Scott A. Berry, Thomas J. Horvath, Brian R. Hollis,
Richard A. Thompson, and H. Harris Hamilton II
NASA Langley Research Center,
Hampton, VA 23681
33rd AIAA Thermophysics Conference
June 28 - July 1, 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 HYPERSONIC BOUNDARY LAYER
TRANSITION
AIAA 99-3560
1
X-33 HYPERSONIC BOUNDARY LAYER TRANSITION
Scott A. Berry
*
, Thomas J. Horvath
*
, Brian R. Hollis
*
, Rick A. Thompson, and H. Harris Hamilton II
*
Abstract
Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally ex-
amined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, surface streamline
patterns, and shock shapes were measured on 0.013-scale models at Mach 6 in air. Parametric variations include
angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 6.6 million; and
body-flap deflections of 0, 10 and 20-deg. The effects of discrete and distributed roughness elements on boundary
layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline,
were investigated. The discrete roughness results on centerline were used to provide a transition correlation for the
X-33 flight vehicle that was applicable across the range of reentry angles of attack. The attachment line discrete
roughness results were shown to be consistent with the centerline results, as no increased sensitivity to roughness
along the attachment line was identified. The effect of bowed panels was qualitatively shown to be less effective
than the discrete trips; however, the distributed nature of the bowed panels affected a larger percent of the aft-body
windward surface than a single discrete trip.
*
Nomenclature
M Mach number
M
e
Mach number at edge of boundary layer
Re unit Reynolds number (1/ft)
Re
L
Reynolds number based on body length
Re
q
momentum thickness Reynolds number
a model angle of attack (deg)
d boundary layer thickness (in)
x longitudinal distance from the nose (in)
y lateral distance from the centerline (in)
L reference length of model (10.00 in)
h heat transfer coefficient (lbm/ft
2
-sec),
=q/(H
aw
- H
w
) where H
aw
= H
t2
h
FR
reference coefficient using Fay-Riddell calcula-
tion to stagnation point of a scaled sphere
q heat transfer rate (BTU/ft
2
-sec)
H enthalpy (BTU/lbm)
k roughness element height (in)
W roughness element width (in)
Introduction
The Access to Space Study
1
by NASA recom-
mended the development of a heavy-lift fully reusable
launch vehicle (RLV)
2,3
to provide a next-generation
launch capability to serve National space transportation
needs at greatly reduced cost. This led to the RLV tech-
nology program, a cooperative agreement between
NASA and industry. The goal of the RLV technology
*
Aerospace Technologist, Aerothermodynamics Branch, Aero- and
Gas-Dynamics Division, NASA Langley Research Center, Hampton,
VA 23681.
Member AIAA.
Copyright Ó1997 by the American Institute of Aeronautics and As-
tronautics, Inc. No copyright is asserted in the United States under
Title 17, U.S. Code. The U.S. Government has a royalty-free license
to exercise all rights under the copyright claimed herein for govern-
ment purposes. All other rights are reserved by the copyright owner.
program is to enable significant reductions in the cost of
access to space, and to promote the creation and delivery
of new space services and other activities that will im-
prove U.S. economic competitiveness. The program
implements the National Space Transportation Policy,
which is designed to accelerate the development of new
launch technologies and concepts to contribute to the
continuing commercialization of the national space
launch industry. As part of the Single-Stage-To-Orbit
(SSTO) RLV program, the X-33 was developed as a
technology demonstrator. The X-33 Program will dem-
onstrate the key design and operational aspects of a
SSTO RLV rocket system so as to reduce the risk to the
private sector in developing such a commercially viable
system. The objective of NASA's technology develop-
ment and demonstration effort, as stated in the National
Space Transportation Policy is to support government
and private sector decisions on development of an opera-
tional next-generation reusable launch system by the end
of this decade. In order to meet its objectives, the X-33
program is an aggressive, focused launch technology
development program, with extremely demanding tech-
nical objectives and milestones. A Cooperative Agree-
ment is used between NASA and the industry partner,
Lockheed Martin Skunkworks, to describe the responsi-
bilities and milestones of both NASA and Lockheed.
The X-33 is a slab-delta lifting body design with sym-
metric canted fins, twin vertical tails, and two outboard
body flaps located at the rear of the fuselage. A linear
aero-spike engine (Ref 4) is used to power the X-33,
which is roughly a half-scale prototype of LockheedÕs
RLV design, the VentureStar. Figure 1 provides a com-
parison of the X-33 to the VentureStar and the Space
Shuttle Orbiter.
