NASA/TM—2015–218940

Flight Tests of a 

Supersonic Natural Laminar Flow Airfoil 

Michael A. Frederick and Daniel W. Banks

Armstrong Flight Research Center, Edwards, California

G. A. Garzon and J. R. Matisheck

Aerion Corporation, Reno, Nevada

 

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NASA/TM—2015–218940

Flight Tests of a 

Supersonic Natural Laminar Flow Airfoil 

Michael A. Frederick and Daniel W. Banks

Armstrong Flight Research Center, Edwards, California

G. A. Garzon and J. R. Matisheck

Aerion Corporation, Reno, Nevada

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National Aeronautics and

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Armstrong Flight Research Center

Edwards, CA 93523-0273

December 2015
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1

Abstract

A flight-test campaign of a supersonic natural laminar flow airfoil has been recently completed. 

The test surface was an 80-inch (203 cm) chord and 40-inch (102 cm) span article mounted on the centerline 

store location of an F-15B airplane (McDonnell Douglas Corporation, now The Boeing Company, Chicago, 

Illinois). The test article was designed with a leading edge sweep of effectively 0 deg to minimize boundary 

layer crossflow. The test article surface was coated with an insulating material to avoid significant heat 

transfer to and from the test article structure to maintain a quasi-adiabatic wall. An aircraft-mounted infrared 

camera system was used to determine boundary layer transition and the extent of laminar flow. The tests 

were flown up to Mach 2.0 and chord Reynolds numbers in excess of 30 million. The objectives of the tests

were to determine the extent of laminar flow at high Reynolds numbers and to determine the sensitivity of 

the flow to disturbances. Both discrete (trip dots) and 2-D disturbances (forward-facing steps) were tested. 

A series of oblique shocks, of yet unknown origin, appeared on the surface, which generated sufficient 

crossflow to affect transition. Despite the unwanted crossflow, the airfoil performed well. The results 

indicate the sensitivity of the flow to the disturbances, which can translate into manufacturing tolerances, 

were similar to that of subsonic natural laminar flow wings.

Nomenclature

CF crossflow (transition mechanism)

CLIP Centerline Instrumented Pylon

Hp pressure altitude

IR infrared

LE leading edge

M Mach number

Minf free-stream Mach number 

M∞ free-stream Mach number

NASA National Aeronautics and Space Administration

NLF natural laminar flow

r recovery factor (0.88 for turbulent boundary layer)

Re Reynolds number

Reft unit Reynolds number

RTD Resistance Temperature Detector

SBJ Supersonic Business Jet

SBLT Supersonic Boundary Layer Transition

SSNLF Supersonic Natural Laminar Flow

Taw adiabiatic wall temperature

TTot free-stream total temperature

Tturb turbulent adiabatic wall temperature
2

T∞ free-stream static temperature

α angle of attack (airplane reference frame)

β flank angle of attack (airplane reference frame)

γ ratio of specific heats (1.4 for air)

Introduction

A current goal for the National Aeronautics and Space Administration (NASA) is to eliminate 

barriers that prevent the development of practical supersonic cruise vehicles. One of the research thrusts

related to this goal is the development of technologies to increase the cruise efficiency of high-speed 

airframes. Drag reduction through the use of natural laminar flow (NLF) is one area that is currently being 

investigated for efficiency improvements. Conventional supersonic designs feature highly swept wings with

subsonic leading edges to minimize inviscid drag (ref. 1). The large sweep angle generates a spanwise 

pressure gradient that leads to three-dimensionality in the boundary layer, while the use of a subsonic 

leading edge typically results in a wing design with a large leading edge radius and thick foil sections that 

create an adverse streamwise pressure gradient. A drawback of this type of wing design is that it exacerbates 

several different boundary layer instability mechanisms, the end result being a wing with low inviscid drag 

but little to no natural laminar flow.

There are four primary instability mechanisms associated with swept wings: attachment-line, 

streamwise, crossflow, and centrifugal. Attachment-line contamination can occur on swept wings with large 

leading edge radii. This form of instability causes transition at the leading edge if the attachment-line 

Reynolds number exceeds a critical value that is dependent on the initial flow condition (laminar or 

turbulent) at the wing root junction. The streamwise instability is caused by the linear amplification of 

disturbances, referred to as Tollmien-Schlichting waves, which propagate in the boundary layer parallel to 

the wall. Tollmien-Schlichting waves are amplified by an adverse pressure gradient, which is common on 

conventional high-speed designs. The crossflow instability is an inflectional instability and is the dominant 

form of instability for wings with moderate to high sweep (sweep angles greater than approximately

25 deg). Wing sweep can generate a significant spanwise pressure gradient on the wing. This pressure 

gradient results in the formation of a velocity component in the boundary layer or crossflow that is 

perpendicular to the inviscid streamline. The crossflow velocity component must be zero both at the wall 

and at the edge of the boundary layer, resulting in an inflection point in the boundary layer velocity profile. 

This inflection point is a source of instability that occurs as a pair of co-rotating vortices whose axes are 

aligned within a few degrees of the local inviscid streamlines (ref. 2). The centrifugal instability, commonly 

referred to as Görtler instability, is associated with surface concavity and results in the formation of counterrotating vortices that are destabilizing to the boundary layer. This mode of instability can be controlled 

through configuration design. A more thorough description of these four transition mechanisms can be 

found in reference 3, while a detailed review of boundary layer stability theory and transition in general can 

be found in references 4-6.

