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Journal ArticleDOI

X-33 Hypersonic Boundary-Layer Transition

01 Sep 2001-Journal of Spacecraft and Rockets (American Institute of Aeronautics and Astronautics (AIAA))-Vol. 38, Iss: 5, pp 646-657
TL;DR: In this article, 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.
Abstract: Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally examined 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.

Summary (4 min read)

Introduction

  • The Access to Space Study 1 by NASA recommended 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.
  • The goal of the RLV technology 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 improve U.S. economic competitiveness.
  • As part of the Single-Stage-To-Orbit (SSTO) RLV program, the X-33 was developed as a technology demonstrator.
  • The purpose of this investigation was to experimentally examine issues affecting boundary layer transition and the effect of transition on the aeroheating characteristics of the X-33.
  • 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.

Boundary Layer Transition in Flight

  • From the perspective of boundary layer transition, the X-33 has many similarities to the Space Shuttle Orbiter.
  • For a majority of these flights, boundary layer transition has been dominated by surface roughness (Ref. 10), in the form of launchinduced damage and/or protruding gap fillers.
  • Early in the Shuttle program the ceramic TPS was recognized to be relatively fragile and studies were per-Figure 2 X-33 windward surface TPS.
  • The sub-orbital trajectory used for the design of the TPS is shown in Fig. 3 and is designated as Old Malmstrom-4.
  • 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).

Transition Prediction Approach

  • A series of wind tunnel tests (see Table 1 ) were performed to investigate the X-33 aeroheating and boundary layer characteristics, while tracking changes to the configuration.
  • 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 D-loft configuration emerged from the end of the Phase I competition and was heavily tested in the early part of Phase II.
  • 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 characteristics across the speed range) and the size of the body flaps and vertical tails was increased.
  • Then during test 6763, the effect of discrete roughness on the centerline of the Rev-F configuration was investigated.

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 blowdown 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 bottom-mounted model injection system can insert models from a sheltered position 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 visualization tests, the model was exposed to the flow for only a few seconds.

Test Techniques

  • The rapid advances in image processing technology which have occurred in recent years have made digital optical measurement techniques practical in the wind tunnel.
  • 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.
  • By acquiring fluorescence intensity images with a color video camera of an illuminated phosphor model exposed to flow in a wind tunnel, surface temperature mappings can be calculated on the portions of the model that are in the field of view of the camera.
  • 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 encompassing the entire 20-in test core.
  • Surface streamline patterns were obtained using the oil-flow technique.

Model Description

  • Additional details about the various model configurations that were tested can be found in Ref. 16 .
  • To meet these requirements, a unique, silica ceramic investment slip casting method has been developed and patented (Ref. 24) .
  • The fiducial marks used for the present study are shown in Fig. 6 and the non-dimensional locations are listed in Table 2 .
  • The roughness elements used in this study were similar to those used in Refs. 6 and 14, which were fabricated to simulate a raised Thermal Protection System (TPS) tile and were cut from 0.0025-inch thick Kapton tape.

Test Conditions

  • The LaRC 20-Inch Mach 6 Air Tunnel provides a freestream unit Reynolds number variation of 0.5 to 8.0 million per foot.
  • For a 0.0132-scale model, this corresponds to a length Reynolds number of approximately 0.41 to 6.7 million.
  • For the baseline data, the model angle of attack (a) was varied from 20-deg to 40-deg in 10-deg increments and the sideslip was maintained at zero for all the runs presented herein.
  • For each model configuration, the unit Reynolds number was varied between 1 and 8 million per foot to obtain the smooth baseline data for comparison to the tripped data.
  • The Reynolds number where significant nonlaminar flow first appears downstream of the roughness element identifies the critical value.

Data Reduction

  • Heating rates were calculated from the global surface temperature measurements using one-dimensional semiinfinite solid heat-conduction equations, as discussed in detail in Refs. 9 and 19.
  • Based on considerations pre-sented in Ref. 9, phosphor system measurement error is believed to be better than ±8%, with overall experimental uncertainty of ±15%.
  • Heating distributions are presented in terms of the ratio of heat-transfer coefficient h/h FR , where h FR corresponds to the Fay and Riddell 26 stagnation-point heating to a sphere with radius 0.629-in (the nose radius of the Rev-F configuration scaled to the model size).
  • Repeatability of the centerline heat transfer distributions was found to be generally better than ±4%.