2
As part of the Cooperative Agreement, NASA Lan-
gley Research Center (LaRC) has been tasked with pro-
viding experimental boundary layer transition and aero-
heating data in support of X-33 aerothermodynamic de-
velopment and design. To satisfy the objectives out-
lined in the task agreements, a combined experimental
and computational approach was utilized. Results from
early wind tunnel heating measurements were compared
to laminar and turbulent predictions (Ref. 5). Prelimi-
nary results associated with the effort to characterize the
boundary layer on the X-33 in flight were reported in
Ref. 6. Since the time of these publications, additional
tests have been completed which supplemented the
original database and accommodated design changes to
the vehicle shape. The most current experimental and
computational aeroheating results are presented in this
report and two companion papers (Refs. 7 and 8).
This report presents an overview of the results to
date of the investigation into boundary layer transition
for the X-33 configuration in NASA Langley Research
Center (LaRC) facilities. The purpose of this investiga-
tion was to experimentally examine issues affecting
boundary layer transition and the effect of transition on
the aeroheating characteristics of the X-33. Over a series
of wind tunnel entries in the LaRC 20-Inch Mach 6
Tunnel, the smooth body transition patterns, the effect
of discrete roughness on and off windward centerline, and
the effect of distributed bowed panels have been exam-
ined. The primary test technique that was utilized during
these tests was the thermographic phosphor technique
9
,
which provides global surface heating images that can be
used to assess the state of the boundary layer. Flow
visualization techniques, in the form of oil-flow to pro-
vide surface streamline information and schlieren to pro-
vide shock system details were also used to supplement
the heating data. Parametrics included in these tests
were the effect of angle of attack (a of 20-deg, 30-deg,
and 40-deg), unit Reynolds number (Re between 1 and 8
million/ft), body flap deflections (d
BF
of 0-deg, 10-deg,
and 20-deg), and roughness. The roughness tests in-
cluded both discrete and distributed trip mechanisms.
The discrete roughness parametrics (which included
height, size, and location) were included in these tests to
provide information to develop roughness transition
correlation for the X-33 vehicle and included results from
both the centerline and attachment lines of the X-33.
The distributed roughness was in the form of a wavy-
wall that simulates the expected metallic TPS Panel
bowing in flight due to temperature gradients across the
panel.
Boundary Layer Transition in Flight
From the perspective of boundary layer transition,
the X-33 has many similarities to the Space Shuttle
Orbiter. Upon descent, both vehicles fly at angles of
attack near 40-deg, which results in a moderately blunt
flowfield that produces similar boundary layer edge con-
ditions on the windward surface (M
e
between 1.5 to 2.0).
The Thermal Protection System (TPS) tiles that protect
the windward surface are laid out in a diamond pattern
similar to the Shuttle (see Fig. 2). The knowledge
gained from the flight experience of the Shuttle forms
the starting point with which to assess transition for the
proposed X-33 flights.
The Shuttle Orbiters have flown numerous reentries
into the earthÕs atmosphere. For a majority of these
flights, boundary layer transition has been dominated by
surface roughness (Ref. 10), in the form of launch-
induced damage and/or protruding gap fillers. The ran-
dom nature of this roughness allows for a wide range of
free-stream conditions for the onset of transition: Mach
numbers between 6 and 18 and length Reynolds numbers
between 2.5 and 13 million. This amount of scatter
does little to induce confidence with regard to reliable
prediction of hypersonic boundary layer transition for
future reentry vehicles.
Early in the Shuttle program the ceramic TPS was
recognized to be relatively fragile and studies were per-
Fi
g
ure 2 X-33 windward surface TPS.
Figure 1 Comparison of X-33 to proposed RLV and
the S
p
ace Shuttle.
3
formed to suggest alternatives which offer more durabil-
ity and operability without sacrificing weight (Ref. 11).
During these studies, a metallic TPS was identified as
the lightest system to provide a significant improvement
in durability and operability for the shuttle program, but
was never implemented. A derivative of this metallic
TPS has been selected for use on the windward surface of
the X-33 (Refs. 12 and 13), as shown in Fig. 2. This
system is expected to offer improved durability against
the random surface defects that continue to plague the
Shuttle. However, a detriment of this system is that it
provides an additional type of surface roughness that has
received very little attention over the years. During a
hypersonic entry, thermal gradients within the metallic
TPS panels will produce an outward bowing of the pan-
els on the order of 0.25-in. The effect of this panel
bowing on hypersonic boundary layer transition is
largely unknown and is part of the current investigation
that will be described in this paper.