One method that is currently being investigated to improve the efficiency of high-speed airframes

is viscous drag reduction through the use of natural laminar flow. Compressibility acts to stabilize the 

boundary layer at supersonic speeds, making it possible to obtain significant runs of supersonic natural 

laminar flow if configuration changes that suppress the different boundary layer instability mechanisms are 

incorporated into the airframe design. Increased range, lower fuel burn, and lower aircraft weight are all 

benefits that could be achieved by increasing the amount of laminar flow on an airplane. For a 110,000-lb 

(49,895-kg) class corporate jet with a cruise Mach number of 1.5, increasing the amount of laminar flow 

on the wing from the typical 5 percent to 80 percent could result in an increase in range of approximately 

48 percent (ref. 7).
3

The Aerion Corporation (Reno, Nevada) has developed a design for a supersonic business jet (SBJ)

that features a wing optimized for natural laminar flow (refs. 8-10). The Aerion SBJ has a design range in 

excess of 4000 nm (7408 km). To meet this goal, the wing must maintain laminar flow for transition 

Reynolds numbers on the order of 30 million. The wing design features a sharp leading edge, a low-sweep 

angle, and thin foil sections shaped to produce a favorable streamwise pressure gradient on both the upper 

and lower wing surfaces. The sharp leading edge eliminates the possibility of attachment-line transition

while the favorable streamwise pressure gradient helps suppress Tollmien-Schlichting instabilities in the 

boundary layer. A major benefit of the low-sweep angle is a reduction in spanwise pressure gradients that 

lead to crossflow instabilities in the boundary layer. Another benefit is improved aerodynamic performance 

in the subsonic regime. One drawback of using a low-sweep angle is that, at supersonic speeds, the flow 

component normal to the leading edge is supersonic, resulting in an increase in wave drag. The sharp 

leading edge and thin foil sections featured in the Aerion design help offset the wave drag increase due to 

the low-sweep angle.

NASA and the Aerion Corporation have teamed together to study supersonic natural laminar flow 

and boundary layer transition through a series of joint flight tests using an F-15B (McDonnell Douglas 

Corporation, now The Boeing Company, Chicago, Illinois) NASA research test bed airplane, shown in 

figure 1, dating back to 1999. In the first series of tests, referred to as the Supersonic Natural Laminar Flow 

(SSNLF) project, a small-chord biconvex test article was mounted on the centerline pylon of the F-15B

airplane and flown out to a free-stream Mach number of 1.8. Images of the boundary layer transition were 

obtained from an analog infrared (IR) imaging system mounted on the forward starboard armament rail of 

the F-15B airplane. The SSNLF test article included an insulative coating material to minimize heat 

conduction to the aluminum substructure. Results from the SSNLF project demonstrated the feasibility of 

maintaining large runs of laminar flow on low-swept wings at chord Reynolds numbers up to 10 million 

(refs. 11, 12). An IR image obtained from one of the SSNLF flights is shown in figure 2. In this figure, the

flow direction is from right to left, and the laminar portion of the boundary layer appears dark while the 

turbulent portion appears light.

Figure 1. The NASA F-15B research test bed airplane.
4

Figure 2. Infrared measurements from the SSNLF test article on the NASA F-15B research test bed airplane.

The Supersonic Boundary Layer Transition (SBLT) project followed the SSNLF testing with the 

objective of expanding the results from SSNLF to more complex surfaces with larger chord lengths (higher 

chord Reynolds numbers). A multi-purpose support structure, or “strong back,” was fabricated from 6061 

aluminum for testing large chord test articles on the NASA F-15B airplane using a modified centerline 

pylon referred to as the Centerline Instrumented Pylon (CLIP). The strong back design allows for test 

articles to be bolted onto the right-hand side (test side) of the strong back. The left-hand side (non-test side) 

of the strong back includes an instrumentation bay and five conical probe mounts with 7.5-inch (19 cm) 

spanwise spacing located near the leading edge.

The SBLT project was comprised of two phases. Phase 1, flown in 2010, involved flight-testing a 

flat-plate test article underneath the F-15B airplane out to a free-stream Mach number of 2.0 (ref. 13). The 

primary objective of this test was to quantify the local flow field underneath the F-15B at the flat-plate test 

article location. The flat-plate test article, which was bolted to the right-hand side of the strong back, had 

an 84-inch (213-cm) chord and a 41-inch (104-cm) span with a leading edge sweep of approximately 

10 deg. An array of 60 surface pressure ports was used to measure the pressure on the test surface. On the 

strong back, five 5-hole conical probes were mounted near the leading edge at 7.5-inch intervals along the 

span for measurement of the inflow velocity and direction. The probes were mounted approximately 

3 inches (7.6 cm) to the left of the leading edge to ensure that the shocks from the probes did not contaminate 

the test surface and that the leading edge shock from the plate would not contaminate the probe 

measurements. Surface pressures, local inflow measurements, and IR thermography images were obtained 

out to chord Reynolds numbers in excess of 26 million. Figures 3(a) and 3(b)show local flow measurements 

and an IR image from the flat-plate test article at Mach 1.57. The flow direction in both figures is from right 

to left.
5

(a) (b)

Figure 3. Infrared (a) and (b) local flow measurements from the SBLT phase 1 flat-plate test article on the 

NASA F-15B research test bed airplane.

The surface pressure measurements from the flat-plate test article were used to aid in the design of 

an NLF article for a second phase of flight tests. In 2013, the NLF test article, hereafter referred to as the

test article, was flown on the bottom centerline location of the F-15B airplane using the CLIP and strong 

back previously used for phase 1. The objectives of phase 2 were to determine the extent of laminar flow 

on the test article at supersonic speeds and high Reynolds numbers, to determine the stability of the laminar 

flow with varying Reynolds number, and to determine the sensitivity of the boundary layer to various 

discreet (trip dots) and 2-D (forward-facing steps) disturbances. The IR imagery results from this second 

phase of flight-testing are the focus of this paper. A quantitative analysis of the effects of the discrete 

roughness elements on transition for the test article can be found in reference 14.

