Wind tunnel tests were conducted to establish a benchmark experimental data base for a genetic hypersonic aircraft vehicle. Comprehensive measurements were made at Mach 7 to give flow visualization, surface pressure, surface convective heat transfer, and flow field Pitot pressure for a delta platform all-body vehicle. The tests were conducted in the NASA/Ames 3.5-Foot Hypersonic Wind Tunnel at Reynolds numbers sufficient to give turbulent flow. Comparisons are made of the experimental results with computational solutions of the flow by an upwind parabolized Navier-Stokes code developed at Ames. Good agreement of experiment with solutions by the code is demonstrated

Cleary, Joseph W.

English

NASA Technical Reports Server

NASA CONTRACTOR REPORT
i
EXPERIMENTAL AEROTHERMODYNAMIC RESEARCH
OF HYPERSONIC AIRCRAFT
Joseph W. Cleary
Eloret Institute
3788 Fabian Way
Palo Alto, CA 94303
Prepared for
Ames Research Center
under Cooperative Agreement NCC2-416
W r"o
Nal_onal AeronautIcs and
Spaco Adm_ntslral_on
b,mes Research Center
Moffell F_etd Cahforn_a94035
)
lb. _, •
https://ntrs.nasa.gov/search.jsp?R=19900016638 2020-03-19T22:21:48+00:00Z
NASA CONTRACTOR REPORT
EXPERIMENTAL AEROTHERMODYNAMIC
OF HYPERSONIC AIRCRAFT
Joseph. W. Cleary
RESEARCH
.41
BP
t. Reooet No. 2, Govlr.ment Accesion No. 3. RecJp,ent'$ Catak_ I No.
,. i
4 Title _ _bzitle
Experimental Aerothermodynamic Research
of Hypersonic Aircraft
7 Autt_De(s)
Joseph W. Cleary
g l_efoeming f)_,uZio. Nlme IcKI Addk'ess
Eloret Institute
3788 Fabian Way
Palo Alto, CA 94303
|;., ¢,,,o.,o,i,_g._.cv N.m..,_ Aaron,
National Aeronautics and Space
Administration, Washington, D.C. 20456
15 _p_e_ntarv Notes
5 Report Dete
23 August 1990
6. Performing OrllZ_zatio- Code
8. Perfoqming Orgz.izltio. Reoort No.
10. wowk U_t No.
11. Contract or Gre.! No.
NCC2-416
o3-01-86 - 07- 31-9o
14. Sponr, orir_ Agm_cV
Point of Contact: William K. Lockman
c/o 230-2 NASA Ames Research Center, Moffett Field, CA 94035
16 Abstract
Wind-tunnel tests were conducted to establish a
benchmark experimental data base for a generic
hypersonic aircraft vehicle. Comprehensive
measurements were made at a Mach mmd_,er of 7 to give
flow visualization, surface pressure, surface convective
heat transfer, and flow-field pitot pressure for a delta
planform all-body vehicle. The tests were conducted in
the NASA/Ames 3.5-Foot Hypersonic Wield Tunnel at
Reynolds numbers sufficient to give turbulent flow.
Comparisons are made of the experimental results with
computational solutions of the flow by an upwind
paraboliz_d Navier-Stokes code developed at Ames. Good
agreement of experiment with solutions by the code is
demonslrated.