Computational Methods

  • Computational predictions for comparison to the wind tunnel aeroheating test results were generated at select angles-of-attack and test conditions using the General Aerodynamic Simulation Program (GASP) code, 27 and the Langley Approximate Three-Dimensional Heating Analysis code.
  • 28 The GASP computations were used to assess the state of the boundary layer for the wind tunnel cases, while the LATCH results were used to generate the boundary layer transition correlation parameters.
  • GASP is a threeÐdimensional, finitevolume Navier-Stokes solver that incorporates numerous options for flux-splitting methods, thermochemical and turbulence models, and time-integration schemes.
  • A perfect gas air model was employed and both fully laminar and fully turbulent solutions were obtained.
  • An integral heating method is used to compute the heating rates along three-dimensional inviscid streamlines.

Smooth Body

  • As will be discussed in more detail subsequently, the increased heating region towards the aft-end of the model identifies the onset of boundary layer transition.
  • Note the changing shape of the transition front as a increases.
  • At the lower angles of attack the forebody generates surface streamlines which curve predominately in towards the centerline .
  • The effect of varying Reynolds number on extracted heat transfer profiles along the model centerline for each a is shown in Fig. 10 with comparison of the smooth body results to tripped cases and to laminar and turbulent heating predictions from Ref. 8.
  • The LATCH results do not compare as favorably along the centerline; however, as shown in Ref. 8, this disparity is only along the centerline as off-centerline the results compare within the experimental uncertainty.

Discrete Roughness along Model Centerline

  • As the turbulent wedge does not start immediately behind the trip, this case would be classified as Òcritical.
  • From the sample results presented in Fig. 11 , one might be tempted to conclude that the X-33 vehicle would be more sensitive to discrete trips at higher angles of attack.
  • Figure 13 illustrates the function in relation to the TPS-design trajectory.
  • Using this value of k/d and the calculated boundary layer thickness at the point on the trajectory that corresponds to the onset of transition, the allowable roughness heights over the windward surface are inferred.

Discrete Roughness Off-Centerline

  • Once the roughness criterion was established for the X-33 centerline, the next step was to verify if the discrete centerline criterion was applicable to off-centerline locations as well.
  • Once the attachment line locations were verified for each angle of attack, the roughness effects were examined.
  • This conclusion is also not supported when the attachment line discrete trip results are compared to the results for the centerline discrete trip correlation (Fig. 12 ).
  • Hollis, et al. (Ref. 8) has shown that the LATCH computations are in very good agreement with Navier-Stokes calculations in the chine region where the attachment lines are located.
  • This information can then be used as input to a swept cylinder boundary layer code (Ref. 30) to compute the momentum thickness Reynolds number for an ÒequivalentÓ swept cylinder.

Distributed Roughness

  • The final transition issue that was investigated for the X-33 was the effect of the bowed metallic TPS panels.
  • The series of tests on the various bowed panel configurations has only recently been completed.
  • The surface streamlines are seen to serpentine around the various bowed panels.
  • The onset of transition results (at x/L = 0.8) for the smooth, discrete, and bowed-panels data for each a was located along these three curves.

Concluding Remarks

  • A series of experimental investigations into several issues affecting boundary layer transition on the X-33 vehicle has been performed in the LaRC 20-Inch Mach 6 Tunnel.
  • These investigations examined natural transition on a smooth body, transition due to discrete roughness on the centerline and attachment lines, and transition due to distributed roughness in the form of Òwavy-wallÓ bowed panels.
  • The experimental heating levels were compared to predictions to determine the onset location of transition and fully turbulent flow.
  • 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.
  • Finally, 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 than a single discrete trip.

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Content maybe subject to copyright    Report

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.

