NASA/TM—2011–215977

Flight Test Results from the Rake Airflow 

Gage Experiment on the F-15B Airplane

Michael A. Frederick and Nalin A. Ratnayake

Dryden Flight Research Center, Edwards, California

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September 2011
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NASA/TM—2011–215977

Flight Test Results from the Rake Airflow 

Gage Experiment on the F-15B Airplane

Michael A. Frederick and Nalin A. Ratnayake

Dryden Flight Research Center, Edwards, California

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National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, CA 93523-0273

September 2011
 

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1

Abstract

The Rake Airflow Gage Experiment involves a flow-field survey rake that was flown on the 

Propulsion Flight Test Fixture at the National Aeronautics and Space Adminstration Dryden Flight 

Research Center using the Dryden F-15B research test bed airplane. The objective of this flight test was to 

ascertain the flow-field angularity, local Mach number profile, total pressure distortion, and dynamic 

pressure at the aerodynamic interface plane of the Channeled Centerbody Inlet Experiment. This new 

mixed-compression, supersonic inlet is planned for flight test in the near term. Knowledge of the 

flow-field characteristics at this location underneath the airplane is essential to flight-test planning and 

computational modeling of the new inlet, and it is also applicable for future propulsion systems research 

that may use the Propulsion Flight Test Fixture. This report describes the flight-test preparation and 

execution, and the local flow-field properties calculated from pressure measurements of the rake. Data 

from the two Rake Airflow Gage Experiment research flights demonstrate that the F-15B airplane, flying 

at a free-stream Mach number of 1.65 and a pressure altitude of 40,000 ft, would achieve the desired local 

Mach number for the future inlet flight test. Interface plane distortion levels of 2 percent and a local angle 

of attack of –2° were observed at this condition. Alternative flight conditions for future testing and an 

exploration of certain anomalous data also are provided.

Nomenclature

CCIE Channeled Centerbody Inlet Experiment

CDE Cone Drag Experiment

LMI Local Mach Investigation

m mean flow quantity

M∞ free-stream Mach number

ML local Mach number

NACA National Advisory Committee for Aeronautics

NASA National Aeronautics and Space Administration

PCM pulse code modulation

PFTF Propulsion Flight Test Fixture

pt local total pressure, psi

q local dynamic pressure, psf

RAGE Rake Airflow Gage Experiment

RBCC rocket-based combined-cycle

s standard deviation

α local angle of attack, deg

αf local flank angle of attack, deg

I. Introduction

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, St. Louis, Missouri) and is powered by twin F100-PW-100 afterburning turbofan engines 

(Pratt & Whitney, West Palm Beach, Florida). The National Aeronautics and Space Administration 

(NASA) Dryden (Edwards, California) 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 in the airplane. A calibrated National Advisory Committee for 

Aeronautics (NACA)-style flight-test nose boom was installed for measurements of free-stream Mach 

number, angle of attack, and angle of sideslip. Figure 1 shows the F-15B airplane with the Rake Airflow 

Gage Experiment (RAGE) during a flight test.
2

Figure 1. The NASA Dryden F-15B research test bed with Propulsion Flight Test Fixture pylon and Rake 

Airflow Gage Experiment rake during flight test.

The Propulsion Flight Test Fixture (PFTF), shown in figure 2, mounts to a standard centerline pylon 

of the NASA Dryden F-15B airplane. It was developed at NASA Dryden for flight-testing 

propulsion-related experiments such as inlets and nozzles at a range of subsonic and supersonic flight 

conditions (ref. 1). The PFTF has a length of 107 in., a height of 19 in., and a maximum thickness of 

10 in. A significant amount of internal space in the fixture is provided for research instrumentation and 

propellant storage tanks. Standard PFTF instrumentation includes the following: (1) a 

six-degree-of-freedom inertial sensing unit, (2) absolute and differential pressure transducers, (3) a total 

temperature probe, and (4) an integrated six-component force balance capable of measuring forces and 

moments in all three axes. Research experiments are flown underneath the centerline of the PFTF 

connected to the force balance hardware. Experimental data obtained during flight tests are sampled and 

transmitted through the pulse code modulation (PCM) system of the PFTF to the airplane, and can be 

recorded on board and telemetered down to the ground for real-time monitoring. 
3

Figure 2. Propulsion Flight Test Fixture with Rake Airflow Gage Experiment.

