X-33 Experimental Aeroheating at Mach 6 Using Phosphor Thermography

17 Pages·1999·5.5 MB·English

by  Horvath, Thomas J.

AIAA 99-3558
X-33 Experimental Aeroheating at Mach 6
Using Phosphor Thermography
T. J. Horvath, S. A. Berry, B. R. Hollis,
D. S. Liechty, H. H. Hamilton II, and N. R. Merski
NASA Langley Research Center
Hampton, Virginia 23681
33rd Thermophysics Conference
28 June - 1 July, 1999 / Norfolk, VA
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics
1801 Alexander Bell Drive, Suite 500, Reston, VA 22091
X-33 EXPERIMENTAL AEROHEATING AT MACH 6 USING PHOSPHOR
THERMOGRAPHY
Thomas J. Horvath*, Scott A. Berry*, Brian R. Hollis*(cid:160), Derek S. Liechty, H. Harris Hamilton II*(cid:160),
and N. Ronald Merski*(cid:160)
Abstract
The goal of the NASA Reusable Launch Vehicle (RLV) technology program is to mature and demonstrate
essential, cost effective technologies for next generation launch systems. The X-33 flight vehicle presently being
developed by Lockheed Martin is an experimental Single Stage to Orbit (SSTO) demonstrator that seeks to validate
critical technologies and insure applicability to a full scale RLV. As with the design of any hypersonic vehicle, the
aeroheating environment is an important issue and one of the key technologies being demonstrated on X-33 is an
advanced metallic Thermal Protection System (TPS). As part of the development of this TPS system, the X-33
aeroheating environment is being defined through conceptual analysis, ground based testing, and computational fluid
dynamics. This report provides an overview of the hypersonic aeroheating wind tunnel program conducted at the
NASA Langley Research Center in support of the ground based testing activities. Global surface heat transfer
images, surface streamline patterns, and shock shapes were measured on 0.013 scale (10-in.) ceramic models of the
proposed X-33 configuration in Mach 6 air. The test parametrics include angles of attack from -5 to 40 degs, unit
Reynolds numbers from 1x106 to 8x106/ft, and body flap deflections of 0, 10, and 20 deg. Experimental and
computational results indicate the presence of shock/shock interactions that produced localized heating on the
deflected flaps and boundary layer transition on the canted fins. Comparisons of the experimental data to laminar and
turbulent predictions were performed. Laminar windward heating data from the wind tunnel was extrapolated to flight
surface temperatures and generally compared to within 50 deg F of flight prediction along the centerline. When
coupled with the phosphor technique, this rapid extrapolation method would serve as an invaluable TPS design tool.
Nomenclature
¥ free-stream conditions
h heat transfer coeff. (lbm/ft2-sec), q˙/(H - H ) L reference length
aw w
where H = H b reference span
aw t,2
H enthalpy (BTU/lbm) n model nose
M Mach number t, 1 reservoir conditions
P pressure, psia t, 2 stagnation conditions behind normal shock
q˙ heat transfer rate (BTU/ft2-sec) w wall
r radius (in.) Introduction
t time (sec) The Access to Space Study1 conducted by NASA
Re unit Reynolds number (1/ft) recommended the development of a fully reusable
T temperature (¡R) launch vehicle (RLV)2,3,4 to provide a next-generation
u velocity (ft/sec) launch capability at greatly reduced cost. This led to
x axial distance from origin (in.) the RLV/X-33 technology program, an industry-led
y lateral distance from origin (in.) effort in partnership with NASA. The primary goal
z vertical distance from origin (in.) of the RLV/X-33 technology program is to enable
a angle of attack (deg) significant reductions in the cost of access to space.
d control surface deflection (deg) The full-scale RLV system must be lightweight
Subscripts enough to achieve orbit and deliver a payload in a
cost-effective manner. As part of the RLV program,
aw adiabatic wall
the X-33 is envisioned as a sub-scale rocket
BF body flap
technology demonstrator for a Single-Stage-To-Orbit
FR Fay-Riddell, stagnation-point, reference heating
(SSTO) RLV. The proposed X-33 vehicle is intended
condition
to demonstrate key design and operational aspects of a
commercially viable system SSTO RLV rocket.
Following a Phase I industry competition for X-33,
*
Aerospace Technologist, Aerothermodynamics Branch, Aero- Lockheed Martin(cid:213)s lifting body concept was selected
and Gas-Dynamics Division, NASA Langley Research Center, by NASA and Lockheed was awarded a Phase II
Hampton, VA 23681.
contract (July 1996) to pursue construction of the
(cid:160) Member, AIAA flight vehicle. Lockheed(cid:213)s current lifting body design,
Copyright (cid:211)1999 by the American Institute of Aeronautics and Fig. 1, shown in the dimensioned sketch at model
Astronautics, Inc. No copyright is asserted in the United States
scale, represents a one-half scale RLV and
under Title 17, U.S. Code. The U.S. Government has a royalty-
free license to exercise all rights under the copyright claimed incorporates symmetric canted fins, twin vertical
herein for government purposes. All other rights are reserved by tails, two body flaps located at the rear of the fuselage
the copyright owner.