Citations
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Journal ArticleDOI
TL;DR: In this article, it is shown that the body, surface panels, and aerodynamic control surfaces are flexible due to minimum-weight restrictions for hypersonic vehicle configurations, and that these flexible body designs will consist of long, slender lifting body designs.
Abstract: H YPERSONIC flight began in February 1949 when a WAC Corporal rocket was ignited from a U.S.-captured V-2 rocket [1]. In the six decades since this milestone, there have been significant investments in the development of hypersonic vehicle technologies. The NASA X-15 rocket plane in the early 1960s represents early research toward this goal [2,3]. After a lull in activity, the modern era of hypersonic research started in the mid-1980s with the National Aerospace Plane (NASP) program [4], aimed at developing a single-stage-to-orbit reusable launch vehicle (RLV) that used conventional runways. However, it was canceled due mainly to design requirements that exceeded the state of the art [1,5]. A more recent RLV project, the VentureStar program, failed during structural tests, again for lack of the required technology [5]. Despite these unsuccessful programs, the continued need for a low-cost RLV, as well as the desire of the U.S. Air Force (USAF) for unmanned hypersonic vehicles, has reinvigorated hypersonic flight research. An emergence of recent and current research programs [6] demonstrate this renewed interest. Consider, for example, the NASA Hyper-X experimental vehicle program [7], the University of Queensland HyShot program [8], the NASA Fundamental Aeronautics Hypersonics Project [9], the joint U.S. Defense Advanced Research Projects Administration (DARPA)/USAF Force Application andLaunch fromContinentalUnited States (FALCON) program [10], the X-51 Single Engine Demonstrator [11,12], the joint USAF Research Laboratory (AFRL)/Australian Defence Science and Technology Organisation Hypersonic International Flight Research Experimentation project [13], and ongoing basic hypersonic research at the AFRL (e.g., [14–20]). The conditions encountered in hypersonic flows, combined with the need to design hypersonic vehicles, have motivated research in the areas of hypersonic aeroelasticity and aerothermoelasticity. It is evident from Fig. 1 that hypersonic vehicle configurations will consist of long, slender lifting body designs. In general, the body, surface panels, and aerodynamic control surfaces are flexible due to minimum-weight restrictions. Furthermore, as shown in Fig. 2, these

257 citations

Journal ArticleDOI
TL;DR: In this paper, the tradeoff between computational cost and accuracy is evaluated for aerothermoelastic analysis based on either quasi-static or time-averaged dynamic fluid-thermal-structural coupling, as well as computational fluid dynamics based reduced-order modeling of the aerodynamic heat flux.
Abstract: The field of aerothermoelasticity plays an important role in the analysis and optimization of airbreathing hypersonic vehicles, impacting the design of the aerodynamic, structural, control, and propulsion systems at both the component and multi-disciplinary levels. This study aims to expand the fundamental understanding of hypersonic aerothermoelasticity by performing systematic investigations into fluid-thermal-structural coupling, and also to develop frameworks, using innovative modeling strategies, for reducing the computational effort associated with aerothermoelastic analysis. Due to the fundamental nature of this work, the analysis is limited to cylindrical bending of a simply-supported, von K arm an panel. Multiple important effects are included in the analysis, namely: 1) arbitrary, nonuniform, in-plane and through-thickness temperature distributions, 2) material property degradation at elevated temperature, and 3) the effect of elastic deformation on aerodynamic heating. It is found that including elastic deformations in the aerodynamic heating computations results in non-uniform heat flux, which produces non-uniform temperature distributions and non-uniform material property degradations. This results in reduced flight time to the onset of flutter and localized regions in which the material temperature limits may be exceeded. Additionally, the trade-off between computational cost and accuracy is evaluated for aerothermoelastic analysis based on either quasi-static or time-averaged dynamic fluid-thermal-structural coupling, as well as computational fluid dynamics based reduced-order modeling of the aerodynamic heat flux. It is determined that these approaches offer the potential for significant improvements in aerothermoelastic modeling in terms of efficiency and/or accuracy.