Methodology and Approach

The F-15B airplane is a two-seat fighter and trainer version of the F-15A high-performance, 

supersonic air-superiority fighter (both built by McDonnell Douglas Corporation, now The Boeing 

Company, Chicago, Illinois) and is powered by twin F100-PW-100 afterburning turbofan engines 

(Pratt & Whitney, West Palm Beach, Florida). The NASA F-15B airplane has been converted from a United 

States Air Force air-superiority fighter to a research test bed airplane. Research instrumentation, recording, 

telemetry, and video systems have been installed on the airplane. A calibrated National Advisory 

Committee for Aeronautics (NACA) -style flight-test nose boom was installed on the airplane and used for 

measurements of free-stream Mach number, angle of attack, and angle of sideslip. A research total 

temperature probe, installed on the bottom side of the forward fuselage of the airplane, was used for 

measurement of the free-stream total temperature. The total temperature and free-stream Mach number 

measurements were used to derive the static temperature, density, and unit Reynolds number of the 

free-stream flow.

The digital IR sensing system on the NASA F-15B airplane is comprised of an IR camera, a camera 

pod, and analog and digital recording systems. The IR camera utilizes an indium antimonide (InSb) 

640 x 512 focal plane array with a 20-micron pitch. The camera is sensitive in the mid-infrared spectrum

of 3.6 to 5 microns. The camera was fitted with a 13-mm lens to capture the entire test article while 

minimizing the amount of background captured. A streamlined camera pod was designed to house the 

camera. The camera pod mounts on the forward starboard armament rail of the F-15B airplane adjacent to 

the #2 engine inlet, as shown in figure 4. A silicon window with an antireflection coating optimized for the 
6

3 to 5 micron spectral range was installed on the inner surface of the camera pod. Inside the camera pod, a

right-angle prism was mounted adjacent to the window to redirect the IR image toward the camera lens. 

The IR camera simultaneously outputs analog (RS-170) and 14-bit digital (high-speed serial interface) 

video at 30 frames per second. The analog video was recorded using an onboard analog (8 mm) tape 

recorder and was converted to DVD format post-flight. An onboard digital data recorder with approximately 

2 hr of recording time was used to record the digital output from the camera.

Figure 4. The IR camera pod and NLF test article on the NASA F-15B research test bed airplane.

Figure 5. The SBLT phase 2 NLF test article mounted underneath the NASA F-15B research test bed 

airplane.
7

The phase 2 test article, shown in figure 5, had an 80-inch (203-cm) chord and a 40-inch (102-cm) 

span. The test article size was chosen to maximize chord length while maintaining adequate clearance 

between the ground plane and the test article. The test article had a sharp leading edge with a sweep angle 

of 0 deg in order to minimize crossflow for the majority of the test surface. A small section of the leading 

edge near the tip included a radius (curved planform) to purposefully induce crossflow in that region and 

to prevent the formation of a tip shock that could spread across the test article. The test article was cambered 

with a maximum thickness to chord ratio of 6.9 percent and was manufactured from 6061 aluminum coated 

with an epoxy that was machined to produce a smooth surface finish. The epoxy was added to insulate the 

test surface from the aluminum substructure in order to provide quasi-adiabatic conditions at the test 

surface. A thin layer of black matte paint was applied over the epoxy to minimize reflections on the test 

surface. The total weight of the test article was approximately 930 lb.

Four Resistance Temperature Detectors (RTDs) were mounted on the test surface for measurement 

of the surface temperature. The sensors were located along the bottom and the trailing edges of the test 

article, where the boundary layer was expected to be turbulent, to ensure they would not cause transition. 

The locations of the sensors are shown in figure 5. Surface pressure measurements were not included on 

the test article for fear of introducing disturbances that could cause premature transition of the boundary 

layer. Four of the five conical probes used during phase 1 were installed on the strong back. The forward 

locations of the probe mounts on the strong back had been designed to place the tips of the probes adjacent 

to the leading edge of the phase 1 flat-plate test article, which had a leading edge sweep angle of 

approximately 10 deg and a leading edge root location close to 1 inch (2.54 cm) downstream of the phase 

2 test article leading edge root location. These two geometric differences resulted in all of the phase 1 

conical probes, when installed on the strong back, being located behind the phase 2 test article leading edge. 

Given their location behind the test article leading edge, the measurements from the four phase 1 probes 

were not expected to be reliable. In order to collect inflow measurements at the mid-span of the test article

leading edge, an alternate 5-hole conical probe (phase 2) with a longer body than the phase 1 probes was 

installed at the mid-span location on the strong back. A new probe holder was designed to allow the alternate

probe to be installed with its apex adjacent to the test article leading edge. Pressure tubing from the probes 

was routed through channels in the strong back to pressure transducers installed in the strong back 

instrumentation bay. Figure 6 shows the locations of the alternate probe (labeled phase 2 probe) and phase 

1 probes on the strong back, as well as the instrumentation bay.

Figure 6. The back side of the phase 2 NLF test article showing the conical probes, strong back, and 

instrumentation bay.
8

Trip dots and 2-D forward-facing steps were installed on the test article for several of the flights

toward the end of the flight-test series after sufficient baseline transition data had been obtained for the 

clean test article. The objective of these flights was to quantify the boundary layer sensitivity to disturbances 

that would be intrinsic to the manufacture and operation of a supersonic cruise airplane (for example, insect 

contamination and manufacturing tolerances). The trip dots were made from various types of adhesive tapes 

with heights of 2, 3, and 4.5 mil (0.051, 0.076, and, 0.114 mm). The trip dots were sized to be at or below 

the expected local boundary layer thickness for the Reynolds numbers that were investigated.