1
for aerodynamic control, and is powered by a linear the LaRC 20-Inch Mach 6 Tunnel with an
aero-spike engine5. appropriately sized model. A Re range of 0.4 to 6.7
L
million is achieved in this tunnel with a 10-in model
As part of the Industry/Government partnership,
length. The current flight trajectory from Edwards
NASA Langley Research Center (LaRC) has been
AFB into Michael AFB has been modified from the
tasked with providing surface heating data, including
preliminary trajectory (9d-3) shown and would place
transition criteria, to Lockheed Martin in support of
vehicle peak heating during ascent near Mach 9 at
X-33 aerothermodynamic development and design.
a=13 deg and during descent near Mach 7 at a=32
To meet the objectives outlined in the tasks, a
deg.
synergistic experimental/computational approach was
utilized. Results from early Phase II wind tunnel Test techniques that were utilized during these
heating measurements were compared to laminar and tests include thermographic phosphors, which
turbulent CFD computations6. Flight peak heating provides global surface heating images; oil-flow,
rates over the X-33 were then predicted with both an which provides surface streamline information; and
(cid:210)engineering(cid:211) code and a Navier-Stokes solver. The schlieren, which provides shock system details.
early data set was also used to formulate and support Parametrics included in these tests at Mach 6 were a
the use of a Re /M criteria to predict transition onset large range of angle of attack to cover ascent and
for X-337. Sqincee the time of these publications, descent conditions (-5<a<40 deg.), unit Reynolds
number (Re/ft between 1 and 8 million), and body
additional heating tests have been completed which
flap deflections (d , = 0, 10, and 20 deg). Emphasis
supplemented the original database and accommodated BF
will be placed on the body flap surface heating
design changes to the vehicle outer mold lines. Key
augmentation due to deflection. Sample
Phase II experimental and computational aeroheating
comparisons with CFD predictions (details provided,
results are presented in this reference and in two
companion papers by Berry et al8 and Hollis9 et al. ref. 9) are included and are used to assess the state of
the windward surface boundary layer. Extrapolation
An extensive testing effort was made to examine the
and comparison of laminar wind tunnel heating
effects of both discrete roughness and distributed
measurement to flight surface temperature predictions
roughness (in the form of a simulated array of
are made.
thermally bowed metallic TPS panels) on transition.
The sensitivity of X-33 boundary layer transition to X-33 Body Shape Evolution and
this (cid:210)wavy wall(cid:211) and the validity of the discrete Description
roughness defined Req/Me transition criteria off The design of the Lockheed Martin X-33 lifting
vehicle centerline have been investigated8. Additional body that emerged from the Phase I competition drew
Navier Stokes computations were performed and upon a synthesis of work performed by the U.S.
detailed comparisons with the more recent wind government and industry over the last few decades11,12.
tunnel data have been made to validate computational The X-33 body resembled a blunted slab with a delta
techniques used in predicting complex three- shaped planform. Two body flaps trailed from the
dimensional flow-fields9. The experimental results lower surface outboard of the base mounted linear
were obtained in the LaRC Aerothermodynamic aero-spike engine. These flaps, along with a pair of
Facilities Complex10. Over 1100 tunnel runs from canted fins and a single vertical tail would be
sixteen different entries in two facilities have been employed for aerodynamic control. Aerodynamic
completed since Aug 1996. Table 1 provides a list of wind tunnel tests conducted in government labs
all the wind tunnel tests that have been completed to (NASA LaRC13,14, NASA MSFC12), and commercial
date in support of X-33 aeroheating since the Phase II facilities continued after Phase I down-select with the
down-select. goal of optimizing aerodynamic performance across
the speed range. Although the basic lifting body
The purpose of this paper is to update and
shape has been maintained throughout the
present an overview of the most current experimental
development, changes to the vehicle have occurred as
measurements to characterize the X-33 windward and
the aero database matured. In the past, aeroheating
leeward aeroheating environments, and will focus on
information has significantly lagged behind
the heating associated with a nominally smooth
aerodynamics due to model and instrumentation
surface. The wind tunnel data in this report was
complexities associated with aerothermodynamic
obtained from the NASA LaRC 20-Inch Mach 6 Air
testing. The timeframe of the X-33 program coupled
Tunnel. In terms of Mach and Reynolds number
with the recent development of the two-color global
simulation, an early X-33 flight trajectory (9d-3)
phosphor thermography technique15-17 presented
considered by Lockheed Martin (Fig. 2) indicates that
NASA Langley the first opportunity to conduct an
the 63-ft long flight vehicle would experience a
aerothermodynamic screening/trade in parallel with
length Reynolds number (Re ), between 4 and 8
L aerodynamic development.
million at a freestream Mach number of 6. This
range of length Reynolds number can be produced in Phase II aerodynamic optimization studies
2
produced a series of X-33 configurations with outer construct molds from which the cast ceramic model
mold line (OML) changes of sufficient significance to configurations were made. A magnesia ceramic was
warrant several aeroheating assessment/trade studies. used to backfill the ceramic shells, thus providing
The configuration that emerged from the Phase I strength and support to the sting support structure.
down-select and was used in the early part of Phase II A photograph of three 0.0132 scale (10-in.)
heating tests was referred to in the aeroheating Rev-F model configurations with the various body
community as the D-loft (603B1001D) - with the flap deflections are shown in Fig.3. Typically, two
industry configuration designation18 shown in casts of each configuration were made; the primary
parenthesis. The subsequent configuration OML, the being immediately prepared for testing and the back-
F loft Rev-C (604B0002C), differed from the D-loft up shell held in reserve, in case of problems with the
in that it had twin vertical tails and incorporated some primary. In order to obtain accurate heat transfer data
modifications which changed the nose shape slightly with the phosphor technique, the models are cast with
and the base region (in the vicinity of the engine). a material with low thermal diffusivity and well
The nosecap changes were made to simplify the defined, uniform, isotropic thermal properties. The
construction of the metallic TPS panels. phosphor coatings typically do not require
Aerodynamic improvements to this vehicle shape refurbishment between runs in the wind tunnel and
resulted in F loft Rev-F (604B0002F), as shown in have been measured to be approximately 0.001 inches
Fig.1. It has the same forebody shape as F loft Rev- thick. Details concerning the model fabrication
C, but the dihedral of the 0012-64 airfoil canted fins technique and phosphor coating can be found in
was lowered from 37-deg to 20-deg (to improve pitch- refs.19 and 20. Fiducial marks were placed on the
trim characteristics across the speed range). At the model surface to assist in determining spatial
same time, the fin incidence was lowered from —6.57 locations accurately.
