A potential flow computer program which has been in use for several years, was discussed. This program has been used primarily as a tool for flow-field analysis in support of test activities in transonic wind tunnels. Analyses have been made over a Mach number range from 0 to 0.9 for a variety of configurations from aircraft to wind tunnels, with excellent agreement between calculated flow fields and measured wind tunnel data. Analytical and experimental data for seven different flow analysis problems are presented in this paper

Palko, R. L.

English

NASA Technical Reports Server

POTENTIAL FLOW COMPUTER PROGRAM* 
Richard L. Palko 
ARO, Inc. 
A potential flow computer program has been in use at the Arnold Engineering De- 
velopment Center (AEDC) for several years. This program has been used primarily a s  
a tool for flow-field analysis in support of test activities in the transonic wind tunnels 
of the Propulsion Wind Tunnel Facility (PWT). Analyses have been made over a Mach k nurnber range from 0 to 0.9 for a variety of configurations from aircraft to wind tun- 
$ nels, with excellent agreement between calculated flow fields and measured wind tunnel 2 
%. data. Analytical and experimental data for seven different flow analysis problems a r e  
f- presented :n this paper. 
Iz f 
f 
-. . 
2. E INTRODUCTION 
e 
The AEDC rhree-Dimensional Potential Flow Computer Program (PFP) in the 
existing form was developed primarily as a result of the need to make calculations of 
the flow field in the vicinity of aircraft fuselages (typically at locations where aircraft 
inlets might be located). This need arose because of the support the theoretical flow- i 
field calculations could lend to a research program carried out at AEDC to simulate the I 
inlet flow fields in a wind-tunnel test of full/scale inlet/en.$ne systems (refs. 1 and 2). 
Much of the computing capability that the PFP  presently has resulted from these flow- 
field calculations which have a s  their primary variables the flow angularity (upwash and 
sidewash) over a y-z plane. After the initial solution of the velocity field for a given 
model attitude m d  Mach number is obtained, the upwash and sidewash can be deter- 
mined for any given point or  over any grid desired. A new solution is required for each 
model attitude o r  Mach number. In addition to computing the upwash and sidewash, tne 
PFP also computes the local Mach number, Cp, and flow streamlines. A computer 
I 
plotting program has been written to supplement the PFP, and computer plots can be 
*The research reported herein was conducted by the Arnold Engineering Develop- ; 
ment Center, Air Force Systems Command. Research results were obtained by per- 
sonnel of ARO, Inc., Contract Operator at AEDC. Further reproduction i s  authorized $ 
I to satisfy needs of the U. S. Government. 
. Q 
a 127 
:: i; 
' *  
;a ' i, 
https://ntrs.nasa.gov/search.jsp?R=19760021083 2020-03-20T02:31:31+00:00Z
obtained for most of the above parameters. A streamline can be traced from any point 
in the stream either upstream o r  downstream (or both). The capability to make a plot 
of the mathematical representation of the model geometry before running the complete 
program allows corrections to be made, if needed, with only a slight loss of computer 
time. A two-volume report (ref. 3) that details the program, modeling techniques, ap- 
plication, and verification has been published. 
This paper briefly outlines the PFP application to seven flow analysis problems in 
support of the transonic wind tunnels in the PWT a t  AEDC. 
SYMBOLS 
Values a r e  given in both SI and U. S. Customary Units. The measurements and 
calculations were made in U. S. Customary Units. 
angle of attack 
lift coefficient 
pressure cot:fficien t 
free-stream Mach number 
ratio of surface static pressure to free stream total pressure 
coordinate along tunnel axis, posi tive downstream 
horizontal coordinate, sign a s  indicated 
vertical coordinate, sign a s  indicated 
model angle of attack, deg, positive up 
model angle of yaw, deg, sign a s  indicated 
upwash, deg, positive up 
sidewash, deg, positive a s  indicated 
P F P  APPLICATION AND UTILIZATION 
The P F P  at AEDC has been used primarily a s  a cool for analysis of the flow in the 
far  field. (Far field refers  to a distance away from tht andysis model surface equal 
to, o r  greater than, the vortex spacing in the direction of flobr. ) The modeling tech- 
niques required for this type of analysis a re  presented in v c u n e  I1 of reference 3. 