A 3/16-inch (4.8-mm) punching die and polyimide adhesive tape were used to make the 2-mil dots, 

whereas a 1/16-inch (1.6-mm) hole punch was used to make the 3- and 4.5-mil trip dots from aluminum 

tape. A total of 19 trip dots were applied to the test article. The trip dot layout incorporated spanwise and 

chordwise spacing to ensure that transition wedges could be differentiated while testing different chord 

Reynolds numbers. Figure 7(a) shows the leading edge of the test article with the layout for the trip dots, 

which are highlighted in red. Adhesive-backed vinyl film, typically used for wrapping automobiles, was 

used to create the 2-D forward-facing steps. The film had a nominal thickness of 4.5 mil. Strips of the film

approximately 30 inches (76 cm) long were applied as a single 4.5-mil step and were also layered to create 

step heights of 13.5 and 22.5 mil (0.343 and 0.572 mm). The heights of the steps were chosen to be 

representative of real-world manufacturing breaks. The top and bottom of the steps were tapered inward

toward the trailing edge to account for turbulent wedge formation from the corners of the forward face of 

the steps. The leading edge of the steps was located approximately 8.5 inches (21.6 cm) downstream of the 

test article leading edge. An upper and lower step, each with a different height in order to maximize test 

efficiency, was applied to the test surface, as shown in figure 7(b); the outline of the steps is highlighted in

blue. Thin strips of aluminum tape were placed near the leading edge corners of both steps and near the 

trailing edge of the upper step to provide a frame of reference for the IR images. Aluminum tape was chosen 

because it shows up well on IR images.

 (a) (b)

Figure 7. (a) Discrete roughness elements near the leading edge highlighted in red; and (b) 2-D steps 

outlined in blue.

Supersonic boundary layer transition data were collected during 11 flights. Data were collected at 

speeds up to Mach 2.0 and unit Reynolds numbers up to Reft= 4.6 million/ft (15.1 million/m), which equates 

to a chord Reynolds number Re = 30.7 million. The majority of the test data were collected at Mach 1.7, as 

this provided a local Mach number that was close to the test article design Mach number. Transition data 
9

were collected during each flight using one of two different profiles. Profile 1 consisted of a straight-andlevel acceleration out to the maximum Mach number at a nominal altitude of 42,000 ft (12,802 m). At the 

end of the acceleration, a short steady-state point was obtained at the maximum Mach number followed by 

a straight-and-level deceleration back to subsonic conditions. Profile 2 consisted of a straight-and-level 

acceleration at 42,000 ft out to a nominal Mach number of 1.3, a constant Mach ascent to 49,500 ft 

(15,088 m), and a straight-and-level acceleration from Mach 1.3 to the maximum Mach number, typically 

Mach 1.7. At the end of the acceleration, a short steady-state point was obtained at the maximum Mach 

number followed by a nominally constant Mach descent down to a target altitude. The target altitude during 

the descent was dependent upon the maximum dynamic pressure desired for each individual flight, but was 

limited to no lower than 30,000 ft (9,144 m) or a maximum dynamic pressure of 1200 psf (57.5 kPa) while 

supersonic. The constant Mach descent provided transition data at the test Mach number for a range of 

Reynolds numbers.

Infrared thermography images from the camera analog output were telemetered in real time to the 

control room and were also recorded onboard using the analog recorder. The analog images received 

processing to enhance contrast using algorithms proprietary to the camera manufacturer prior to being 

output into the telemetry and recording streams. The digital images underwent minimal processing by the 

camera and were recorded onboard using the digital data recorder. Both analog and digital images were 

analyzed post-flight; however, the results presented in this paper focus on the analog images as they more 

clearly show the flow features when compared to the minimally-processed digital data. In the future, the 

digital data will be used along with the RTD measurements for temperature mapping of the test article 

surface.

Due to the camera position on the armament rail (forward, above, and to the side of the test article)

and due to the wide-angle lens used in the camera, there was considerable distortion in the IR images. 

A frame of the analog IR from flight 452 showing this distortion is presented in figure 8. In this image and 

all subsequent IR images presented in this paper, the flow direction is from right to left. The analog images 

were transformed into a side-facing perspective post-flight for analysis purposes. In order to perform the 

image transformation, a calibration grid was created using strips of aluminum tape applied to the test surface

along the span and chord directions. A small amount of IR video of the test article with the calibration grid 

was recorded and photographs of the calibration grid were taken from the desired perspective for the 

transformation. A frame from the analog IR video and a photo of the calibration grid are shown in figures 

9(a) and 9(b), respectively. The images from figures 9(a) and 9(b) were loaded into commercially-available 

imaging processing software and a total of 127 control points pairs were located on each figure using the 

calibration grid. A fourth-order polynomial transform was generated from the control point pairs and was 

applied to the analog image to spatially transform the image into the desired perspective. The transform 

was applied frame-by-frame to the IR videos post-flight for analysis purposes.
10

Figure 8. Analog IR image from the SBLT phase 2 flight showing image distortion.

(a) (b)

Figure 9. (a) Analog IR image; and (b) photo of the calibration grid used for IR image transformation.

The spatial transformation of the IR image from figure 8 is shown in figure 10. The image has been 

cropped to minimize unnecessary background clutter. Selected flight parameters appear as text near the top 

of the image, and significant flow phenomenon have been labeled. The transformation has removed the 

majority of the distortion and rotated the image into a side-facing perspective, allowing for more accurate 

determination of the extent of laminar flow. The image clearly shows the laminar-turbulent transition 

boundary. At this flight condition, laminar regions appear dark while turbulent regions appear light. Areas 

where the transition line has a wispy, sawtooth pattern of turbulent wedges, as at the bottom of the test 

article, indicate transition due to crossflow. Portions of the transition line that exhibit a less-distinct 

sawtooth appearance are regions where transition is thought to be due to crossflow induced by weak oblique 

shocks. A turbulent wedge is located at the mid-span location. The wedge is somewhat faint and the apex 
11

location is hard to discern, indicating that the transition here is incipient or in its initial stage. The IR image 

also shows several shock waves impinging on the test surface. A shock emanating from the leading edge 

of the camera pod covers the entire span of the test article, and is located approximately mid-chord in the 

figure. The chordwise impingement location of this shock was observed to be a function of airplane Mach 

number.

Figure 10. Transformed analog IR image showing important flow phenomena.