The PFTF has been flown in two previous flight experiments. Using a NACA air data boom mounted 

to a cylinder and conical nose cap, the Local Mach Investigation (LMI) flights (ref. 2) quantified the local 

Mach number and flow angle at a single point underneath the F-15B airplane at various flight conditions. 

Additionally, the Cone Drag Experiment (CDE) flights (ref. 3) tested the drag-measuring capabilities of 

the PFTF’s integrated force balance from subsonic speeds to Mach 2.

To support future propulsion research using this unique flight test fixture, the flow quality underneath 

the F-15B airplane must be quantified and correlated to the airplane free-stream conditions. The RAGE 

was designed to measure the Mach number, flow angularity, total pressure distortion, and dynamic 

pressure in front of the PFTF at the aerodynamic interface plane of the Channeled Centerbody Inlet 

Experiment (CCIE), an experimental mixed-compression, supersonic inlet (ref. 4) that will be flown on 

the PFTF in the near future. The data from this experiment are aimed at providing the CCIE and future 

PFTF experiments with estimates of flow speed, direction, and distortion prior to flight tests. 

II. Rake Airflow Gage Experiment

The RAGE involves a flow-field survey rake that was flight-tested on the PFTF using the NASA 

Dryden F-15B research test bed airplane. This section describes the experiment design and setup in detail.

A. Experiment Design

The RAGE, designed by NASA Dryden, consists of a flow-field survey rake, cylindrical boom, and 

cone cylinder. Figure 2 shows the RAGE and PFTF mounted to the NASA Dryden F-15B research 

airplane. The rake, shown in the far right of the picture, is secured to an aluminum boom that is attached 

to an aluminum cone cylinder. This cone cylinder, known as the crayon because of its shape, is 

mechanically attached to the PFTF force balance through four bars that bolt to two sleeves on the crayon. 

The flow-field survey rake consists of nine five-hole conical probes mounted in a cruciform 

configuration, as shown in figure 3. The span of the rake is 9.5 in., and the chord is 2.5 in., with a 
4

maximum thickness of 0.375 in. A span of 9.5 in. was chosen to represent the size of a typical experiment 

on the PFTF. The rake is a two-piece design, comprised of a front and rear section that bolt together. The 

rake is made of 7075-T6 aluminum, chosen for its high specific strength and ease of machining. A 

channel between the front and rear sections allows for routing of the probe pressure tubing. This geometry 

was chosen to accommodate pressure tube routing, bending stress constraints, machining requirements, 

and wind-tunnel flow blockage concerns.

Figure 3. Rake Airflow Gage Experiment flow-field survey rake consisting of nine five-hole conical 

probes.

The conical probes, manufactured by Aeroprobe Corporation (Blacksburg, Virginia), are fabricated 

from stainless steel and are comprised of a conical forebody and cylindrical afterbody. Each probe has a 

diameter of 0.125 in. and a length of 3.75 in. The cap of the probe is a 60-degree right-circular cone with 

the tip of the cone replaced by a 0.015-in. pitot port situated normal to the longitudinal axis of the cone. 

Four static pressure ports are located on the surface of the cone. Two ports lie diametrically opposed on 

the vertical meridian of the cone, and the remaining two are situated in the same fashion on the horizontal 

meridian. Inconel® alloy 600 (Special Metals Corporation, New Hartford, New York) exit tubulations are 

connected to the ports for transmission of the cone surface and pitot pressures. Figure 4 shows a 

schematic of the probe and port numbering convention. The conical probes extend 3.0 inches in front of 

the leading edge of the rake. This placement helps to ensure that the pressure ports are located forward of 

any bow shocks originating from the rake body at supersonic conditions.
5

Figure 4. Rake layout and probe and port numbering convention (flight 1).

