzles of the nozzle. Static pressures on the nozzle
adversely affects the performance when boattail downstream of x/L=0.4 increased with
compared to the baseline nozzle, especially at increasing external contour. The effect of
the higher nozzle pressure ratios. In figures 11 contouring on the double sawtooth pressure
and 12 it is shown that increasing the rate of distributions was much less pronounced,
change in internal angle decreases probably as a result of the generally smaller
aeropropulsive performance as it did statically. boattail and internal divergence angles.
Also, it can be seen that increasing the rate of
change in external boattail angle increases The static pressure distributions
aeropropulsive performance. The fact that the presented for Mach = 1.2 and NPR = 10.0 are
wind-on effects do not change from the static shown in figure 14. As seen, the characteristic
effects would indicate that the internal geometry shape of the distributions are somewhat
changes affect performance more than external different than for the subsonic cases. The
geometry changes. The effects of rate of overexpansion region is farther downstream
change in boattail angle on thrust removed drag and the recovery process is considerably
are not as clear. Overall it appears that delayed. In fact, pressures generally remained
negative (unfavorable) over the entire nozzle
increasing @/ay decreases external drag. This
length. The data indicate that a shock and
would indicate that larger amounts of external
ensuing separation generally occurred on the
contouring for a scarfed nozzle provides the
nozzle boattail (as indicated by the steep
lowest external drag.
compression resulting from the shock,
followed by a "plateau" region of pressure).
External Pressure Results- Static
The baseline nozzle did not have lower
pressure coefficient data on the nozzle boattail
pressures than the rest of the configurations
are presented in figures 13 and 14 for Mach =
0.8 at NPR = 3.9 and Mach = 1.2 at NPR = over the entire length of the nozzle as it did at
Mach 0.8 and NPR =3.9. This matches the
6
results in figure 9 where the thrust removed Alp increases thrust performance. As stated
drag for the baseline indicates that it has the
previously in the nozzle design section,
lowest drag. This change in pressure results
increasing the number of sawteeth decreases
from subsonic to supersonic is most likely just
A6 and AD. This shows empirically why there
that different NPR's were plotted for the
are conflicting effects of the number of
different speeds and not the overall nozzle
characteristics since drag for the baseline nozzle sawteeth based on whether internal (As) or
decreased much faster with increasing NPR
external (Alp) effects are considered. The only
than for the scarfed nozzles.
general conclusions that can be drawn is that
for internally contoured nozzles, the double
EfSect of Nozzle Planform
sawtooth (or low A6) had higher performance
Results showing the effects of nozzle and for the externally contoured nozzles, the
planform on the static performance are shown
single sawtooth (or high AD) had higher
in figure 15. The aeropropulsive performance
performance.
results are presented in figures 16 and 17 for
Mach = 0.80 and Mach = 1.20, respectively
Wind-On Performance- The effects of
and the thrust removed drag performance is
nozzle planform on aeropropulsive
presented in figures 18 and 19 for Mach = 0.80
performance are presented in figures 14 and
and Mach = 1.20, respectively. The effects of
15. As seen, results are again highly
the total of change in internal divergence angle,
dependent upon contour location. However, it
A6, and total change in boattail angle, AD, on can be noted that as long as any internal
static, aeropropulsive and thrust removed drag contour is required, the double sawtooth
performance for both the single and double planform provided better performance than the
sawtooth nozzles (no baseline nozzles) are single sawtooth planform subsonically as
presented in figure 20 for static conditions at shown in figure 14. Both scarfed planforms
NPR = 2.0, 3.9 & 10.0, in figure 21 for Mach had lower performance than the baseline
nozzle. Result were significantly different
= 0.8 and a = 0.0" at NPR = 2.0, 3.9 & 10.0
when all of the contouring took place
and in figure 22 for Mach = 1.2 and a = 0.0" externally. Aeropropulsive performance of the
at NPR = 8.0 & 10.0. The effects of nozzle double sawtooth planform was lower than
planform on external pressure at the nozzle either the baseline nozzle or the single sawtooth
sidewall centerline, divergent flap edge and planform. At Mach = 1.2 (fig. 15), the single
divergent flap centerline are shown in figures sawtooth configuration generally provided
23 and 24 for Mach = 0.80 at NPR = 3.9 and slightly higher aeropropulsive performance
Mach = 1.20 at NPR = 10.0 respectively. Data than the double sawtooth regardless of contour
are compared with the unscarfed baseline method. Performance of the baseline nozzle
nozzle for the area ratio 1.2 nozzle.
was higher than either scarfed nozzle planform.