As shown in figure 10, a series of weak oblique shocks parallel to one another, or a “shock train” 

are present across the span toward the front third of the test article. The “shock train” is primarily located 

in the laminar flow region at this flight condition and appears as a dark series of lines, as opposed to the 

lighter camera pod shock in the turbulent region. The difference in shock color between the laminar and 

turbulent regions seen in the IR image is thought to be an artifact of the local image processing performed

by the IR camera. A close examination of the figure reveals at least six shocks in the train. Only four of 

them stand out clearly in the IR image; these four shocks are identified in the figure. These shocks, hereafter 

referred to as train shocks, are undesirable, because they produce a spanwise pressure gradient, similar to a 

swept wing, which causes crossflow. The pressure rise through the shocks also creates an adverse 

streamwise pressure gradient that would exacerbate Tollmien-Schlichting instabilities. At higher Reynolds 

numbers, several of the individual train shocks caused premature transition on the test article. An example 

of this forced transition is shown in figure 10. The transition line in the area near one-third span marked 

“CF transition” has moved upstream compared to the laminar area directly above, and is coincident with 

the aft train shock indicating that the shock is forcing early transition.

The source of the train shocks is still under investigation. The shocks exhibit a small increase in 

sweep angle with Mach number suggesting that they are not being generated from any airplane structure 

that is located at a considerable distance from the test surface, such as the wing or inlet ramps. If the shocks 

were being generated from a considerable distance away from the test article, they would be expected to

sweep back across the entire surface of the test article as the aircraft Mach number increases, similar to the 

behavior of the camera-pod shock. One hypothesis under investigation is that the shocks are a result of the 

complex flow pattern created by the shape of the fuselage underbody and the spillage from the inlets. If this 
12

were the culprit, one would expect to have seen the train shocks during phase 1 testing of the flat-plate test 

article. The train shocks were not apparent on the phase 1 IR images; however, this may be due to the lack 

of insulation and the higher reflectivity of the phase 1 test article, which resulted in flow features being less 

distinguishable in the IR images from that phase. A few theories regarding the shock formation due to the 

test article shape have been postulated, but all have been deemed to be not credible. The source of these 

shocks requires further investigation.

Results

Data are presented for three of the eleven data flights, with the majority of data from flight 452

(flight 5 of 11). A few of the IR images from two of the roughness element flights (454 and 456) are

presented to show the effects of roughness elements on transition. Profile 2, with a nominal maximum Mach

number of 1.7, was flown on all three of these flights. A time history of aircraft pressure altitude and Mach 

number for the test portion of flight 452 is shown in figure 11. The F-15B airplane was accelerated out to 

Mach 1.7 and 49,500 ft, followed by a constant Mach descent to approximately 33,000 ft (10,058 m). Small 

variations in Mach number occurred throughout the descent due to piloting tolerances.

Figure 11. Flight 452, profile 2 with Mach 1.7 descent to 33,000 ft.

Test Article Conditions with Mach Number

Figure 12 shows the test article surface temperature measurements from the four RTDs during the 

test portion of flight 452. The aircraft total temperature and an estimate of the turbulent adiabatic wall 

temperature are also included in the figure. The adiabatic wall temperature was estimated from equation 

(1), where the recovery factor r = 0.88 for a turbulent boundary layer (ref. 15).
13

Figure 12. Flight 452, free-stream and test article surface temperatures.

Taw = T∞ (1 + r

γ − 1

2

M∞

2 ) (1)

The temperature and Mach number terms on the right-hand side of equation (1) correspond to the

temperature and Mach number at the edge of the boundary layer for each RTD location. As these quantities 

weren’t measured in flight, free-stream values for T∞ and M∞ were used in equation (1), with the 

understanding that this is not a rigorous use of the equation. The laminar adiabatic wall temperature was 

not estimated because the RTDs were located in regions where the boundary layer was known to be 

turbulent. As shown in figure 12, the four surface temperatures and the free-stream total temperature at the 

start of the acceleration maneuver were between -14 and -20 °F (-25.5 and -28.9 °C). The adiabatic wall

temperature was approximately -28 °F (-33.3 °C). Three of the four surface temperature measurements were 

approximately 15 °F (8.3 °C) greater than the adiabatic wall temperature from the start of the acceleration,

indicating “hot wall” conditions. From Mach 1.22 out to Mach 1.7, the surface measurements were all less 

than the adiabatic wall temperature, indicating “cold wall” conditions where heat was being transferred to 

the test article through the boundary layer. The peak surface temperature was experienced at the end of the 

constant Mach descent and was approximately 133 °F (54.4 °C) at the RTD1 location (figure 5) while the 

free-stream total temperature was approximately 185 °F (85 °C). At this same flight condition, the lowest 

surface temperature was 121 °F (49.4 °C) at the RTD4 location. The adiabatic wall temperature at this flight 

condition was 158 °F (70 °C), which was 25 °F (13.9 °C) greater than the highest temperature measured 

on the surface.

Figure 13 shows the local Mach number and flow angularity measured by the phase 2 probe 

installed mid-span on the strong back. The angle-of-attack and flank angle-of-attack values given in figure
14

13 are with respect to the test article. For example, positive angle of attack with respect to the test article is 

equivalent to negative flank angle of attack in the airplane reference frame, whereas positive flank angle of

attack with respect to the test article is equivalent to positive angle of attack in the airplane reference frame. 

The abrupt change in local flow properties near the beginning and end of the profile was due to the passage 

of the camera-pod shock past the probe. At Mach numbers below approximately Mach 1.22 the camerapod shock was in front of the test article leading edge, resulting in transonic conditions at the probe that 

were below the calibration range of the probe. After passage of the camera-pod shock, the flow at the test 

article was supersonic and the conical probe was within its functional calibration range. As shown in figure

13, the test article experienced a steady increase in angle of attack from approximately -1.0 to 1.1 deg as 

the airplane accelerated out to Mach 1.7. The flank angle of attack during this time period varied from 

approximately 0 to -4.3 deg. At the test Mach number of Mach 1.7, the local Mach number at the test article

was approximately Mach 1.65.