Before the flight test, the rake was calibrated in the Boeing Polysonic Wind Tunnel in St. Louis, 

Missouri. Calibration data were taken at various combinations of angles of attack and sideslip out to 

± 10° at Mach numbers of 1.46, 1.51, and 1.61. A calibration algorithm for the calculation of local Mach 

number, flow angularity, total pressure, and dynamic pressure at each of the nine probes was developed 

and coded for real-time calculations in flight. Table 1 shows the estimated uncertainties in the 

probe-calculated flow quantities. More detail on the design and calibration of the rake can be found in 

reference 5.

Table 1. Rake uncertainties.

Quantity Uncertainty

Mach number ±0.02

α ±0.50°

αf ±0.50°

pt ±1%

q ±1%

B. Experiment Setup

The RAGE flow-field rake was positioned at the aerodynamic interface plane of the Channeled 

Centerbody Inlet, as depicted in figure 5. The rake was secured in this location by a cylindrical boom that 

was attached to a cone cylinder. The cone cylinder allows for experiment integration with the PFTF 

through the force balance assembly. For the RAGE flights, the cone cylinder was inclined towards the 

underbody of the airplane by 5° measured from a horizontal plane parallel to the waterlines of the 

airplane. The purpose of inclining the rake was to better align the probes with the local velocity vector 

during flight at supersonic Mach numbers. Data from the LMI experiment suggest that at supersonic 

conditions, there is a variable downwash underneath the airplane that ranges from approximately 3 to 8° 
6

over a free-stream Mach range of M∞= 1.3–2.0. Although the downwash underneath the airplane was 

expected to exceed 5° at the RAGE test points, the cone cylinder could not be inclined further because of 

minimum ground clearance requirements when the PFTF is mounted underneath the airplane.

Figure 5. Comparison of Rake Airflow Gage Experiment configuration with Channeled Centerbody Inlet 

Experiment configuration, showing the aerodynamic interface plane.

Flexible urethane tubing was drawn through the cone cylinder and boom hardware to connect the 

probe tubulations, which terminated at the back of the rake, to differential pressure scanners mounted on 

the instrumentation tray located underneath the PFTF force balance. A single ± 30-psid 32-port 

electronically scanned differential pressure transducer was used to measure the pitot pressures of each of 

the nine conical probes, while two ± 15-psid 32-port electronically scanned differential pressure 

transducers were used to measure the surface static pressures on the RAGE probes. A known reference 

pressure was plumbed to the reference ports on the three differential pressure transducers. The reference 

pressure selected for these three transducers was provided by four cone surface static measurements taken 

from a separate conical probe that is located on the front of the PFTF. The four static ports of the PFTF 

probe were connected to a manifold inside the PFTF. The output from the manifold was connected to the 

reference pressure ports on each differential pressure transducer and to a 15-psia absolute pressure 

transducer for an accurate measurement of the reference pressure. Before the flight, the pressure tubing 

from the conical probes to the differential pressure transducers was checked extensively to identify and 

correct any excessive leaks in the system. All 45 pressure ports had leak rates that were within the noise 

level of the differential pressure transducers (0.05 psi) and were deemed acceptable for the purposes of 

this experiment.

III. Flight Test

The experimental data for the RAGE program were obtained during two separate research flights that 

were conducted in the second half of 2009. This section describes the objectives, approach, and results 

from the flight test in detail.
7

A. Objectives and Approach

The primary objective of the RAGE was to quantify the local flow field (Mach number, flow 

angularity, total pressure distortion, and dynamic pressure) at the aerodynamic interface plane of the 

CCIE and to correlate this flow field to airplane free-stream conditions. The CCIE was designed and sized 

to provide a specific mass flow to a notional rocket-based combined-cycle (RBCC) (ref. 6) engine at a 

local Mach number, ML, of 1.5 with a local dynamic pressure, q , of 1,000 psf. As a result, the 

identification of the free-stream Mach number that results in ML = 1.5 at the inlet interface plane, along 

with measurements of the local flow angularity, total pressure distortion, and dynamic pressure at 

ML = 1.5, were of primary importance to the RAGE. Of secondary importance was the determination of 

ML = 1.6, because this test point is a possible test point for the CCIE. This particular Mach number also 

represents the upper end of the rake’s calibration. Using nine probes in a cruciform array, as opposed to a 

single probe, allows for the determination of a more robust average interface plane value for each flow 

property and also provides information on flow uniformity over the interface plane. Because the CCIE is 

not an axisymmetric inlet, the identification of nonuniformities in the local flow field is important to 

assist with test point planning and the interpretation of the flight test results for that program. The 

flow-field data obtained from the RAGE will be used primarily for flight-test planning and computational 

flow analysis of the CCIE, but they are applicable to any other propulsion-related experiment that might 

fly on the PFTF in the future.