The sub-orbital trajectory used for the design of the
TPS is shown in Fig. 3 and is designated as Old Malm-
strom-4. This is a high Mach number trajectory that
would land the X-33 at Malmstrom Air Force Base in
Montana. The initial flights of the X-33 are lower
Mach number trajectories to Michaels Air Force Base in
Utah (preliminary Michaels trajectories are also shown
for comparison). Some select points (relevant to the
TPS design) along the Old Malmstrom-4 trajectory are
illustrated in Fig. 3. On ascent the X-33 windward sur-
face is assumed, for the purpose of the TPS design, to
remain turbulent until Re
L
= 2 million. Thus, during
peak heating on ascent, the windside boundary layer will
be laminar and the vehicle will be at a low angle of at-
tack. As the top of the trajectory is reached, the vehicle
pitches up to high angles of attack for most of the hy-
personic descent. A second peak heating point occurs on
descent at a Mach number near 11. Based on earlier re-
sults published in Ref. 6, the X-33 boundary layer is
expected to transition back to a turbulent state at a Mach
number near 9. However, when viewed against the var-
ied results experienced on the Shuttle, also shown in
Fig. 3, the need to minimize known transition by-pass
mechanisms, which could force earlier transition (prior
to the descent peak heating point), is evident.
Transition Prediction Approach
A series of wind tunnel tests (see Table 1) were per-
formed to investigate the X-33 aeroheating and boundary
layer characteristics, while tracking changes to the con-
figuration. Over 1100 tunnel runs from 16 entries in
two facilities have been completed on four X-33 con-
figurations since Aug 1996. The evolution of the X-33
configuration from the onset of Phase II has necessitated
multiple entries into LaRC facilities to investigate the
effects of outer mold line (OML) changes to X-33's
aeroheating environment. The OML for the X-33 started
with the original D-Loft concept, then the F-Loft Revi-
sion C (Rev-C), the F-Loft Revision F (Rev-F), and
finally the F-Loft Revision G (Rev-G). These four con-
figurations have been tested in LaRC facilities for both
baseline aeroheating data (i.e. wide ranges of angle of
attack, yaw, Reynolds number, body flap deflection,
etc.), and to investigate the effects of surface roughness
(both discrete and distributed), and test technique (model
scale, blade vs. sting support, etc.). The D-loft configu-
ration emerged from the end of the Phase I competition
and was heavily tested in the early part of Phase II. The
Rev-C configuration instituted small modifications to
the nose shape (to simplify the construction of the me-
tallic TPS panels) and to the base region (in the vicinity
of the engine). The Rev-F has the same forebody shape
as Rev-C, but the dihedral of the canted fins was lowered
from 37-deg to 20-deg (to improve pitch-trim character-
istics across the speed range) and the size of the body
flaps and vertical tails was increased. Finally the Rev-G
had some minor modifications to the leeside and canted-
fin fillet. Additional details regarding the OML changes
can be found in Ref. 7.
The testing sequence and model configurations
tested are listed in Table 1. First, the effect of discrete
roughness elements on the centerline of the D-Loft fore-
body was investigated in test 6737. Then during test
6763, the effect of discrete roughness on the centerline
of the Rev-F configuration was investigated. These
tests utilized the same approach that was used during an
investigation into discrete roughness elements on the
Shuttle Orbiter (Ref. 14) that has shown good agreement
0
50
100
150
200
250
300
0 5 10 15 20
Old Malmstrom 4
Michael 10a-1
Michael 9d
Shuttle entry corridor
Altitude, Kft
Mach Number
Worst Case
Shuttle
Experience
Best Case
Shuttle
Experience
Ascent Re-Laminarization
Mach = 7
Re = 2 million
a = 0-deg
Ascent Peak Heating
Mach = 14.8
Re = 0.4 million
a = 6-deg
Descent Peak Heating
Mach = 11.4
Re = 3 million
a = 32-deg
Descent Transition
Mach = 9.4
Re = 4 million
a = 20-deg
Figure 3 Malmstrom-4 (Old) trajectory.