Figures 18 and 19 present the effects of
Static Performance- The effects of
planform on the thrust removed drag. As seen,
nozzle planform on internal performance are
the single sawtooth nozzles almost always had
presented in figure 15. As seen, contouring
lower thrust removed drag than the double
location had a large impact on the impact of
sawtooth nozzle regardless of the speed
planform shape on internal performance and no
regime. Supersonically the double sawtooth
general conclusions could be drawn. A
nozzles had significantly higher drag than the
conclusion could be drawn that planform
single sawtooth nozzles, especially the
effects are secondary in nature when compared
internally contoured nozzles. In figures 21 and
to contouring effects which are primary in
22 can be seen the effects of AD and A6 on
nature. However, it can be noted that as long
aeropropulsive and thrust removed drag
as any internal contour is required, the double
sawtooth planform provided better performance performance. The trends are similar to the
static results for the aeropropulsive
than the single sawtooth planform. Also in
figure 20 it can be seen that increasing the total performance. In addition for the thrust
removed drag performance it can be seen that
change in internal divergence angle, A6,
increasing AP decreases thrust removed drag
decreases the thrust performance and increasing
throughout the speed regime, though results are
7
not as clear as the aeropropulsive data. This occurs in a circular pattern on the nozzle
would indicate that increasing the number of divergent flap with the separation occurring
farther upstream at the nozzle centerline than at
sawteeth (decreasing AP) would increase the
the nozzle sidewall. Also, recirculation and
thrust removed drag as a general trend, though
flow separation occurs on the peak(s)
there are exceptions to this rule.
(minimum external boattail angle) of the nozzle
contour for both the single and double
External Pressure Results- The effects
sawtooth external contour configurations
of nozzle planform (for various contour
shown. The large differences in boattail angle
locations) on boattail pressure distribution are
across the nozzle width for both nozzle
presented in figures 23 and 24 for Mach = 0.8
configurations most likely cause the transverse
at NPR = 3.9 and Mach = 1.2 at NPR = 10.0,
flow and recirculation that occurs.
respectively. As seen in figure 23 for subsonic
speeds, pressures on the single sawtooth
CONCLUSIONS
planform tended to be higher than those
measured for either the baseline or double
The effects of contouring were
sawtooth planforms indicating that from an
consistent throughout the speed regime were
external drag viewpoint, the single sawtooth
and the effects of nozzle planform were not
planform would be the most favorable. This is
consistent between subsonic or supersonic
the case as indicated in figure 18 for NPR =
speeds. For this reason great care will be
3.9 where the single sawtooth had the lowest
needed to optimize a design that works well in
or equal thrust removed drag. At Mach = 1.2,
both subsonic and supersonic flight unless only
the effect of planform was generally more
point cruise performance is desired. The main
difficult to see, although it does appear as
conclusions from the test are as follows:
though the single sawtooth static pressure
distributions are slightly more favorable then
1) Increasing the rate of change internal
those of the double sawtooth, but still both of
divergence angle across the width of the nozzle
them less favorable than the baseline nozzle. In
or increasing internal contouring will decrease
figure 19 it was shown clearly that the baseline
static, aeropropulsive and thrust removed drag
nozzle had much lower thrust removed drag at
performance regardless of the speed regime.
NPR = 10.0 (actually off of scale shown) than
the single and double sawtooth nozzles and the
2) Increasing the rate of change in boattail
single sawtooth nozzle had lower drag than the
angle across the width of the nozzle or
double sawtooth nozzle at Mach = 1.2 thus
increasing external contouring will provide the
matching the pressure results.
lowest thrust removed drag.
Flow Visualization
3) Scarfing of the nozzle trailing edges reduces
the aeropropulsive performance for the most
Oil, kerosene and dry powder paint
part and adversely affects the nozzle plume
mixture flow visualization results are shown in
shape at higher nozzle pressure ratios thus
figure 21 for Mach = 0.8, NPR = 3.8 and a = increasing the thrust removed drag.