Figure 13. Flight 452, test article flow conditions at the mid-span measured from the phase 2 probe.

Transition Pattern with Mach Number

On flight 452, an initial acceleration out to Mach 1.3 was performed at a nominal altitude of 

42,000 ft before performing a constant Mach climb up to 49,500 ft. Figures 14-16 show the start of the 

acceleration from Mach 0.95 out to Mach 1.3. As a reminder, the flow direction in these and all subsequent 

IR figures in this paper is from right to left. Aerodynamic heating due to skin friction at these Mach numbers 

is low. The test article, still retaining some heat from the ground, is warmer than the adiabatic wall 

temperature. The laminar boundary layer acts as an insulator for the test article whereas the turbulent 

boundary layer, with its larger convective-heat transfer coefficient, transfers more heat to the atmosphere. 

As a result, the laminar region of the test article appears lighter than the turbulent portion in these three 

figures. Toward the upper portion of the test article, the laminar region extends beyond the 75 percent chord 
15

location. The small black dots on the test article are areas where the epoxy insulation had to be re-applied 

during manufacturing due to pitting. As the airplane accelerates to Mach 1.2, there is a very small increase 

in the extent of the laminar flow region on the test article. At Mach 1.3, the extent of laminar flow has not 

changed, but the intensity of the laminar region has begun to fade slightly.

Figure 14. Mach = 0.95 during initial acceleration at 42,000 ft.

Figure 15. Mach = 1.2 during initial acceleration at 42,000 ft.
16

Figure 16. Mach = 1.3 during initial acceleration at 42,000 ft.

Figure 17. Mach = 1.35 during constant Mach climb to 49,500 ft.

At higher Mach numbers, the test article temperature increases, but at a slower rate than the 

adiabatic wall temperature. The test article becomes “cold” relative to the adiabatic wall temperature. The 
17

laminar boundary layer is now insulating the “cooler” surface while the turbulent boundary layer transfers 

more heat to the surface, resulting in the laminar regions now appearing darker and the turbulent regions 

appearing lighter in the IR images. Figure 17 shows an image captured at Mach 1.35 during the climb up 

to 49,500 ft. The laminar region extends from the leading edge of the test article to approximately 7 percent 

chord near the root and 13 percent at the tip. The short extent of laminar flow is a result of forced transition 

from the camera-pod shock impinging upon the test article surface. The shock impingement location, 

located at the transition between the darker and lighter areas near the front of the test article, covers the 

entire span of the test article. A spurious turbulent wedge is also apparent near the outboard leading edge 

of the test article.

Figures 18-22 show the shape of the transition line at five different Mach numbers (1.4, 1.5, 1.55, 

1.6, and 1.7) during the straight-and-level acceleration at 49,500 ft. At Mach 1.4, the camera-pod shock

(the aft edge of the dark line) is forcing transition for the entire span. The transition line does, however, 

extend a few inches aft of the shock. The extent of laminar flow has increased because the shock has swept 

further aft due to the increase in Mach number. The transition line is now located at roughly 13 percent

chord at the root and 22 percent chord at the tip. The spurious wedge near the tip of the test article has 

disappeared. Two white lines, located at roughly 65 percent chord, are shocks of a yet unknown origin. At 

Mach 1.5, transition for the entire span of the test article is no longer being forced by the camera-pod shock. 

On the inboard section of the test article, the laminar region extends slightly beyond the camera-pod shock

location and is marked by a slight sawtooth appearance. Directly below this region, at approximately

25 percent span, a small portion of laminar flow extends to approximately 39 percent chord. Moving 

outboard, the transition line moves back toward the leading edge. A large percentage of the boundary layer 

in this region experiences forced transition due to the camera-pod shock. The boundary layer from the tip 

to approximately 5 inches (12.7 cm) inboard has transitioned upstream of the camera-pod shock. Crossflow

induced from the tip geometry (radius) is responsible for the early onset of transition in this region. 

Figure 18. Mach = 1.4 during acceleration at 49,500 ft.
18

Figure 19. Mach = 1.5 during acceleration at 49,500 ft.

A new phenomenon is apparent in the laminar region of the boundary layer at Mach 1.55, shown 

in figure 20. A series of parallel black lines swept back at approximately 30 deg with respect to the leading 

edge are present on the front portion of the test article. These black lines are the shock train described in 

the “Methodology and Approach” section above. The camera-pod shock appears as the furthest aft dark 

line located at the 25 percent chord location. The camera-pod shock and the last train shock intersect at

approximately mid-span. The transition pattern is similar to that seen at Mach 1.5; however, the extent of 

laminar flow has increased as the camera-pod shock has moved aft. A long run of laminar flow at the 

one-third span location persists significantly aft of the camera-pod shock just past 75 percent chord. At 

these higher Mach numbers, the camera-pod shock impingement on the test article occurs at a more oblique 

angle due to the sweep back of the shock with Mach number. The shock, now more oblique than at lower 

Mach numbers, is weaker, allowing the laminar flow to persist beyond the impingement line. The extended 

run of laminar flow finally transitions just upstream of another shock impinging on the test article surface, 

shown as a bright white line.
19

Figure 20. Mach = 1.55 during acceleration at 49,500 ft.

At Mach 1.6, shown in figure 21, a significant portion of the transition line is located beyond the 

camera-pod shock. Between one-quarter and three-quarters span, a large portion of laminar flow persists 

out to approximately 75 percent chord. A turbulent wedge, originating just aft of the camera-pod shock, is 

located just inboard of the mid-span location. The transition line outboard of the large run of laminar flow 

has a wispy, sawtooth appearance indicative of crossflow and is located just aft of the camera-pod shock. 