to —8.57 deg. Finally, the size of the body flaps and
Once the phosphor testing was completed, the
vertical tails was increased (to improve trim
untested backup models were prepared (spray-coated
characteristics and low-speed lateral-directional
and kiln fired with a thin black glazing) for use as oil-
stability, respectively). The last OML iteration that
flow and schlieren models.
has undergone heating tests at LaRC is known as F
loft Rev-G (604B0002G). Small protrusions on the Facility D escription
leeward surface near the vertical tails were added to The X-33 aeroheating test series was conducted
accommodate internal structural changes and the in two hypersonic blowdown facilities that are part of
canted fin body fillet was modified. While other more the LaRC Aerothermodynamic Facilities Complex.
recent systematic configuration changes have resulted Detailed descriptions of the facilities within the
in minor vehicle OML modifications, these changes complex can be found in ref. 10. A brief description
were not significant enough to warrant construction of the 20-Inch Mach 6 Tunnel follows, since a
of new models and additional wind tunnel testing. majority of the tests were in this facility.
Experimental Methods The 20-Inch Mach˚6 Tunnel uses heated, dried,
Models and filtered air as the test gas. Typical operating
conditions for the tunnel are: stagnation pressures
A majority of the cast ceramic aeroheating
ranging from 30 to 500 psia; stagnation temperatures
models are a 10-inch long 0.0132-scale representation
from 760-deg to 1000-degR; and freestream unit
of the proposed 63-ft long (from nose to the end of
Reynolds numbers from 0.5 to 8 million per foot. A
the aerospike engine base) X-33 F loft Rev-F
two-dimensional, contoured nozzle is used to provide
(604B0002F) flight vehicle. In addition to
nominal freestream Mach numbers from 5.8 to 6.1.
configuration assessment of the X-33 OML revisions
The test section is 20.5 by 20 inches; the nozzle
(L=10-in.), an attempt to qualitatively characterize the
throat is 0.399 by 20.5-inch. A bottom-mounted
influence of model size on tunnel partial blockage
model injection system can insert models from a
effects was made. This was accomplished by limited
sheltered position to the tunnel centerline in less than
testing of two smaller scale Rev F models (L=5.7-in.
0.5-sec. Run times up to 15 minutes are possible
and L=6.5-in.).
with this facility, although for the current heat
Over 70 ceramic models were fabricated in transfer and flow visualization tests, the model
support of the LaRC X-33 aerothermodynamic residence time in the flow is limited to only a few
program, all of which share a common construction seconds. Flow conditions were determined from the
technique. A rapid prototyping technique was first measured reservoir pressure and temperature and the
used to build a resin stereolithography (SLA) model measured pitot pressure at the test section.
with various, detachable body flaps on both the port
Test C onditions a nd S etup
and starboard region of the base of the vehicle. The
SLA resin model was then assembled with the desired Nominal reservoir and corresponding free
control surface settings and served as a pattern to stream flow conditions for the 20-Inch Mach 6
3
Tunnel is presented in ref. 20. The freestream and red intensity images to global temperature
properties were determined from the measured mappings. With temperature images acquired at
reservoir pressure and temperature and the measured different times in a wind tunnel run, global heat
pitot pressure at the test section. The reservoir transfer images are computed assuming one-
pressure pt,1 was measured with two silicon sensors dimensional semi-infinite heat conduction. The
having a full scale rating of 150 psia or 500 psia, primary advantage of the phosphor technique is the
depending on the operating condition of the tunnel. global resolution of the quantitative heat transfer data.
The reservoir temperature Tt,1 was measured with Such data can be used to identify the heating footprint
of complex, three-dimensional flow phenomena (e.g.,
two iron-constantan thermocouples inserted through
transition fronts, turbulent wedges, boundary layer
the wall of the settling chamber. Test section wall
vortices, etc.) that are extremely difficult to resolve
static and pitot pressures were monitored and
by discrete measurement techniques. Because models
compared to tunnel empty conditions to assess if
are fabricated and instrumented more rapidly and
model blockage effects existed. No significant
economically, global phosphor thermography has
differences in pitot pressure were measured and it was
largely replaced discrete heating instrumentation in
concluded that significant blockage did not exist (up
Langley’s AFC.