224 citations

Journal ArticleDOI
TL;DR: Considerable work still needs to be done before the understanding of hypersonic flow will allow for the accurate prediction of vehicle flight characteristics throughout the flight envelope from launch to orbital insertion.
Abstract: The challenges in understanding hypersonic flight are discussed and critical hypersonic aerothermodynamics issues are reviewed. The ability of current analytical methods, numerical methods, ground testing capabilities, and flight testing approaches to predict hypersonic flow are evaluated. The areas where aerothermodynamic shortcomings restrict our ability to design and analyze hypersonic vehicles are discussed, and prospects for future capabilities are reviewed. Considerable work still needs to be done before our understanding of hypersonic flow will allow for the accurate prediction of vehicle flight characteristics throughout the flight envelope from launch to orbital insertion.

210 citations


Cites background from "X-33 Hypersonic Boundary-Layer Tran..."

  • ...And one article by Berry et al. (2001b) related to objective 5: “Obtain data that can be used to develop empirical correlations for phenomena that resist analytical and/or numerical modeling, such as boundary-layer transition and turbulence modeling.”...

    [...]

Journal ArticleDOI
TL;DR: In this article, the Reynolds number based on height k and edge conditions at k was proposed to measure roughness element height, where k = roughness elements height, N k = average roughness component height, ft L = vehicle length, ft M = Mach number n = exponent, and ft Y = generalized transition parameter ® = angle of attack.
Abstract: Nomenclature a = constant; Fig. 1 C , C 0 = constants k = roughness element height, ft N k = average roughness element height, ft L = vehicle length, ft M = Mach number n = exponent; Fig. 1 N R = Poll’s transition parameter; Eq. (1) Reke = roughnessReynolds number based on height k and edge conditions Rekk = roughnessReynolds number based on height k and conditions at k Reμ = Reynolds number based on height μ and edge conditions U = velocity component parallel to test surface or velocity component perpendicularto attachment line, ft /s V = velocity component parallel to attachment line, ft/s X = generalized disturbanceparameter or axial coordinate along windward centerline x = coordinate perpendicular to attachment line, ft Y = generalized transition parameter ® = angle of attack, deg ± = smooth-wall laminar boundary-layer thickness, ft = Poll’s length scale [Eq. (2)], ft μ = smooth-wall laminar boundary-layermomentum thickness, ft 1 = viscosity, lbm/ft ¢ s o = kinematic viscosity, ft2/s 1⁄2 = density, lbm/ft

203 citations

Journal ArticleDOI
TL;DR: In this article, boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels, including the NASALangleyResearch Center 20-Inch Mach 6 Air and 31-inch Mach 10 Air tunnels and in the HYPULSE Reeected Shock Tunnel at the General Applied Sciences Laboratory.
Abstract: Boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels.Fivedifferenttripconegurationswerecomparedinthreehypersonicfacilities:theNASALangleyResearch Center 20-Inch Mach 6 Air and 31-Inch Mach 10 Air tunnels and in the HYPULSE Reeected Shock Tunnel at the General Applied Sciences Laboratory. Heat-transfer distributions, utilizing the phosphor thermography and thin-elm techniques, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles of attack of 0, 2, and 4 deg; Reynolds numbers based on model length of 1.2 ££ 10 6‐15.4 £ 10 6 ; and inlet cowl door simulated in both open and closed positions. Comparisons of boundary-layer transition as a result of discrete roughness elements have led to the selection of a trip coneguration for the Hyper-X Mach 7 eight vehicle.