The transition line in this region is located at roughly 44 percent chord, whereas the transition line near the 

tip is located at roughly 25 percent chord. The difference in transition line chord locations between these 

two regions suggests that the crossflow due to the tip radius is significantly stronger than the crossflow 

generated by the shock impingement.

Figure 21. Mach = 1.6 during acceleration at 49,500 ft.
20

The acceleration at 49,500 ft was terminated at Mach 1.7. The transition line for this Mach number 

is shown in figure 22. A small section of the inboard portion of the boundary layer has transitioned to 

turbulent slightly aft of the camera-pod shock just upstream of the 50 percent chord location.

A large region of laminar flow is present between one-quarter and three-quarters span, extending 

slightly beyond 88 percent chord at mid-span. The turbulent wedge that existed in this region at Mach 1.6 

has disappeared. The crossflow region near the test article tip has extended inboard several inches; however, 

the chord location of the transition line in this region is effectively the same as it was at lower Mach 

numbers. A few small areas of incipient transition appear as light streaks in the laminar region. One of these 

streaks is located aft of the last train shock at about one-third span. Four of the train shocks stand out very 

well in the IR image. The train shocks have swept back a small amount, as compared to their location at 

Mach 1.55, now making an angle of approximately 36 deg with the leading edge. Another interesting 

change has taken place near the leading edge: The region from the leading edge up to roughly 6 percent

chord now appears light instead of dark even though the flow is still laminar in this region. This is a result 

of the test article heating up disproportionately in this area. The thickness of the insulation goes to zero just 

downstream of the leading edge, and the thickness of the aluminum substructure also decreases toward the 

leading edge, resulting in a lower local thermal inertia in this region. The end result is a localized heating 

of the aluminum substructure near the leading edge.

Figure 22. Mach = 1.7 at end of acceleration at 49,500 ft.

Reynolds Number Effects on Transition at Mach 1.7

After the airplane reached Mach 1.7, a constant Mach descent was flown to approximately 

33,000 ft to study the effects of Reynolds number on boundary layer transition for the Mach 1.7 condition. 

Images from the descent with unit Reynolds numbers of Reft = 2.67 million/ft (8.76 million/m) and Reft =

3.49 million/ft (11.45 million/m) are shown in figures 23 and 24. A yellow line showing the transition 

pattern at the beginning of the descent at Reft = 2.14 million/ft (7.02 million/m) from figure 22 has been 

added to both figures for comparison purposes. At Reft = 2.67 million/ft (figure 23) the laminar flow no 

longer extends beyond the camera-pod shock. The largest extent of laminar flow is approximately
21

56 percent chord at the three-quarters span location. The aft train shock appears to be forcing transition in 

the region of one-third to mid-span. The transition line in this region has the familiar wispy, sawtooth 

appearance indicative of crossflow transition. A small wedge indicating incipient transition is present 

behind the second train shock.

Figure 23. Mach = 1.7, Reft = 2.67 million/ft.

Figure 24. Mach = 1.7, Reft = 3.49 million/ft.
22

Figure 25. Mach = 1.7, Reft = 4.31 million/ft.

At Reft = 3.49 million/ft (figure 24) the camera-pod shock is no longer forcing transition on the test 

article. The train shocks appear to be forcing transition for a large portion of the span. The transition line 

has moved almost entirely forward of the 25 percent chord line, and is characterized by a wispy, sawtooth 

pattern in the inboard and outboard regions, and a more jagged pattern with a few distinct wedges mid-span. 

The furthest extent of laminar flow is near the upper portion of the test article. The transition line in this 

region is coincident with the fourth train shock. The aft two train shocks, labeled in figure 24, appear as 

white lines now that the flow on the majority of the test article is turbulent. The transition line in the 

mid-span region is located directly in between the first and second train shocks. The transition region near 

the tip has moved upstream and is now located at approximately 19 percent chord. Figure 25 shows the 

transition results at Reft = 4.31 million/ft (14.14 million/m). The transition pattern has not changed 

significantly from what is shown in figure 24 at Reft = 3.49 million/ft. The two main differences are that the 

transition in the mid-span region has moved slightly forward and has lost its jaggedness, and two faint 

wedges have begun to cause transition in the outboard region that had previously been dominated by 

crossflow.

Roughness Elements Effects on Transition at Mach 1.7

Transition patterns for the trip dots from flight 454 are shown in figures 26-29. A yellow line 

showing the transition line for the clean test article from flight 452 at similar Reynolds numbers (figures 

22-25) has been added to the figures for comparison. The locations of the trip dots have been highlighted 

in red on these figures as well. Starting from the inboard section of the test article and proceeding 

outboard, the trip dot heights in thousandths of an inch (mil) are as follows: 2, 3, 3, 3, 3, 4.5, 4.5, 4.5, 4.5, 

4.5, 2, 2, 2, 2, 2, 3, 3, 3, 3. The last four trip dots were a secondary experiment to see if changes in the 

diameter and shape of the dots had an effect on transition. The fourth to last trip dot had a diameter of 

0.0625 inches while the third to last trip dot had a diameter of 0.25 inches (0.635 cm). The last two trip 

dots were 0.0625-inch squares, with the last square being rotated 45 deg so that a corner of the square was 
23

facing the incoming flow. The beginning of the Mach 1.7 descent at Reft = 2.21 million/ft (7.25 million/m)

is shown in figure 26. At this condition, the majority of the trip dots have not yet caused transition. Two 

turbulent wedges are located at approximately mid-span. The growth of these wedges has eliminated the 

large extent of laminar flow that existed on the clean test article at this Reynolds number. The apex of the 

top wedge is located directly upstream of the second train shock, and has a faint appearance until it has 

passed the shock. Incipient transition starts upstream of the shock and looks to be caused by a 4.5-mil trip 

dot located approximately 6 inches (15.24 cm) downstream of the leading edge. The second train shock 

appears to accelerate transition from this dot. The lower turbulent wedge shows incipient transition 

originating from a 2-mil trip dot that was located approximately 1.375 inches (3.49 cm) from the leading 

edge. It appears that the first train shock may be accelerating the transition of this wedge. Figure 27 shows 

the transition pattern for Reft = 2.67 million/ft. The apex of the first turbulent wedge has moved upstream 

but is still located downstream of the 4.5-mil dot, whereas the apex of the second wedge is clearly located 

at the location of the 2-mil dot. Incipient transition from several of the other trip dots is apparent at various 

locations along the span. At Reft = 3.49 million/ft (figure 28) eight of the nineteen trip dots have caused the 

formation of turbulent wedges. Three of the trip dots located near the test article tip have forced transition 

in the region that, at lower Reynolds numbers, was primarily dominated by crossflow from the tip radius. 