to 40 deg angle of attack, the limit of the tests). The
ratio of projected model frontal area to tunnel cross Flow visualization techniques, in the form of
sectional area for the present test was 0.1 (@L=10- schlieren and oil-flow, were used to complement the
in., a=40 deg). surface heating tests. The LaRC 20-Inch Mach 6
All models were supported by a base mounted Tunnel is equipped with a pulsed white-light, Z-
cylindrical sting with the exception of two models pattern, single-pass schlieren system with a field of
(L=6.5-in and 5.7-in..) which were blade supported view encompassing the entire 20-in test core. Images
from the leeside to assess support interference effects. were recorded on a high-resolution digital camera,
Details of the X-33 ceramic heat transfer model
enhanced with commercial software and electronically
installation in the NASA LaRC 20-Inch Mach 6 Air
placed into this report. Surface streamline patterns
Tunnel can be found in ref 20. The various model
were obtained using the oil-flow technique. Backup
configurations were pitched through angle of attack in
ceramic models were spray-painted black to enhance
increments of 5 deg. A limited number of runs were
made at incidence angles outside these increments to contrast with the white pigmented oils used to trace
match conditions at points in the X-33 trajectory of streamline movement. A thin basecoat of clear
interest to the aerodynamic community. Sideslip was silicon oil was first applied to the surface, and then a
held fixed at 0 deg with the exception of the tests mist of pinhead-sized pigmented-oil drops was applied
conducted in the 20-Inch Mach 6 CF Tunnel with a onto the surface. After the model surface was
4
sideslip angle of 4 deg. prepared, the model was injected into the airstream
Test T echniq u e s and the development of the surface streamlines was
The rapid advances in image processing recorded with a conventional video camera. The
technology which have occurred in recent years have model was retracted immediately following flow
made digital optical measurement techniques practical establishment and formation of streamline patterns,
in the wind tunnel. One such optical acquisition and post-run digital photographs were taken.
method is two-color relative-intensity phosphor Data R eduction a nd U ncertainty
thermography15-17 which is currently being applied to
A 16-bit analog-to-digital facility acquisition
aeroheating tests in the hypersonic wind tunnels of
system acquired flow condition data on all channels at
NASA Langley Research Center.21-23 With this
a rate of 20 samples per second. Measured values of
technique, ceramic wind tunnel models are fabricated
and coated with phosphors that fluoresce in two Pt,1 and Tt,1 are believed to be accurate to within –2
regions of the visible spectrum when illuminated percent. Heating rates were calculated from the global
with ultraviolet light. The fluorescence intensity is surface temperature measurements using one-
dependent upon the amount of incident ultraviolet dimensional semi-infinite solid heat-conduction
light and the local surface temperature of the equations, as discussed in detail in ref. 17. As
phosphors. By acquiring fluorescence intensity discussed in this reference, the accuracy of the
images with a color video camera of an illuminated phosphor system measurement is dependent on the
phosphor model exposed to flow in a wind tunnel, temperature rise on the surface of the model. For the
surface temperature mappings can be calculated on the windward side heating measurements, the phosphor
portions of the model that are in the field of view of system measurement accuracy is believed to be better
the camera. A temperature calibration of the system than –8%, and the overall experimental uncertainty of
conducted prior to the study provides the look-up the heating data due to all factors is estimated to be
tables that are used to convert the ratio of the green –15%. In areas on the model where the surface
4
temperature rise is only a few degrees (i.e. leeside or presentation (except where noted) to maintain
aerospike engine), the estimated overall uncertainty consistency when viewing or comparing the images.
increases to at least –25%. Repeatability for the On the contour scale, the colors tending towards red
normalized windward centerline (laminar) heat transfer indicate areas of higher heating (temperatures) while
measurements was found to be generally better than the colors towards blue represent areas of lower
–4%. heating. In areas where the local heating exceeds the
stagnation point reference value (h/h > 1), such as
FR
the deflected body flaps or fin leading edges, a gray
Prediction M ethod
(cid:210)overscale(cid:211) will be evident.
X-33 Rev-F flow field computations for Windward F uselage
selected angles-of-attack and test conditions were
The global effect of Reynolds number on the
performed using the General Aerodynamic Simulation
windward surface heating at Mach 6, a = 40 and 20
Program (GASP) code24 developed by Aerosoft Inc.
deg, and d = 20 deg is shown in Figs. 5a-d and 6a-d,
GASP is a three-dimensional, finite-volume code BF
respectively. These two incident angles were selected
which incorporates numerous options for flux-
to illustrate the distinctly different global heating
splitting methods, thermochemical and turbulence
patterns found on the windward surface near the model
models and time-integration schemes. Predicted
base. In order to understand the nature of this elevated
heating from GASP has been validated against flight
heating, windward centerline heating distributions at
data obtained from the Shuttle Orbiter.25 In the
a = 40 deg over a range of Reynolds numbers are
present work, a perfect gas air model was employed
compared with laminar and turbulent prediction in
with a Jacobi time integration scheme. Full viscous
Fig. 7. Comparisons to laminar and turbulent
terms were retained for all three directions. A third-
predictions at other angles of attack can be found in
order Roe flux splitting scheme was used in the
refs. 8 and 9. At the highest Reynolds numbers
normal direction to accurately capture shock and
tested, the experimental heating distributions reveal a
boundary layer gradients and third-order van Leer flux
departure from the laminar prediction near X/L = 0.4.
splitting was used in the streamwise and
This suggests the boundary layer is transitioning to a
circumferential directions to promote stability. All
turbulent state (for these test results, the model
cases were treated as either fully laminar or turbulent
surface was considered smooth and transition was not
(the turbulent computations were performed using a
forced via applied discrete roughness). Thus, it is
modified algebraic Baldwin-Lomax model as discussed
reasonable to hypothesize that the progression of
in ref. 9).
heating images in Figs. 5 and 6 display the evolution
Grid description, sensitivity studies, turbulence
of the laminar windward boundary layer at Re /ft = 1
¥
model, and further details regarding the computational
x 106 to the transitional/turbulent state observed at
solutions presented in this report may be found in ref.