4
with flight data (Ref. 15). These early test results were
presented in Ref. 6. The effect of distributed roughness
in the form of a wavy-wall surface (simulating bowed
metallic thermal protection system tiles) on Rev-F was
investigated in test 6769. The windward attachment line
was examined during test 6770 for comparison to trends
found on the centerline. These results are detailed in
Ref. 16. And finally, extended bowed panels on Rev-G
were investigated during test 6786. Collectively, these
tests provide a systematic investigation of several differ-
ent boundary layer trip mechanisms while tracking the
OML changes of the X-33 vehicle.
Experimental Methods
Test Facility
The present experiments were conducted in the
LaRC 20-Inch MachÊ6 Air Tunnel. Miller (Ref. 17)
provides a detailed description of this hypersonic blow-
down facility, which uses heated, dried, and filtered air as
the test gas. Typical operating conditions for the tunnel
are stagnation pressures ranging from 30 to 500 psia,
stagnation temperatures from 760 to 940-degR, and
freestream unit Reynolds numbers from 0.5 to 8 million
per foot. A two-dimensional, contoured nozzle is used
to provide nominal freestream Mach numbers from 5.8
to 6.1. The test section is 20.5 by 20 inches; the nozzle
throat is 0.399 by 20.5-inch. A bottom-mounted model
injection system can insert models from a sheltered posi-
tion to the tunnel centerline in less than 0.5-sec. Run
times up to 15 minutes are possible with this facility,
although for the current heat transfer and flow visualiza-
tion tests, the model was exposed to the flow for only a
few seconds. Flow conditions were determined from the
measured reservoir pressure and temperature and the
measured pitot pressure at the test section.
Test Techniques
Surface Heating
The rapid advances in image processing technology
which have occurred in recent years have made digital
optical measurement techniques practical in the wind
tunnel. One such optical acquisition method is two-
color relative-intensity phosphor thermography, which
is currently being applied to aeroheating tests in the
hypersonic wind tunnels of NASA LaRC. Details of
the phosphor thermography technique are provided in
Refs. 9, 18, and 19, while Refs. 6, 14, 20, and 21 are
recent examples of the application of the technique to
wind tunnel testing. With this technique, ceramic wind
tunnel models are fabricated and coated with phosphors
that fluoresce in two regions of the visible spectrum
when illuminated with ultraviolet light. The fluores-
cence intensity is dependent upon the amount of incident
ultraviolet light and the local surface temperature of the
phosphors. By acquiring fluorescence intensity images
with a color video camera of an illuminated phosphor
model exposed to flow in a wind tunnel, surface tem-
perature mappings can be calculated on the portions of
the model that are in the field of view of the camera. A
temperature calibration of the system conducted prior to
the study provides the look-up tables that are used to
convert the ratio of the green and red intensity images to
global temperature mappings. With temperature images
acquired at different times in a wind tunnel run, global
heat transfer images are computed assuming one-
dimensional heat conduction. The primary advantage of
this technique is the global resolution of the quantitative
heat transfer data. Such data can be used to identify the
heating footprint of complex, three-dimensional flow
phenomena (e.g., transition fronts, turbulent wedges,
boundary layer vortices, etc.) that are extremely difficult
to resolve by discrete measurement techniques. Phos-
phor thermography is routinely used in Langley's hyper-
sonic facilities as quantitative global surface heating
information is obtained from models that can be fabri-
cated quickly (within a few weeks) and economically
(cost an order of magnitude less than the thin-film tech-
nique). Recent comparisons of heat transfer measure-
ments obtained from phosphor thermography to conven-
tional thin-film resistance gauges measurements (Ref.
22) and CFD predictions (Ref. 5, 6, 20, and 23) have
shown excellent agreement.
Flow Visualization
Flow visualization techniques, in the form of
schlieren and oil-flow, were used to complement the
surface heating tests. The LaRC 20-Inch Mach 6 Air
Tunnel is equipped with a pulsed white-light, Z-pattern,
single-pass schlieren system with a field of view en-
compassing the entire 20-in test core. Surface stream-
line patterns were obtained using the oil-flow technique.
Both schlieren and oil-flow images were recorded with a
high-resolution digital camera.
-FS0_00-
-BL0_00-
70.0¡
40.2¡
10.0
-WL0_00-
-BL0_00-
6.08
20¡
R 0.629
Front View Port Side View
Leeward View
Windward View
Fi
g
ure 4 X-33 Rev-F Confi
g
uration.