186 citations

References
More filters
Journal ArticleDOI
TL;DR: In this article, the Reynolds number based on height k and edge conditions at k was proposed to measure roughness element height, where k = roughness elements height, N k = average roughness component height, ft L = vehicle length, ft M = Mach number n = exponent, and ft Y = generalized transition parameter ® = angle of attack.
Abstract: Nomenclature a = constant; Fig. 1 C , C 0 = constants k = roughness element height, ft N k = average roughness element height, ft L = vehicle length, ft M = Mach number n = exponent; Fig. 1 N R = Poll’s transition parameter; Eq. (1) Reke = roughnessReynolds number based on height k and edge conditions Rekk = roughnessReynolds number based on height k and conditions at k Reμ = Reynolds number based on height μ and edge conditions U = velocity component parallel to test surface or velocity component perpendicularto attachment line, ft /s V = velocity component parallel to attachment line, ft/s X = generalized disturbanceparameter or axial coordinate along windward centerline x = coordinate perpendicular to attachment line, ft Y = generalized transition parameter ® = angle of attack, deg ± = smooth-wall laminar boundary-layer thickness, ft = Poll’s length scale [Eq. (2)], ft μ = smooth-wall laminar boundary-layermomentum thickness, ft 1 = viscosity, lbm/ft ¢ s o = kinematic viscosity, ft2/s 1⁄2 = density, lbm/ft

203 citations

Journal ArticleDOI
TL;DR: In this article, boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels, including the NASALangleyResearch Center 20-Inch Mach 6 Air and 31-inch Mach 10 Air tunnels and in the HYPULSE Reeected Shock Tunnel at the General Applied Sciences Laboratory.
Abstract: Boundary-layer trip devices for the Hyper-X forebody have been experimentally examined in several wind tunnels.Fivedifferenttripconegurationswerecomparedinthreehypersonicfacilities:theNASALangleyResearch Center 20-Inch Mach 6 Air and 31-Inch Mach 10 Air tunnels and in the HYPULSE Reeected Shock Tunnel at the General Applied Sciences Laboratory. Heat-transfer distributions, utilizing the phosphor thermography and thin-elm techniques, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. Parametric variations include angles of attack of 0, 2, and 4 deg; Reynolds numbers based on model length of 1.2 ££ 10 6‐15.4 £ 10 6 ; and inlet cowl door simulated in both open and closed positions. Comparisons of boundary-layer transition as a result of discrete roughness elements have led to the selection of a trip coneguration for the Hyper-X Mach 7 eight vehicle.

186 citations

Journal ArticleDOI
TL;DR: In this paper, a weighted two-color relative intensity e uorescence theory for quantitatively determining surface temperatures on hypersonic wind-tunnel models and an improved application of the one-dimensional conduction theory for use in determining global heating mappings is described.
Abstract: Detailed aeroheating information is critical to the successful design of a thermal protection system (TPS) for an aerospace vehicle. NASA Langley Research Center’ s (LaRC) phosphor thermography method is described. Development of theory is provided for a new weighted two-color relative-intensity e uorescence theory for quantitatively determining surface temperatures on hypersonic wind-tunnel models and an improved application of the one-dimensional conduction theory for use in determining global heating mappings. The phosphor methodology at LaRC is presented including descriptions of phosphor model fabrication, test facilities, and phosphor video acquisition systems. A discussion of the calibration procedures, data reduction, and data analysis is given. Estimates of the total uncertainties (with a 95% cone dence level ) associated with the phosphor technique are shown to be approximately 7 ‐10% in LaRC’ s 31-Inch Mach 10 Tunnel and 8 ‐10% in the 20-Inch Mach 6 Tunnel. A comparison with thin-e lm measurements using 5.08-cm-radius hemispheres shows the phosphor data to be within 7% of thin-e lm measurements and to agree even better with predictions via a LATCH computational e uid dynamics (CFD) solution. Good agreement between phosphor data and LAURA CFD computations on the forebody of a vertical takeoff/vertical lander cone guration at four angles of attack is also shown. In addition, a comparison is givenbetween Mach 6phosphordata andlaminarandturbulentsolutionsgeneratedusing theLAURA,GASP, and LATCH CFD codes on the X-34 cone guration. The phosphor process outlined is believed to provide the aerothermodynamic community with a valuable capability for rapidly obtaining (three to four weeks ) detailed heating information needed in TPS design. Nomenclature A = area of camera array element, m 2 a = effective aperture factor of camera optics, sr b = vehicle wing span from wing tip to wing tip, m C = heat transfer coefe cient constant, h.iw=Tw/ c = specie c heat of model substrate, J/ (kg-K) D = driver constant, iaw.Tw=iw/iTinit F = e ux of light, W/m 2 h = heat transfer coefe cient, kg/ (m 2 -s) I = radiant intensity, W/ (m 2 -sr)