Re /ft = 8 x 106. As expected, boundary-layer
¥
9.
transition first appeared at the aft end of the model and
Results and Discussion was found to move forward with increasing Reynolds
Preface number. The transition onset Reynolds numbers
should not be applied directly to flight due primarily
The X-33 aeroheating results presented in this
to the adverse effect of tunnel noise and inherent
paper are organized around specific locations on the
surface roughness of the phosphor coated ceramic
vehicle such as body flap, canted fin, etc. Presented in
models (which may be different than that found on the
Fig. 4 is a sketch of the X-33 windward surface
metallic surface of the flight vehicle27,28).
detailing the various nomenclature and flow features
that will be referred to in the discussions. Flow Consistent with earlier observations7 made on
visualization in the form of Schlieren and surface oil an X-33 D-loft (603B1001D) forebody, distinctive
flows are used to assist in the analysis of the patterns of the transition front are observed with angle
experimentally measured surface heating. The of attack. Fig. 8 a-c illustrates the significant
discussion of results will highlight some of the more differences found at a = 20, 30, and 40 deg angles of
relevant conclusions to date; a more complete attack at a constant unit Reynolds number of 8 x 106.
presentation of flow visualization observations can be At a = 20 deg (Fig. 8a) two transition fronts
found in ref. 20. symmetric about the centerline are observed. As
incidence angle is increased, the fronts merge (Fig.
Heating distributions are presented in terms of
8b) and eventually coalesce into a single parabolic
the ratio of enthalpy based heat-transfer coefficients shape at a = 40 deg (Fig. 8c). The corresponding
h/h , where h corresponds to the Fay and Riddell26
FR FR experimental surface streamline patterns, Fig. 9a-c,
stagnation-point heating to a sphere with radius equal
indicate boundary layer inflow towards the model
to the wind tunnel model nose (rn=0. 629-in for the centerline at a = 20 deg. The inflow of surface
0.0132 scale 10-in long model). A color bar having a
streamlines, Fig. 9a, results in a flow convergence on
maximum value of h/h = 1 was selected for data
FR the windward centerline, which would thicken the
5
boundary layer locally. The degree of inflow suggests streamlines away from the body flap centerline. This
that crossflow may be the mechanism responsible for outboard directed flow spillage did not appear to
transition at lower incidence angles. By increasing the influence the flow over the nearby canted fin (as was
sensitivity of the color scale (not shown) the presence observed on several X-33 phase I configurations).
of heating striation patterns at low incidence angles The flow reattachment downstream on the body flap
was revealed. It is believed that an array of was observed to be in close proximity and nearly
streamwise-orientated boundary layer vortices exist on parallel to the hinge line at all angles of attack
the windward surface and reveal themselves through suggesting locally higher heating in this area. The
local increases in surface shear and thus heating. This body flap hinge line gap is planned to be sealed to
vortex formation is believed to be indicative of the prevent the circulation of this high energy flow into a
onset of three-dimensional, crossflow-induced cavity. Expansion of the flow around the body flap
transition from laminar to turbulent flow. The edges is indicated by the local streamline curvature;
appearance of heating striatia from vortices entrained the outboard curvature is most pronounced at a = 40
within supersonic/hypersonic boundary layers has deg and is indicative of a stronger lateral pressure
been inferred from flight measurements as well as gradient at the higher angles of attack.
observations in hypersonic wind tunnels; the reader is A 20 deg deflected body flap at angles of attack
referred to ref. 29 for a more complete review of this of 30 and 40 deg produced a disturbance in the flap
phenomenon. streamline pattern, Fig. 10b-c. Computational
Asymmetric boundary layer transition on the predictions of the corresponding flowfield (not
windward surface chine in the vicinity of the fin/body presented) indicated the interaction of the deflected flap
junction was observed on a few isolated occasions shock with the bow shock. The outboard location of
during the test series (see Fig. 8b). This transition the X-33 body flaps (in contrast, for example, to the
was attributed to (unintentional) isolated surface shuttle orbiter body flap) locates the body flap shock
roughness in the phosphor coating created from the system in close proximity to the bow shock.
impacts of small particulates (of less than 5 microns) Computed flow field results show that the resulting
on the ceramic model. Although this observation interaction produced an expansion fan that impinged
initially suggested heightened sensitivity of the chine on the body flap in the same spatial orientation as
to transition, a correlation of isolated roughness data seen in the experimental heating (Fig 5a) and surface
revealed that the off-centerline locations were no more streamline patterns (Fig. 10b-c). A typical side view
sensitive than those on centerline (see ref. 8). schlieren image, Fig. 11, at a = 30 deg, d = 20 deg,
BF
Windward B ody F lap and Re¥/ft = 2 x 106 is presented to characterize the
shock system in this region. The sensitivity of the
Body flap deflections of 0, 10, and 20 deg were
schlieren system did not permit the detection of the
tested to cover the anticipated deflection range. The
expansion wave disturbance. This type of surface
flow features and resulting surface heating on the
disturbance on the flap had not been observed on
deflected flaps was complex and was in large part
earlier X-33 configurations due to shorter length body
determined by the extent of boundary layer separation
flaps. Localized increases in body flap surface heating
ahead of the flap and subsequent flow reattachment.
corresponding to flow reattachment near the hingeline
In turn, the control surface flow separation was
and from the expansion fan disturbance were observed,
largely determined by the state of the boundary layer
Fig. 5a, and will be discussed in more detail at a later
approaching the flaps. As inferred from the a = 40
point. The local heating peak from the disturbance
deg surface heating patterns in Fig. 5a-c, the flow
was observed at all Reynolds numbers but because the
leading up to the deflected control surface is laminar
color bar scale for this report was set to provide the
(as transition is limited to the center region only) and
best sensitivity for presentation of windward surface
appears to separate ahead of the flap hinge line (the
transition images (on the fuselage), the body flap data
cooler region upstream of the hingeline). The off
is off-scale.