However, work is underway to develop the modeling technique to allow accurate analysis 
of the surface pressure. Results of some of this continuing effort a re  reported in 
reference 4. All the flow problems presented here a re  of the far-field type. 
Flow Field Between TWO Hollow Circular Cylinders 
The analysis of these cylinders was part of a research program in which the objec- 
tive was to create flow fields by some auxiliary method to simulate the flow entering a 
full-scale inlet/engine at high mgles of attack and yaw. The device was to deflect (or 
induce) the flow upward a s  it passed between the inclined cylinders. A mathematical 
model of the cylinders is shown in figure 1. The last circumferential ray on each cyl- 
inder had trailing vortices that were trailed at an angle equal to one-half the cylinder 
pitch angle. A comparison between the theoretical and experimental flow angularity 
data is shown in figure 2. The theoretical results a r e  shown a s  lines of constant flow 
angle (both upwash and sidewash), and the solid symbols show the relative location of 
experimental data with the magnitude of the measured angles indicated. The Mach num- 
ber a t  which these data were taken was 0.9. It can be seen that the P F P  overestimated 
the flow inclination angles by approximately lo. 
Flow Field Around an Aircraft Fuselage 
Primary purpose of this analysis was to verify the results obtained from the PFP. 
Experimental flow-field data used for comparison with theory were available from wind- 
tunnel measurements made during the Tailor Mate test series. The objective of the 
windtunnel test was to determine the flow field (upwash and sidewash) at a typical en- 
gine inlet fuselage loca tion. The ma thema tical model of the fuselage configuration is 
shown in figure 3. The comparison between the predictions from the PFP  and the wind- 
tunnel data for a pitch angle of 25' and a free-stream Mach number of 0.9 a r e  shown in 
figure 4. Here again, excellent agreement was obtained. 
Flow Field Under a Fuselage-Wing Configuration 
The purpose of this analysis was also for program verification; again experimental 
data obtained du~ing  the Tailor Mate studies were used. The fuselage-wing configura- 
tion was analyzed to compute the flow field under the wing a t  the wing-fuselage junction. 
The computer math model used in the analysis is shown in figure 5. A comparison be- 
tween the upwash and sidewash predictions and the experimental data for a Mach num- 
ber of 0.9 and an angle of attack of lo0  is shown in figure 6. Analytical and experi- 
mental data trends show excellent agreement, although the predicted d a k  show some- 
what higher flow angularity gradients across the survey area than the measured data. 
Streamtube Entering M e t  Behind Wing 
The purpose of this analysis was to determine the origin of the streamtube entering 
the inlet in support of an inlet hot gas ingestion investigation. The comparison between 
experimental and theoretical data a r e  shown in figure 7. These data were taken during 
a store separation study in an effort to verify the accuracy of the PFP to predict the 
correct flow field above and behind the wing. Data were taken a t  a Mach number of 0.3 
with an angle of attack of 8O, and show excellent agreement between the experimental 
and theoretical values. The mathematical model and the predicted streamtube a r e  
shown in figure 8. The streamtube was determined by tracing streamlines from four 
locations beginning just upstream of the inlet and extending forw .rd to just upstream of 
the aircraft nose. A Mach number of 0.3 and an angle of attack of 8' were also used 
for the streamtube analysis. This mathematical model is the largest analyzed to date, 
with 1559 loop vortices and 20 horseshoe vortices, and required approximately 4 hours 
run time on the AEDC IBM 370/1b5 computer. 
Inle t/Engine in Crosswind 
This analysis was made in support of a crosswind experiment conducted during an 
inlet study in the AEDC 16-ft (4.88-m) Transonic Wind Tunnel (PWT-16T). The objec- 
tive of the analysis was to determine if  a 0.91-m-diameter (3-ft-diameter) crosswind 
simulator would adequately simulate the crosswind when used in conjunction with the in- 
let model, and to determine the position for the simulator to give best results. The 
theoretical analysis was made with the inlet/engine in an infinite crosswind, The math- 
ema tical model included only a portion of the experimental model a s  shown in figure 9. 
A computer plot of the mathematical model is shown in figure 10. The engine ducts 