17 Key Word_ (SucJIq_ted by Author(ll_
I IR O4,tr,b_It_(x_Statement
Wind Tunrlel Test, s
Hyperson]c Aircrsft
Flow Visual_zatiou
Surface Convective }{eat Transfe,
19 _'_ih, C_lwif iof th,I ,e_t) 20 _itV C_.wf (of this I_m[
Unc lass] fied Unc ]assJ fied
Unclassified, Unlimited
'2! No of P0g_ 72 P_Ice"
• to, ..aI_ by th_ F.lt*o-al Tlchr.cM l.fo_.'zetio_ S41rv*Ce. S(_'_n,gfield. V,rg..,l 22181
1
"1
EXPERIMENTAL AEROTHERMODYNAMIC RESEARCH
OF HYPERSONIC AIRCRAFT
Final Technical Report
for the period
March I, 1986 - July 31, 1990
Submitted to
National Aeronautics and Space Administration
Ames Research Center
Moffett Field, California 98035
Aerothermodynamics Branch
George S. Deiwert, Chief
William K. Lockman, Technical Monitor
NASA-Cooperatlve Agreement NCC2-416
Prepared by
ELORET INSTITUTE
1178 Maraschino Drive
Sunnyvale, CA 94087
Phone: (408) 730-8422 and (415) 493--4710
Telefx: 408 730-8422
K. Heinemann, President and Grant Administrator
Joseph W. Cleary, Principal Investigator
23 August, 1990
The need for long-range, high-speed modes of transportation has
generated significant interest in high-performance hypersonic aircraft. A
concept known as the National Aero-Space Plane, (NASP) envisions a lifting
air-breathing vehicle that can operate across the Mach number range up to
orbital velocity with the added capability of conventional aircraft-type
takeoffs and landings. A concerted theoretical and experimental program
is underway to give proper definition to the design of NASP. The
theoretical effort entails numerical solutions of the Navier-Stokes
equations for perfect and real-gas flows using state-of-the-art and
advanced computer codes. Because of their versatility, the codes are a
valuable asset in all stages of the design process. However, comparisons
with experimental data are necessary to validate the codes and to give
confidence in their applicability to practical designs.
A program has been initiated at the Ames Research Center to provide
these much-needed experimental aerodynamic data. Comprehensive
wind-tunnel tests were made in the 3.5-Foot Hypersonic Wiad Tunnel of a
generic hypersonic airplane model of a type envisioned for NASP. Briefly,
the model is an all-body concept having a delta planform with 75 °
sweepback of the leading edges, elliptical cross sections, and a sharp
trailing edge. Horizontal and vertical control surfaces can be appended to
the outboard tips at the trailing edge. Previous research has shown that
this configuration is stable and controllable over the oper_tional rang': of
Mach numbers. The present tests were conducted without control
surfaces.
maP'
The present Final Technical Report covers the period from the beginning of
the grant on March l, 1986, to completion on May 31, 1990. The protocol
for conducting the research was established at the beginning of the grant
(Cooperative Agreement No. NCC2-416). Under terms of the agreement,
the Principal Investigator, Joseph W. Cleary, and the Technical Monitor,
William K. Lockman, collaborated on conducting the experimental
aerodynamic research and the wind-tunnel test programs.
A major endeavor of the experimental program was to provide
aerodynamic data that would be useful for computational code validation.
A selected series of tests were made to give surface pressure, convective
heating, shock-wave shape, surface skin-friction lines, and flow-field
properties of the shock layer. Results were presented at the Third, Fourth,
Sixth, and Seventh National Aero-Space Plane Technology Symposia. A list
of the papers that were presented and will be published is given as an
addendum. The experimental data were shown to be in good agreement
with an upwind parabolized Navier-Stokes (UPS) code. Copies of NASP
papers have been submitted previously as progress reports.
As a Principal Investigator, the research covered four major work areas:
(1) Stress analysis and detail designs; (2) Aerothermodynamic study tasks:
(3) Wind-tunnel test operations; and (4) Data analysis and report writing.
Although the Principal Investigator collaborated in all areas, he was given
primary responsibility for certain research aspects of the investigation.
Participation in each of the four work areas will be discussed briefly. Then
-2-
a more detailed report will be given from which research aspects of
primary responsibility will be apparent.
The model was designed and built by Convair Corporation specifically for
wind-tunnel tests at Ames. To assure integrity of the model and sting
support, a complete review was made of the Convair stress analysis for the
high-temperzture and dynamic-pressure conditions encountered in the
3.5-Foot Hypersonic Wind Tunnel. Loads were estimated from previous
wind-tunnel force tests of a smaller model and factors of safety were
evaluated.
Detailed designs and stress analyses were completed of five related
components essential to the wind-tunnel test program. A new circulating-
water, constant-temperature heater was designed to accommodate six
electronic pressure scanners for surface pressure measurements.
Commercially available heaters were too large to fit in the restricted
internal volume of the model.
For aerodynamic convective heating tests of the model, a new thin-skin
heat-transfer cover plate was designed. A unique feature of the design
was to attach a thin sheet of stainless steel of constant thickness to the
cover-plate frame, thereby eliminating the need for expensive detailed
machining of thin-skin areas for the thermocouples.