centerline transition found at the higher Reynolds
numbers for a = 20 deg, Fig. 6b-d, reduces or nearly In order to provide heating trends on the flaps
in the absence of separation, the entire flap surface
eliminates this separated region.
heating was averaged and presented, Fig. 12, as a
At low Reynolds numbers, however, the
function of Reynolds number for an undeflected body
laminar separation bubble at the flap hinge line flap (d = 0 deg). As expected, the flap surface
persists. The surface streamlines associated with heatingB iFncreased with angle of attack. At a = 30 and
laminar separation at Re /ft = 2 x 106 from d = 20
¥ BF 40 deg the averaged heating to the undeflected flap did
deg is shown, Fig. 10a-c, for a = 20, 30, and 40 deg.
not vary with Reynolds number suggesting laminar
This flow separation bubble becomes smaller with flow. The off centerline transition noted earlier at a
increasing angle of attack. The circulation of
= 20 deg is responsible for the increase in the
separated flow upstream of reattachment was highly
averaged flap heating observed at Re /ft = 4 x 106.
¥
three-dimensional with a strong curvature of the
6
The effect of control surface deflections on the secondary peak correlates with a predicted drop in
averaged flap surface heating is shown Fig. 13a and b. surface pressure from the expansion fan tail.
The figures present data obtained with a laminar (Fig. Canted F in
13a) and turbulent (Fig. 13b) boundary layer
The shock shape about the X-33 F Rev-F
approaching the flap (details of the boundary layer
model is shown in planform view, Fig. 15, for a =
tripping method can be found in refs. 9 and 21). The
30 deg, d = 20 deg, and Re /ft = 2 x 106. The bow
averaged flap heating has been normalized to the BF ¥
measured average with d = 0 deg so as to present shock was observed to interact with the canted fin
BF leading edge at all angles of attack tested. Inherently
heating amplification factors above undeflected levels
three-dimensional, the observed shock orientations
(in each figure, the undeflected flap reference heating
(inset, Fig. 15) most closely resemble a two-
was obtained for both laminar and turbulent
dimensional type VI interaction described by Edney30
conditions). Presented in this fashion, these heating
which is known to produce a shear layer and
amplifications may be used with analytic solutions
with d = 0 deg to provide an estimate of what might expansion fan. Depending upon model angle of
BF attack, this shear layer is thought to be swept over
be expected in flight. For an approaching laminar
the fin upper and/or lower surface producing local
boundary layer (Fig. 13a), maximum heating
augmentation (i.e. for d = 20 deg) above laminar perturbations in fin surface heating. The surface
BF effect from the shock interaction was not easily
undeflected flap levels are 7.5 and 5.3 for 30 and 40
discernable from the streamline patterns on the fins
deg angles of attack. The amplification levels are
(Fig. 9a-c) but was more readily identified in the
reduced by approximately 50% when the flap heating
surface heating images (Fig. 6a-d). The fin windward
is normalized to turbulent undeflected flap heating.
surface heating resulting from this interaction is
The effect of tripping the boundary layer approaching
complex and more distinct at lower angles of attack.
the deflected flap has little effect on the averaged
On the canted fin windward surface, elevated heating
deflected flap heating augmentation factors. This
from boundary layer transition (associated with the
suggests that the boundary layer downstream of
bow/fin-shock interaction and possible leading edge
reattachment is transitional/turbulent.
attachment line contamination31) occurs for unit
While the averaged body flap heating was
Reynolds numbers in excess of Re /ft = 4 x 106 (Fig.
¥
useful to examine trends, the local flow on the
5c-d, and 6c-d). Asymmetric boundary layer
deflected flaps was observed to be quite complex.
transition on the windward fin surface was observed
With the exception of the lowest Reynolds numbers,
on a few isolated occasions during the test series (see
the heating levels found on the deflected body flaps
Fig. 5c) and was attributed to (unintentional) isolated
generally meet or exceed the theoretical reference
surface roughness (discussed earlier) on the fin leading
stagnation level (h ). Right and left body flap
FR edge.
centerline heating distributions extracted from images
taken at a = 40 deg, d = 20 deg, and Re /ft = 2 x At high angles of attack characteristic of
BF ¥ hypersonic entry, the X-33 canted fin lower surface
106 are shown, Fig. 14, for both a laminar and
would be exposed to the flow and, in the traditional
turbulent (tripped) boundary layer approaching the
sense, called a (cid:210)windward(cid:211) surface. At low angle of
flap. For a laminar boundary layer approaching the
attack representative of ascent, however, this same fin
deflected flap, the local centerline heating peaks from
surface would be shadowed from the flow (due to the
flow reattachment (h/h = 1.6) and expansion fan
FR negative fin incidence). Because of this situation, the
impingement (h/h = 1.3) are measured at
FR fin surfaces in the following section are referred to as
approximately 5% and 55% flap chord length and are
(cid:210)upper(cid:211) and (cid:210)lower(cid:211). Canted fin upper and lower
spatially consistent with surface oil flow
surface heating distributions extracted from images
observations. The local heating peak from flow
taken at a = 0 and 30 deg, d = 20 deg, and Re /ft =
reattachment was reduced nearly 60% when the BF ¥
4 x 106 are shown, in Fig. 16a-b. The heating
boundary layer was tripped upstream of the deflected
distributions were taken at the fin half chord station.
flap. Based on this observation, it is believed that the
At a low incidence angle characteristic of ascent, the
tripped boundary layer approaching the flap results in
surface heating to the canted fin lower surface was
a smaller separation bubble and turbulent
less than that measured on the upper fin surface (due
reattachment heating levels. The laminar separation
to the negative fin incidence). At a = 0 deg (Fig.