were closed on the downstream end and a negative source was located near the rear 
center of each engine duct to produce the correct inlet mass flow when that particular 
engine was in operation. Streamlines were traced upstream from near the four corners 
of the inlet, for each engine in operation, to determine the flow pattern of the air- 
stream entering the inlet. By tracing the streamlines, a fan position was deter- 
mined that would influence the inlet flow for all engine power settings and cross- 
wind velocities required. A mica1  flow pattern for the analysis is shown in figure 
11 for a crosswind velocity of 20.57 m/sec (67.5 ft/sec) with both engines operat- 
ing. 
Pressure Distribution in PWT-16T Contraction Section 
The objective of this analysis was to determine the pressure distribution \ i
along the bottom and side walls of the PWT-16T contract i~n section. Pressure dis- Y 
tributions were needed for use in a theoretical boundary-layer analysis of the wind- 
tunnel nozzle to support a test-sec tion flow angularity study. Mathematical modeling 
used in the analysis ie  shown in figure 12. The flow in the test section area  was speci- 
fied to give the pressure ratio desired for Mach number 0.6. The analysis provided 
1 
1 
130 
streamline information at  a distance of 0.305 m (1 ft) from the walls, and the calculated 
Cp was converted to p/po. Following the calculation of the theoretical pressure dis- 
tribution, the pressure distribution was experimentally measured in the contraction 
section. A comparison between the theoretical and experimental pressure distribution 
is shown in figure 13, with excellent agreement indicated. 
Strut Effects Analysis 
The objective of this analysis was to determine the strut effect corrections to mea- 
sured force and moment data for a slender winged vehicle with a mid-strut mount. The 
vehicle wing was located just forward of the strut, For this analysis the upwash angle 
was determined with the PFP for the body alone (fig. 14) and the body with strut (fig. 15). 
An ir,cremental upwash angle was then determined a t  the wing location from these two 
sets of data. In this case the incremental values were negative because of the down flow 
around the strut. The incremental values along the wing location were averaged and the 
ACL correction calculated from the average angle-of-at tack change. A comparison of 
the calculated corrections and those measured with a subscale model are shown in fig- 
ure 16. Excellent agreement is shown in both the trend with Mach number and the ab- 
s o h  te values. 
CONCLUDING REMARKS 
The AEDC Potential Flow Program is used primarily a s  a tool for flow-field 
analysis in support of the test activities in the transonic wind tunnels of PWT. This 
paper has covered seven different problems that have utilized the PFP including both 
external and internal analysis, A11 but one of the examples have experimental data to 
verify the calculated flow fields, and all comparisons show excellent agreement. The 
PFP at AEDC has not been used as  a tool to obtain absolute values, but rather a s  a tool 
to predict and verify flow fields in sapport of the test activities. In addition to the 
problems presented, the PFP has been used to predict the flow angularity at the model 
resulting from sting and support systems, to predict the flow around various types of 
support systems, and many other general flow analysis pmblems directly connected 
with wind-tunnel testing. 
1 REFERENCES 
i i/ 
1. Palko, R. L. : Full-scale Inlet/Engine Testing at High Maneuvering Angles at 
Transonic Velocities. AIAA Paper No. 72-1026, Presented at the A M  Seventh 1 
Aerodynamic Testing Conference, Palo Alto, California, September 13-15, I . 
, - 1972. , 
i 
2. Palko, R. L. : A Method to Increase the Full-scale Inlet/~ngine System Testing 
Capability of the AEDC 16-F t Transonic Wind Tunnel. AEDC -TR-73-9 
4 (AD 762912). June 1973. 
3. Todd, Donald C. and Palko, Richard L. : The AEDC Three-Dimensional, Potential 
Flow Computer Program. Vol. I and 11. AEDC-TR 75-75, February 1976. 
4. Heltsley, Fred L.: Report on the Status of a Slotted Wind-Tunnel Wall Representa- 
tion Using the Vortex-Lattice Technique. Vortex-Lattice Utilization, NASA 
SP-405, 1976. (Paper no. 9 of this compilation.) 
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Figure 3.- Mathematical model of the fuselage configuration. 
W I ND-TUNNEL DATA 
ANALYTICAL DATA 
r s I DEWAS H 
Y Y 
Figure 4.- Comparison between analytical and wind-tunnel data for 
Mach number 0.9 at an angle of attack of 25O. 