A total of two flights were conducted for the RAGE. The first flight, occurring in August of 2009, 

was the primary research flight for this program in which the bulk of the data was obtained. Therefore, the 

data from this flight are discussed in greater detail. A second flight was conducted in October of 2009 to 

spot-check some anomalous data that had been observed at a particular location on the interface plane 

during the first flight. For this flight, the rake was rotated 180° about its longitudinal axis to ensure that 

the data anomaly was not the result of a particular probe. 

For both flights, experimental data were obtained at steady-state supersonic test points and during the 

straight and level acceleration segments out to these points. Because the relationship between free-stream 

and local conditions was not known before the first flight, the desired steady-state test points had to be 

identified in real time during the flight by using the calculated local flow properties, available in the 

control room, to guide the pilot to the target free-stream Mach number that produced the desired local 

Mach number. Flow-field data were obtained primarily at a pressure altitude of 40,000 ft, because this 

altitude allowed for test points to be flown up to M∞ = 2, if necessary, to achieve the desired local flow 

conditions. For the first run on the first flight, a steady-state test point at M∞ = 1.65 was chosen to develop 

an initial correlation between the free-stream and local flow quantities. The data from this first run 

identified M∞ = 1.56 as a desired test point for the second run. On the second run, the airplane accelerated 

out to M∞ = 1.56, held this condition for 10 s, then accelerated out to M∞ = 1.86, at which point the pilot 

was instructed to hold the condition for another 10 s, because the upper Mach limit of the rake’s 

calibration had been reached. For the first flight, a total of four supersonic runs, with aerial refueling in 

between each run, were necessary to obtain test data at a range of supersonic Mach numbers. During this 

flight, test points were also flown at pressure altitudes of 35,000 and 45,000 ft to check for altitude effects 

on the local flow properties. Figure 6 shows a summary of the flight conditions in which data were 

obtained. A total of seven unique steady-state test points over a range of free-stream Mach numbers from 

1.56 to 1.86 were flown. To check for data repeatability, the M∞ = 1.56 and 1.86 test points at 40,000 ft 

were repeated on run 3 of the first flight. For the second flight, the M∞ = 1.56, 1.65, and 1.86 test points at 

40,000 ft were flown for comparison with the data obtained at these conditions on the first flight. Table 2 

shows the aggregate of steady-state test points for both flights.
8

Table 2. Steady-state test points.

Pressure altitude, ft Mach number Run Flight

40,000 1.65 1 1

40,000 1.56 2 1

40,000 1.86 2 1

40,000 1.56 3 1

40,000 1.86 3 1

35,000 1.75 3 1

45,000 1.56 4 1

45,000 1.86 4 1

35,000 1.56 4 1

40,000 1.56 1 2

40,000 1.65 1 2

40,000 1.86 1 2

Figure 6. Rake Airflow Gage Experiment steady-state test points.

B. Flight Test Results

Note that probes 1, 6, and 7 were unintentionally bent by small amounts during the aforementioned 

leak checks of the pressure tubing. The probes were straightened as much as possible; however, 

reorienting them in the exact same configuration as they had been during the wind-tunnel calibration of 

the rake was impossible. Consequently, the flow calculations from these three probes have an additional 

error associated with them. Although the data obtained from these three probes are presented, they are not 

discussed in detail in this report. Also note that the local angle of attack that is discussed in this report is 

measured with respect to the inclined rake. Therefore, it is necessary to subtract 5° to these values to 
9

obtain the total angle of attack underneath the airplane with respect to the longitudinal axis of the 

airplane.