At Reft = 4.31 million/ft (figure 29) thirteen of the nineteen trip dots have turbulent wedges associated with 

them. Three of the 3-mil trip dots located in the inboard section have forced transition in that region. All

four of the 3-mil trip dots in the outboard region have wedges associated with them, with the second to last 

trip dot having transition located upstream of the trip location. 

Figure 26. Mach = 1.7, Reft = 2.21 million/ft with trip dots (red highlight).
24

Figure 27. Mach = 1.68, Reft = 2.67 million/ft with trip dots (red highlight).

Figure 28. Mach = 1.71, Reft = 3.49 million/ft with trip dots (red highlight).
25

Figure 29. Mach = 1.71, Reft = 4.31 million/ft with trip dots (red highlight).

Figures 30-33 show transition patterns from flight 456 with the 2-D forward-facing steps. A yellow 

line showing the transition line for the clean test article from flight 452 at similar Reynolds numbers 

(figures 22-24) has been added to the figures for comparison. On this flight, the upper step had a height of

0.0135 inches (0.0343 cm) and consisted of three layers of film. The lower step had a height of

0.0045 inches (0.0114 cm) and consisted of a single layer of film. Figure 30 shows the transition pattern at 

the beginning of the descent at Reft= 2.23 million/ft (7.32 million/m). The leading edges of both steps show

up as a dark vertical line just downstream of the leading edge. The trailing edge of the upper step shows up 

as white vertical line located several inches behind the camera-pod shock. The trailing edge of the bottom 

step has the same longitudinal location as that of the upper step, but is not visible in the figure. A large 

turbulent wedge originating from the test article leading edge exists on the upper step. A post-flight 

inspection of the test article revealed a nick on the leading edge at this location. Runway debris kicked up 

from the airplane nose gear is thought to have caused the nick. A second turbulent wedge on the upper step 

is located at approximately 28 percent chord. This wedge is a direct effect of the disturbance caused by the 

step height. At approximately mid-span, a large turbulent wedge has formed from the bottom right corner 

of the forward face of the upper step. Even though both steps were tapered toward the trailing edge to 

account for turbulent wedge formation from the corners of the leading edge of each step, the bottom half of

this turbulent wedge has encroached onto the upper portion of the lower step test area. A small section of 

the flow over the lower step remains laminar past the aft portion of the step and beyond the camera-pod 

shock. An incipient wedge can be seen approximately midway down the lower step, but the majority of the 

flow over the lower step is laminar. Two small turbulent wedges are apparent near the test article tip 

upstream of the step. The source of these wedges is not known at this time.

Figure 31 shows the transition pattern for Reft = 2.68 million/ft (8.79 million/m). A small turbulent 

wedge from the top corner of the bottom step has formed. With the exception of the two wedges near the 

tip leading edge, the transition line on the lower step is not significantly different from that of the clean test 

article (yellow line). At Reft = 3.49 million/ft (figure 31) the transition line on both steps has moved 
26

appreciably forward. Two small areas of incipient transition have formed on the lower portion of the upper 

step. The transition pattern on the lower step closely resembles that of the clean test article. Figure 33 shows 

the results for Reft = 4.31 million/ft. On the upper step, there is a very small amount of laminar flow that 

persists beyond the leading edge. The transition pattern on the bottom step has not appreciably changed 

from what it was at Reft = 3.49 million/ft.

Figure 30. Mach = 1.7, Reft = 2.23 million/ft with 2-D forward-facing steps.

Figure 31. Mach = 1.73, Reft = 2.68 million/ft with 2-D forward-facing steps.
27

Figure 32. Mach = 1.69, Reft = 3.49 million/ft with 2-D forward-facing steps.

Figure 33. Mach = 1.69, Reft = 4.31 million/ft with 2-D forward-facing steps.
28

Summary

A flight-test series was undertaken to investigate the extent of supersonic laminar flow on a 

specially designed natural laminar flow test article using the NASA F-15B (McDonnell Douglas 

Corporation, now The Boeing Company, Chicago, Illinois) research test bed airplane. The test article was 

designed with no leading edge sweep to eliminate spanwise pressure gradients that result in boundary layer 

crossflow. An onboard digital infrared imaging system was used to characterize the transition front, and 

thus, the extent of laminar flow on the test article. The infrared images clearly show the boundary layer 

transition line and other flow features such as shock waves impinging on the test article. Results from the 

flights showed laminar flow in excess of 80 percent chord for portions of the span at chord Reynolds number 

in excess of 14 million. At higher Reynolds numbers, transition was dominated by crossflow induced by a 

series of parallel shocks located over approximately the front third of the test article. The presence of these 

shocks was undesirable, as they created a spanwise pressure gradient and an adverse chordwise pressure 

gradient, both of which are destabilizing to boundary layer transition mechanisms. A number of flights were 

flown with discrete and 2-D roughness elements installed on the test article. The results from the disturbance 

flights will be used to benchmark allowable roughness tolerances intrinsic to the construction and operation 

of a supersonic business jet class airplane.
29

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