yields flow reattachment that is transitional and
16a) the heating to the fin lower surface is relatively
heating that is characteristically higher than turbulent
uniform at this fin chord location with h/h = 0.1
results. Computations (not shown) suggest the FR
and with no evidence of locally large increases in
secondary heating peak downstream of reattachment is
surface heating from the bow/fin shock interaction. It
the result of a thinning boundary layer produced from
is believed the disturbances (shear layer) from the
the head of the expansion wave system impinging on
interaction are swept over the fin upper surface
the flap. The rapid decrease in heating downstream of
producing locally elevated heating. For example,
heating to the undisturbed outboard 40% of the fin
7
was found to be near h/h = 0.15 while the disturbed At angles of attack more representative of
FR
inboard 60% of the fin had a heating peak of ascent, the placement of the X-33 TPS split lines in
approximately h/h = 0.35. As anticipated, an the vicinity of the nosecap becomes more crucial. At
FR
increase in angle of attack to 30 deg, Fig. 16b, low incidence angles, the upper fuselage is effectively
produced heating levels on the fin lower surface that a (cid:210)windward(cid:211) surface (recall, vehicle peak heating
are elevated above that measured on the upper surface. during ascent occurs near Mach 9 corresponding to
As noted earlier, boundary layer transition a=13 deg). Laminar centerline heating distributions
associated with the bow/fin-shock interaction occurs along the vehicle upper surface at Re¥/ft = 4 x 106 for
for unit Reynolds numbers in excess of Re /ft = 4 x a = -5, 0, 5, and 10 deg are shown, Fig. 20. These
¥
106. This is more apparent in the extracted fin low angle of attack distributions were obtained as a
heating distributions measured over a range of small part of a validation process aimed at
Reynolds numbers and shown in Fig. 17a-b. The determining TPS split lines.
disturbance from the bow shock interaction on the fin Extrapolation to Flight
lower surface at a = 30 deg, Fig. 17a, is not readily A feature of the phosphor thermography
discernable at the fin half chord station until Re¥/ft = analysis package16 (IHEAT) is the ability to
4 x 106. Similarly, the transitional effects from the extrapolate ground based heating measurements to
shock interaction on the fin upper surface at a = 0 flight radiation equilibrium wall temperatures. The
deg, Fig. 17b, are not apparent below Re¥/ft = 4 x successful application of this technique to predict
106. flight surface temperatures for both laminar and
Engine turbulent conditions was demonstrated in the X-34
program (refs. 16 and 22). Based on the initial
Impingement of the separated flow off the
success with the X-34 data and the good agreement
windward fuselage base onto the lower surface of the
between the X-33 laminar data and GASP prediction
linear aerospike nozzle was inferred from elevated
presented in this report, phosphor data were
heating measured at this location (see Fig. 5b-d).
Extracted heating distributions at a = 40 deg and extrapolated to flight surface temperatures and
compared to an equilibrium laminar GASP flight
Re /ft = 4 x 106, Fig. 18, indicated that at
¥
prediction, Fig. 21. The tunnel data was obtained at
reattachment, peak heating was on the order of h/h =
FR a = 40 deg, d = 20 deg, and Re = 2 x 106. Flight
0.1. Surprisingly, support system interference effects BF ¥,L
conditions at this angle of attack correspond to an
(leeside blade mount vs. base sting mount) in this
altitude of 146,730 ft., velocity of 7,045 ft/s, Mach
region were not apparent. Flowfield computations at
number of 6.6, and a length Reynolds number of 5 x
flight conditions also suggest that the boundary layer
106. The body flap was omitted for the flight
separating off the body is likely to reattach on the
computation. As with the X-34 flight case, no
nozzle. In flight, circulating residual hydrogen
significant real gas effects were anticipated at this X-
through the structure during descent will actively cool
33 flight condition. The phosphor images were
the nozzle.
mapped to the three-dimensional vehicle surface
L eeward F uselage
geometry via a new option in the IHEAT code. The
Much lower heating levels generally global comparison of windward surface temperature
characterize the thermal environment of a vehicle presented in Fig. 21 is shown along with data
leeward surface at hypersonic entry angles of attack. extracted along the centerline. Generally, the surface
Comparison of leeward centerline heating at Re /ft = temperatures compare well over the entire image.
¥
2 x 106 and 4 x 106 with laminar prediction (Re /ft = The extrapolated wind tunnel data generally compares
¥
4 x 106) at a = 40 deg is shown, Fig. 19. The to within 50 deg F of prediction along the centerline.
laminar comparison presented here and at lower angles This type of information provided to the designer
of attack (see ref. 9) was generally within the early in the TPS evaluation process would be
experimental uncertainty (–25%) with the exception invaluable and could potentially result in significant
near the end of the fuselage. This discrepancy may be savings of computational time required for flight
due to the fact the model base and wake flow was predictions.
computationally simplified9. The actual flow Concluding Remarks
expansion from the leeside into the wake during the
One of the key technologies being
tests might result in the higher heating. No
demonstrated on the Lockheed Martin X-33 RLV
significant Reynolds number effects were observed
demonstrator is an advanced metallic Thermal
experimentally which suggests either the leeward flow
Protection System (TPS). The heating environment
remained laminar or the difference between laminar
definition for X-33 incorporates conceptual analysis,
and turbulent leeside heating is small. A heating
ground based testing, and computational fluid
maxima of 7.5% of the stagnation reference value was
dynamics into the TPS design process. This report
measured at X/L=0.2.