0.0" with AJ& = 1.2 for the single sawtooth
internal contour nozzle, single sawtooth 4) The effects of contouring were primary in
external contour nozzle and the double nature and the effects of planform were
sawtooth external contour nozzle. Flow is left- secondary in nature.
to-right across the nozzle. The single sawtooth
nozzle with internal contouring had a constant
5) Larger losses occur supersonically than
external boattail angle of 20.96". The single subsonically when scarfing of nozzle trailing
sawtooth nozzle with external contouring had edges occurs.
external boattail angles that varied from 16.8" at
the nozzle contour peak to 29.7" at the contour 6) The single sawtooth nozzle almost always
valleys. The double sawtooth nozzle with provided lower thrust removed drag than the
external contouring had external boattail angles
double sawtooth nozzles regardless the speed
that varied from 18.6" at the nozzle contour
regime.
peaks to 24.2" at the contour valleys. For the
constant external boattail angle of the single
sawtooth internal contour nozzle, separation
8
7) If internal contouring is required, the double 10. Cler, Daniel L. and Mason, Mary L.:
sawtooth nozzle planform provides better static Static Internal Performance of Three Spherical-
and aeropropulsive performance than the single Convergent-Flap Nozzles with In-Flight
sawtooth nozzle and if no internal contouring is Deployable Plugs, NASA TP-3565, U.S.
required the single sawtooth provides the Government and U .S . Government
highest static and aeropropulsive performance. Contractors Only.
REFERENCES 11. Asbury, Scott C. and Carlson, John R. :
Transonic Aeropropulsive Performance of
1. Herbst, W. B.: Future Fighter Advanced Exhaust Nozzles Designed for
Technologies, Journal of Aircraft, Vol. 17, Reduced Radar Cross-Section Signatures,
NO. 8, August 1980, pp. 561-566. NASA TP-3505, July 1995, U.S. Government
and U.S. Government Contractors Only.
2. Dollyhigh, Samuel M.; Foss, Willard E.:
The Impact of Technology on Fighter Aircraft 12. Propulsion Aerodynamics Branch Staff A
Requirements, SAE Technical Paper Series User's Guide to the Langley 16-Foot Transonic
85 1841, October 1985. Tunnel Complex, Revision 1, NASA TM-
102750, September 1990.
3. Mace, J., Doane, P.: Integrated Air
VehiclePropulsion Technology for a Multi- 13. Mercer, C. E.; et. al.: Data Reduction
Role Fighter - a MCAIR Perspective, AIAA Formulas for the 16-Foot Transonic Tunnel
Paper 90-2278, July 1990. NASA Langley Research Center, Revision 2,
NASA TM-107646, July 1992.
4. Brown, Squire L.; Snyder, Steven P.:
Exploration of Concepts for Multi-Role
Fighters ,AIAA Paper 90-2276, July 1990.
5. Herrick, Paul W.: Air-to-Ground Attack
Fighter Improvements Through Multi-Function
Nozzles, SAE Paper 901002, April 1990.
6. Results from NASA Langley Experimental
Studies of Multiaxis Thrust Vectoring Nozzles,
SAE Technical Paper 88 1481 , October 1988.
7. Brown, Alan C.: Low Observable
Propulsion Design, International Symposium
on Air Breathing Engines, loth, Nottingham,
England, Sept. 1-6, 1991, Proceedings. Vol. 1
(A91-56101 24-07), Washington DC, AIAA,
1991, pp 54-61.
8. MacLean, M.: Static Internal Performance
Tests of Single-Expansion Ramp Nozzle
Concepts Designed with LO Considerations,
AIAA Paper 93-2429, June 1993.
9. Cler, Daniel L.: Internal Performance of
Two Gimballed Nozzle Concepts with Multi-
Axis Thrust Vectoring and Reduced Observable
Design, NASA TP-3464, August 1994,
1TAW.S. Government and U.S. Government
Contractors Only.
9
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Orm
U
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FS 27.mO FS S3RI
Man: bruk
Figure 1. General arrangement of model and support system showing details of the single-engine propulsion
simulation system. All dimensions are in inches unless otherwise noted.
Single sawtooth, combined contour nozzle
Double sawtooth, external contour nozzle
Figure 2. Photographs of the single-engine propulsion simulation system installed in the tunnel with different nozzle
afterbodies installed
10