Be,'ause the wind-tunnel test section is not equipped for flow-field
surveys, a new survey platform was designed. The platform can be used
to position pitot-pressure rakes (or other type probes) within the (x, y, z)
it,
coordinate system of the model flow field. A new flow-field and
boundary-layer rake was also designed. The rake consists of multiple
stainless steel probes closely spaced in the boundary-layer region and
more widely spaced in the outer shock layer. The probes were flattened to
give better definition to the boundary-layer profiles. In addition to the
foregoing components, a 5 ° sideslip wedge was designed to yaw the model.
All of the foregoing components were built and used in the all-body wind-
tunnel test program. Furthermore, most of these components are believed
useable for other future wind-tunnel test programs.
In the area of aerothermodynamlc study tasks, equations were derived for
hypersonic Newtonian flow over the elliptical-forebody and -afterbody of
the all-body model. The equations were extended and a computer code
was developed (jointly with the Technical Monitor) to give tangent-cone
solutions of the flow. The solutions show the effects of angle of attack,
angle of sideslip, Mach number, and specific-heat ratio on spanwise
pressure distributions. Comparisons were made with experimental results
and UPS computational solutions of the flow in a previously mentioned
NASP document.
Before beginning the pressure distribution tests, a brief analysis was made
of unpublished pressure distributions from tests of the model in the
Unitary Plan Wind-Tunnel Facilities. The tests were made at Mach
numbers of 2.5, 3.0, and 3.4. The data were analyzed to show the effects
of angle of attack, nose bluntness, Mach number, and forebody/afterbody
juncture on windward and leeward spanwise-pressure distributions. The
//llmlr
_
results provided a basis for analyzing these same effects at hypersonic
Mach numbers.
A simplified inviscid model of the forebody and afterbody flow-field was
developed to demonstrate flow-field structure on model centerline.
Because the forebody of the model is an elliptical cone, the bow-shock
wave is basically conical as is the entire forebody shock layer except for
the boundary layer. For such flows, an approximation using the known
wave angle in the angle-of-attack plane was previously published in
NASA TN D-5951 (by the Principal Investigator). The method yields
accurate inviscid flow-field estimates for various conical bodies. The
afterbody flow-field for the all-body model is quite complex; it is
composed of an outer conical region, a central expansion region, and a
viscoas region at the wall. Using the above approximate method, a
simplified inviscid model of this flow-field was developed and compared
with experiment. The main features of the centerline flow-field structure
compare favorably with experiment for a range of angles of attack. In the
work area of wind-tunnel operations, the Principal Investigator assumed a
hands on position for the numerous tasks involved. Briefly, the wind-
tunnel tests required model 0reparation and installation, data acquisition,
and data reduction. Calibration and instrumentation checks were made
before and after a wind-tunnel run to assure valid data. Each of the
various tes _ spanned a duration of one to two (or more) months.
The final work area, data analysis, and report writing, can be time
consuming. Large amounts of data were produced during the wind-tunnel
test program. Test results were published in a timely manner in the four
_
previously mentioned NASP Technology Symposia papers. Moreover, an
AIAA paper entitled "Experimental and Computational Surface and Flow-
Field Results for an All-Body Hypersonic Aircraft" will be given in at the
AIAA 8th Applied Aerodynamics Conference in Aug. 1990. The paper
gives detailed comparisons of flow visualization, surface pressure, heating,
and shock-layer pitot-pressure measurements with Navier-Stokes
solutions of the flow from the Ames UPS code. Generally good agreement
is shown by the comparisons. The experimental results fJ-om this ._tudy
are invaluable for the validation and or calibration of the UPS code and
other codes being developed at Ames and by other organizations.
With regard to specific aspects of the research program, the Principal
Investigator was given primary responsibility for the flow-field area of
research. Detailed pitot-pressure measurements were made of the
afterbody flow-field of the all-body model. It is noteworthy that turbulent
flow over the afterbody is quite complex and affords a realistic test case
for the capabilities of the UPS code. In addition to pitot-pressure,
measurements of bow shock-wave shape and surface oil-flow patterns
were essential to the flow-field analysis. Tests were completed for the
sharp-nose model. Effort was expended in preparation for tests to
determine the effects of nose blunt0ess. However, wind-tunnel scheduling
conflicts precluded tests of the effects of bluntness.