To facilitate presentation of the steady-state data, the calculated flow properties for each probe were 

time averaged over the test point duration. The time-averaged flow properties from the six unbent probes 

(2, 3, 4, 5, 8, and 9) were used to compute the mean flow conditions over the interface plane. In addition 

to the steady-state flow properties, the level acceleration flow-field calculations also are presented. 

Because the pressure transducers are located several feet away from the pressure ports of the probes, the 

pressures measured during the acceleration portion of the run are different from the true pressures because 

of the lag caused by the tubing. Based on the methods given in Gracey (ref. 7) and Huston (ref. 8), the 

pressure lag of the system during the acceleration was estimated to have a mean value of approximately

0.005 psi. Because this value was within the measurement resolution of the differential pressure 

transducers, the pressure lag of the tubing was deemed negligible. It should be noted, however, that the 

calculation method is only rigorously valid for straight lengths of tubing. As a result, the acceleration data 

should be considered an approximation of the behavior of the local flow properties during those 

maneuvers.

C. Local Mach Number, Total Pressure Distortion, and Dynamic Pressure 

The local Mach number measured during the RAGE flights was expected to be less than the 

free-stream value as a result of the airplane flying faster than the speed of sound. At supersonic speeds, 

the airplane surfaces, particularly the nose and inlet ramps, produce oblique shock waves that alter the 

flow underneath the airplane. As the free-stream flow passes through these shock waves, it undergoes a 

reduction in Mach number and is turned away from the airplane underbody. In addition to the 

airplane-generated shock waves, several other factors, including but not limited to, variable inlet spillage, 

dissimilar trim settings between port and starboard engines, and airplane attitude are thought to affect the 

local flow conditions underneath the airplane. 

1. Steady-State Results

Tables 3–7 present the time-averaged steady-state data for both flights. The local Mach number, total 

pressure, and dynamic pressure for the unbent probes were used to calculate the mean and standard 

deviation of these quantities over the interface plane. For the first flight, the 40,000-ft test points, shown 

in table 3, indicate that the average local Mach number at the inlet interface plane varies from 1.48 to 1.60 

over a free-stream Mach range of 1.56–1.86. The data in table 3 indicate that a free-stream Mach number 

of 1.65 will achieve the desired local Mach number of 1.5 for the CCIE. Small deviations from the mean 

Mach number of up to 0.03 are apparent from a probe-to-probe comparison. The variances in local Mach 

number between the probes show that the flow at the interface plane is not entirely uniform. As a check of 

the data repeatability, the M∞ = 1.56 and 1.86 test points at 40,000 ft were repeated on the third run of 

flight 1. The average local Mach number from the repeated test points showed excellent agreement 

between the two runs, as seen in table 4.
10

Table 3. Rake Airflow Gage Experiment flow-field summary (flight 1, 40,000 ft).