provides an overview of the hypersonic aeroheating
8
wind tunnel program conducted at the NASA Langley Sheila Wright and Bert Senter for data acquisition
Research Center in support of the ground based assistance; Bill Wood for CFD analysis support; and,
testing activities. In the past, aeroheating Richard Wheless for documentation assistance. The
information has significantly lagged behind authors gratefully acknowledge their contributions and
aerodynamic information due primarily to model and behind-the- scenes work.
instrumentation complexities associated with References
aerothermodynamic testing. The X-33 program was
1.Bekey, I., Powell, R., and Austin. R., (cid:210)NASA Studies
able to take advantage of recent developments in a
Access to Space,(cid:211) Aerospace America, May 1994,
two-color global phosphor thermography technique,
pp. 38-43.
providing the first opportunity to conduct an
aerothermodynamic screening/trade study concurrent 2.Cook, S. A., (cid:210)X-33 Reusable Launch Vehicle Structural
with aerodynamic tests. The work reported herein Technologies,(cid:211) AIAA Paper 97-10873, Nov. 1996.
served as a baseline for a parallel effort to determine 3.Freeman Jr., D. C., Talay, T. A., and Austin, R. E.,
the effect of discrete roughness and distributed (cid:210)Reusable Launch Vehicle Technology Program,(cid:211)
roughness due to a bowed panel (TPS) array on AIAA Paper IAF 96-V.4.01, Oct. 1996.
boundary layer transition (AIAA-99-3560). The study
4.Powell, R.W., Lockwood, M.K., and Cook, S. A., (cid:210)The
also provided a laminar and turbulent heating database
Road from the NASA Access-to-Space Study to a
with a wide range of parameters from which
Reusable Launch Vehicle,(cid:211) AIAA Paper IAF
engineering and benchmark CFD codes were validated
98-V.4.02, Sept. 1998.
against (AIAA-99-3559).
5.Baumgartner, R. I., and Elvin, J. D., (cid:210)Lifting Body —
Global surface heat transfer images, surface
An Innovative RLV Concept,(cid:211) AIAA Paper 95-3531,
streamline patterns, and shock shapes were measured
Sept. 1995.
on 0.013 scale (10-in.) ceramic models of the
6.Hamilton, H., Berry, S., Horvath, T., and
proposed X-33 configuration in Mach 6 air. The test
Weilmuenster, J., (cid:210)Computational/ Experimental
parametrics included angles of attack from -5 to 40
degs, unit Reynolds numbers from 1x106 to 8x106/ft, Aeroheating Predictions for X-33 Phase II Vehicle,(cid:211)
AIAA Paper 98-0869, January 1998.
and body flap deflections of 0, 10, and 20 deg.
Comparisons of the laminar and turbulent 7.Thompson. R. A., Hamilton, H. H., Berry, S. A., and
experimental data were performed which also served to Horvath, T. J., (cid:210)Hypersonic Boundary Layer
assess the state of the windward boundary layer. It Transition for X-33 Phase II Vehicle,(cid:211) AIAA Paper
was determined that natural transition occurred on the 98-0867, January 1998.
windward surface. The smooth body transition 8.Berry, S. A.,, Horvath, T. J., Hollis, B.R., Thompson.
patterns observed on the windward surface were R. A., Hamilton, H. H., (cid:210)X-33 Hypersonic Boundary
strongly dependent on angle of attack and were Layer Transition,(cid:211) AIAA Paper 99-3560, June 1999.
consistent with early phase II observations (AIAA-98-
9.Hollis, B.R., Berry, S. A.,, Horvath, T. J., Hamilton,
0867). At hypersonic entry incidence angles, a
H. H., and Alter, S., (cid:210)X-33 Computational
complex surface flow environment was observed on
Aeroheating Predictions and Comparisons with
the deflected body flaps downstream of reattachment.
Experimental Data,(cid:211) AIAA Paper 99-3559, June
Experimental and computational results indicated the
1999.
presence of shock/shock interactions that produced
localized heating on the deflected flaps and boundary 10.Miller, C. G., (cid:210)Langley Hypersonic
layer transition on the canted fins. At an incidence Aerodynamic/Aerothermodynamic Testing
angle of 40 deg, a laminar boundary layer approaching Capabilities - Present and Future,(cid:211) AIAA Paper 90-
a 20 deg deflected flap, produced a local heating peak 1376, June 1990.
(h/h = 1.3) from a shock/shock interaction. 11.Reed, R. D. (cid:210)Wingless Flight The Lifting Body
FR
Laminar windward centerline heating data from the Story,(cid:211) NASA SP-4220, 1997.
wind tunnel was extrapolated to flight surface
12.Barret, C., (cid:210)Lifting Body Stability and Control,(cid:211)
temperatures and generally compared to within 50 deg
NASA TM —1999-209255, March 1999.
F of flight prediction.
13.Hollis, B.R., Thompson, R. A., Murphy, K., and
Acknowledgments
Nowak, R., (cid:210)X-33 Aerodynamic/Aerothermodynamic
Without the assistance of the following CFD Validation and Flight Predictions,(cid:211) AIAA Paper
individuals this work would not have been possible: 99-4163, August 1999.
Mark Cagle, Joe Powers, Mike Powers, Mark
14.Murphy, K., Nowak, R., Thompson, R. A., and
Griffith, Ed Covington and Tom Burns for model
Hollis, B.R., (cid:210)X-33 Hypersonic Aerodynamic
design/ fabrication/ instrumentation/ surface
Characteristics,(cid:211) AIAA Paper 99-4162, August 1999.
inspection support; John Ellis, Rhonda Manis, Grace
Gleason, Melanie Lawhorne, Harry Stotler, Steve
Jones, and Jeff Warner for wind tunnel support;
9