II,"
h
At present, a final report is being prepared and near completion on the all-
body sharp-nose flow-field data. Completed results of the afteihody pitot-
pressure flow-field tests will be given in tabulated form for centerline and
outboard (lateral) survey stations. In addition, the report will have figures
_
that show the effects of angle of attack and longitudinal station on pitot-
pressure distributions. Since the boundary layer is quite thick and the
rake probes are closely spaced, detailed boundary-layer profiles are shown
for complex turbulent flows.
Brief excerpts from the final report are given now to show the more
interesting aspects of the flow-field investigation. In figure 1, the model is
shown mounted on the sting support in the wind-tunnel. The model is
instrumented with centerline and outboard cantilever mounted rakes
which can be located at various longitudinal positions. A shadowgraph of
the flow-field is shown in figure 2 for ot = 5 ° . Pertinent features of the flow
on the windward side are the bow-shock wave, an expansion of the flow at
the forebody/afterbody juncture, and the turbulent boundary layer on the
forebody. A thick turbulent boundary layer on the afterbody is barely
visible. The modei is shown with the alternate pitot-pressure rake that
was used in preliminary tests. The boundary-layer and flew-field rakes
shown in figure 1 were used for the main investigation.
A comparison of centerline pitot-pressure data with an inviscid model of
the flow-field is shown in figure 3. This figure demonstrates the structure
of the afterbody flow-field. Since the bow-shock wave is straight, the
flow-field has a conical-flow region adjacent to the wave. The expansion of
the flow at the forebody/afterbody juncture is similar to the Pral,dtl-
Meyer expansion of the flow model. The expansion turns the flow parallel
to the afterbody boundary layer, which is represented as a two-
dimensional flow (planar) region by the flow-field model.
7
iiiF_ r--
Tt_._, effects of x/L (rake station) on pitot-pressure distributions for ix = 0 °
are shown in figure 4. At each station, the distributions are composed of
conical, expansion, and viscous regions ,af the flow-field as previously
discussed. The edge of the boundary layer (viscous region) was estimated
from shadowgraphs. It can be seen that the relative thicknesses of the
flow-field regions are highly dependent on x/L. Just after the
forebody/afterbody juncture (x/L -- 0.70), the expansion and the thick
turbulent 0oundary layer are closely coupled. Further aft (x/L _>0.80), it is
hypothesized that the expansion terminates external to the boundary
layer. The somewhat abrupt change in pitot-pressure near the boundary
layer edge (figure 4) at x/L = 0.90 is believed to be a manifestation of the
phenomenon.
The effects of varying angle of attack on pitot-pressure distributions are
shown in figure 5. Similar to ix = 0 °, ti?_ distributions on the windward side
are composed of conical, expansion, and viscous regions. Increasing angle
of attack has a significant effect on the jur cture expansion because the
forebody Mach number, which modulates the expansion, is reduced. On
the leeward side, the distributions for ix = -5 ° and -10 ° appear related as a
family of curves to those on the windward side.
Pitot-pressure distributions at the outboard station, y/L = 0.167, are
shown in figure 6 for ct = 0 °. Previous results on centerline are shown for
comparison. The outboard distributions, similar to centerline, have conical.
expansion, and viscous regions of the flow-field. However, for the same
distance from the surface, pitot-pressure is greater outboard ti:an on
centerline. In summary, a comprehensive wina-tunnel investigation was
_
conducted for a generic airplane model in support of NASP technology
development. Tests were made to give surface pressure, heating, shock-
wave shape, surface skin-friction lines, and flow-field pitot pressure. The
experimental data were found to be in good agreement with the theoretical
UPS code. Results were presented at NASP Symposia and will be given at
the AIAA Applied Aerodynamics Conference.
A list of the publications from this investigation is presented here.
• Lockman, William K.; Lawrene, e, Scott L,; and Cleary, Joseph W.: Flow-
Visualization Results for an Ali-Body Hypersonic Aircraft. Third
National Aero-Space Plane Technology Symposium, Paper No. 6, June
1987.
.