139 citations

Journal ArticleDOI
TL;DR: In this paper, the evolutionary development of subsonic, supersonic, and hypersonic wind tunnels for the study of aerodynamic, aerothermodynamic, and fluid-dynamic characteristics of the flow about models, including transition from laminar to turbulent boundary layers, is discussed.
Abstract: The evolutionary development of subsonic, supersonic, and hypersonic wind tunnels for the study of aerodynamic, aerothermodynamic, and fluid-dynamic characteristics of the flow about models, including transition from laminar to turbulent boundary layers, is discussed. Currently, three supersonic and seven hypersonic wind tunnels are operational at Langley, and two additional tunnels are scheduled to become operational by 1990. In the present work, an effort is made to provide a 'tour'of selected supersonic and hypersonic wind tunnels at NASA-Langley used for aerodynamic and aerothermodynamic testing of models, and to present the evolution of quiet-tunnel technology at this facility over the last decade. It is noted that upgrades to the hypersonic facilities complex are underway in order to provide the high flow quality and improved data accuracy required to calibrate advanced computational fluid-dynamic computer codes. Also to be provided are increased productivity required for configuration development and improved reliability to support major hypersonic programs in an efficient and timely manner.

118 citations

Frequently Asked Questions (17)
Q1. What are the contributions in "X-33 hypersonic boundary layer transition" ?

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. 

In order to obtain accurate heat transfer data using the one-dimensional heat conduction equation, models need to be made of a material with low thermal diffusivity and well-defined, uniform, isotropic thermal properties. 

By iterating on this velocity gradient, the momentum thickness Reynolds number from the swept cylinder calculation can be matched to the approximate three-dimensional momentum thickness Reynolds computed by LATCH. 

Run times up to 15 minutes are possible with this facility, although for the current heat transfer and flow visualization tests, the model was exposed to the flow for only a few seconds. 

The distributed roughness was in the form of a wavywall that simulates the expected metallic TPS Panel bowing in flight due to temperature gradients across the panel. 

During a hypersonic entry, thermal gradients within the metallic TPS panels will produce an outward bowing of the panels on the order of 0.25-in. 

Variations on the roughness heights (k) were obtained by stacking multiple layers of Kapton tape (k = 0.0025, 0.0050, and 0.0075-inch). 

For a majority of these flights, boundary layer transition has been dominated by surface roughness (Ref. 10), in the form of launchinduced damage and/or protruding gap fillers. 

Phosphor thermography is routinely used in Langley's hypersonic facilities as quantitative global surface heating information is obtained from models that can be fabricated quickly (within a few weeks) and economically (cost an order of magnitude less than the thin-film technique). 

The LATCH code (Ref. 5) was used to predict the location of the attachment lines for angles of attack of 20, 30, and 40-deg (with nominal tunnel flow conditions as inputs). 

Examples of typical heating images, which illustrate flow symmetry, and extracted heating profiles along the attachment line are provided in Figs. 16 and 17. 

The following individuals were critical to the successful completion of this work: Mark Cagle, Joe Powers, Mark Griffin, Mike Powers, Rhonda Manis, Grace Gleason, Johnny Ellis, Bert Senter, Sheila Wright, Glenn Bittner, Steve Alter, Matt Kowalkowski, Derek Liechty, and Richard Wheless. 

Further analysis of the distributed bowed panel results is required to determine if bowed panels in the vicinity of the chine region might influence the crossflow dominated flow field at lower angles of attack. 

The Rev-C configuration instituted small modifications to the nose shape (to simplify the construction of the metallic TPS panels) and to the base region (in the vicinity of the engine). 

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 than a single discrete trip. 

Based on earlier results published in Ref. 6, the X-33 boundary layer is expected to transition back to a turbulent state at a Mach number near 9. 

For each model configuration, the unit Reynolds number was varied between 1 and 8 million per foot to obtain the smooth baseline data for comparison to the tripped data.