EXPERIMENTAL THEORETICAL 
7 E -6.15' E * -6.540 
EXPERIMENTAL THEORETICAL 
Figure 7.- Compsrison between experimental and theoretical data for 
Mach number 0.3 at an angle of attack of 8O. 
Figure 8.- Flow streamtube entering the inlet at Mach number 0.3 
at an angle of attack of 8O. 
--..----- 
-a CROSSWIND 
-
Figure 9.- Sketch of the section of the test model 
wirh the mathematical model. 
duplicated 
TOP VIEW 
LEADING-EDGE NEGATIVE SOURCE 
POSITIVE 
INBOARD) 
- 
NEGATIVE 
(OUTBOARD1 
NEGATIVE 
SOURCE 
NEGATIVE 1 POSITIVE 
(OUTBOARD) (INBOARD) 
FRONT VIEW CROSSWIND- NEGATIVE 
LEAD I NG-EDGE LIP SOURCE 
INLET RAMP/ 
Figure 10.- Basic mathematical model for the inletlengine 
crosswind analysis. 
BOTH ENGINES ON 
VOL. FLOW . 4  74 m3lsec (167.5 ft31sa) PER ENGl NE 
CROSSWIND VELOCITY M 57 mlsa (67.5 ftlsa) 
CONFIGURATION 2 
POSITION 
TOP VIEW 
-4 0 1  SIDE VIEW 
0 ,  
-LO 
-2.0 
:I SIMUUIOR POSITION FRONT VIEW 
-3.0 
-5.0 -40  -3.0 -2.0 -LO 0 LO 
Y 
Figure 11. -  Flow streamlines for both engines on with a 
crosswind v e l o c i t y  a>f 20.57 m/sec (67.5  f t / s e c )  . 
SIMUUTOR POSITION 

SYM LOCATION 
- SOURCE 
- BOTTOM WALL 
;'iP I CALCULATED FOR 0.38 m -- - S I DE WALL PFP (1 ft) OFF WALL 
0 BOrrOM WALL EXPERIMENTAL DATA 
BOITOM WALL EXPERIMENTAL DATA (2175) 
A SIDE WALL EXPERIMENTAL DATA (2175) 
Figure 13.- Pressure distribution in PWT-16T for Mach number of 0.6. 
35 * 
ARO OM91 THRU AP R 24, 1975 
30' NOSTRUT b 4 . 7  A . 2  
25 ., 
Figsre 14.  - Upwash about c y l i n d r i c a l  model a t  &, = 0 . 7  and a = 2'. 
35 '  ARO 00649 THRU APR. 24, 1975 
30; AFT STRUT M m x Q 7  A . 2  
Figure 15.-  Upwash about cyl indrical  model with s trut  
a t  Mm = 0.7  and a = 2 O .  



Utilization of the AEDC three-dimensional potential flow computer program

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