,
,
Lockman, William K.; Cleary, Joseph W., and Lawrence, Scott L.: Flow
Visualization and Pressure Distributions for an All-Body Hypersonic
Aircraft. Fourth National Aero-Space Plane Technology Symposium,
Paper No. 53, Feb. 1988.
Lockman, William K.; Cleary, Joseph W.;and Lawrence, Scott L.:
Experimental and Computational Flow-Field Results for an All-Body
Hypersonic Aircraft. Sixth National Aero-Space Plane Technology
Symposium, Paper No. 37, April 1989.
Lockman, William K.; Lawrence, Scott L.; and Cleary, Joseph W.: Surface
Heat Transfer and Pitot-Pressure Survey Results for an All-Body
Hypersonic Aircraft. Seventh National Aero-Space Plane Technology
Symposium, Papel No. 6, Oct. 1980
_
5. Lockman, William K.; Lawrence, Scott L. and Cleary, Joseph W.:
Experimental and Computational Surface and Flow-Field Results for an
All-Body Hypersonic Aircraft. AIAA Paper No. 90-3067, Aug. 1990.
, Cleary, Joseph W. and Lockman, William K.: Experimental Flow-Field
Results for an All-Body Hypersonic Aircraft Model. Proposed NASA
TM, 1990.
, Cleary, Joseph W. and Lockman, William K.: Newtonian and Tangent-
Cone Approximations of the Spanwise Pressure Distributions for an All-
Body Hypersonic Aircraft. Proposed NASA TM, 1990.
D
10
Fig. 1 Photograph of pitot-pressure survey setup with rakes on
centerline and outboard.
!
Fig. 2 Shadowgraph of flow field: c_ = 5 °, Moo = 7.4, Reo. L = 15 x 10 6.
ORIGINAL PAGE IS
OF POOP QUALITY
N,<:
0.15
(./1
0.10
_L_ 0.05
E
Z 0.00
Fig. 3
_Nw 0.20
0.16
o=
n=
0.10
_,, 0.06
N
,<
=E
E
0 o.oo
Z
EXPERIMENT; a -- I_DMO0_
0
, , : "_--P"_,
o I / S.OCK wave
4 _i : •
..°1 ,/._o,,_,L
0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0
PITOT-PRESSURE RATIO, Pt2/(Pt2)oo
Comparison of centerline pitot pressure with inviscid model of
flow; Moo = 7.4.
(3
0
0
t
a •
I I I
0.0 0.6 1.0 1.8 2.0
Fig. 4
PITOT-PRESSURE RATIO. Pt.2/(Pt.2 ).
Effect of rake station on pitot-pre'_sure distributions; 0t = 0 °,
Moo = 7.4, ReooL = 15 x 10 6 .
v . k _t,=-,,.,_e
m O20N
0.16
O.10
Z
a
a o.os
tU
--I
<
_" 0.00
0
Z
EXPERlUENT; a _ CODE
-18" /_ 111" _ TURBULENT
LEEWARD
-10" 0 10"
O M. vZ=
¢
A
0
0 o WINDWARD
O
0.0
Fig. 5
0.6 1.0 1.6 2.0 2.6 3.0 3.6 4.0
PITOT-PRESSURE RATIO. Pt,2/(Pt,2).=
Effect of angle of attack on pitot-pressure distributions; Mo. =
7.4, Reoo,L = 15 x 10 6.
N= (U) CENTERUNE, y/L : 0 (U) OUTBOARD, y/L = 0.167
I o
_XPERIMENT; IlL EXPERIMENT; x/L/ <>o:s _ "_ I I <>0.=
¢/_ 0.16 _- _o.=0
/---,-_ _""--- t --------_0 q I
i 0.10
0.06
B
0.00 _ 1 a
rr 0.0 0.6 1.0 1.6 0 0.0 0.6 1.0 1.6
o
z PITOT-PRESSURE RATIO, Pt,2/(Pt,2)oo PITOT-PRESSURE RATIO, Pt,2/(Pt_.)oo
Fig. 6 Comparison of pitot-pressure distribution_ at centerline and
outboard stations; Mo. = 7.4, Reo.,L = 15 x 10 6 .



Experimental aerothermodynamic research of hypersonic aircraft

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