M∞ = 1.56

Probe ML α, deg αf, deg pt, psi q , psf

1* 1.49 –0.60 0.22 10.96 678

2 1.48 –0.81 –0.11 10.97 679

3 1.48 –0.01 –0.68 10.94 677

4 1.49 –1.30 0.25 10.95 677

5 1.50 –1.40 –0.19 10.99 679

6* 1.47 0.67 1.85 11.03 683

7* 1.49 –1.12 –0.77 10.95 677

8 1.48 –1.15 0.15 10.84 671

9 1.46 –0.53 0.87 10.89 675

 m 1.48 –0.87 0.05 10.93 676

s 0.01 0.53 0.52 0.06 3

Distortion 1.4%

M∞ = 1.65

Probe ML α, deg αf, deg pt, psi q , psf

1* 1.51 –1.63 1.05 12.33 761

2 1.50 –2.73 0.12 12.41 767

3 1.52 –0.83 –0.35 12.37 763

4 1.52 –1.75 0.16 12.27 757

5 1.52 –2.06 –0.22 12.39 764

6* 1.49 –1.36 1.78 12.50 773

7* 1.51 –2.41 –0.30 12.37 763

8 1.51 –2.07 0.01 12.20 753

9 1.47 –1.05 0.20 12.24 758

 m 1.50 –1.75 –0.01 12.31 760

s 0.02 0.71 0.22 0.09 5

Distortion 1.7%

M∞ = 1.86

Probe ML α, deg αf, deg pt, psi q , psf

1* 1.62 –3.34 1.30 16.43 993

2 1.62 –4.82 0.27 16.68 1009

3 1.62 –3.26 0.07 16.46 995

4 1.59 –3.31 –0.28 16.10 979

5 1.61 –4.19 0.18 16.46 998

6* 1.60 –2.82 1.28 16.79 1019

7* 1.63 –4.37 0.28 16.67 1004

8 1.58 –3.92 0.22 15.75 961

9 1.58 –3.19 –0.06 16.38 998

 m 1.60 –3.78 0.07 16.30 990

s 0.02 0.65 0.20 0.33 17

Distortion 5.7%

* Denotes bent probe
11

Table 4. Rake Airflow Gage Experiment flow-field summary (flight 1, 40,000 ft, repeated test points).

M∞ = 1.56

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.48 –0.65 0.29 10.93 677

2 1.48 –0.96 –0.11 10.96 678

3 1.48 0.05 –0.66 10.93 676

4 1.49 –1.18 0.25 10.95 677

5 1.50 –1.38 –0.04 10.97 678

 6* 1.47 0.71 2.01 11.01 682

 7* 1.49 –1.05 –0.82 10.94 677

8 1.48 –1.06 0.17 10.82 670

9 1.46 –0.49 0.93 10.89 675

m 1.48 –0.84 0.09 10.92 676

s 0.01 0.53 0.52 0.06 3

Distortion 1.4%

M∞ = 1.86

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.61 –3.26 1.26 16.08 974

2 1.61 –4.73 0.30 16.30 987

3 1.61 –3.12 0.09 16.10 975

4 1.58 –3.32 –0.33 15.76 962

5 1.60 –4.10 0.14 16.13 979

 6* 1.60 –2.67 1.33 16.51 1002

 7* 1.62 –4.33 0.28 16.26 982

8 1.57 –3.88 0.25 15.44 944

9 1.58 –3.12 –0.16 16.15 985

m 1.59 –3.71 0.05 15.98 972

s 0.02 0.64 0.24 0.32 17

Distortion 5.4%

* Denotes bent probe
12

Table 5. Rake Airflow Gage Experiment flow-field summary (flight 1, 45,000 ft).

M∞ = 1.56

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.51 –0.67 0.81 8.73 539

2 1.49 –0.51 0.37 8.72 539

3 1.49 –0.27 –0.41 8.69 538

4 1.51 –0.99 0.09 8.73 538

5 1.51 –1.19 0.39 8.73 539

 6* 1.48 0.55 1.85 8.77 543

 7* 1.50 –1.22 –0.40 8.70 538

8 1.49 –0.89 0.54 8.60 532

9 1.47 –0.20 1.06 8.66 537

m 1.49 –0.67 0.34 8.69 537

s 0.01 0.41 0.49 0.05 3

Distortion 1.5%

M∞ = 1.86

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.61 –3.07 1.60 12.37 747

2 1.60 –4.17 0.28 12.50 756

3 1.60 –2.71 0.28 12.38 748

4 1.58 –3.29 0.00 12.12 736

5 1.60 –3.60 0.28 12.36 747

 6* 1.57 –2.30 1.48 12.20 743

 7* 1.61 –3.59 0.42 12.60 759

8 1.56 –3.90 0.42 11.86 723

9 1.56 –2.76 0.41 12.23 746

m 1.58 –3.41 0.28 12.24 743

s 0.020 0.597 0.152 0.229 11

Distortion 5.2%

* Denotes bent probe
13

Table 6. Rake Airflow Gage Experiment flow-field summary (flight 1, 35,000 ft).

M∞ = 1.56

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.48 –0.76 0.31 13.64 845

2 1.47 –1.16 –0.06 13.67 847

3 1.48 –0.11 –0.43 13.67 846

4 1.48 –1.09 0.64 13.61 843

5 1.48 –1.46 0.06 13.68 847

 6* 1.47 0.16 1.84 13.79 854

 7* 1.48 –1.61 –0.19 13.68 847

8 1.48 –1.35 0.05 13.52 837

9 1.46 –0.48 0.10 13.61 844

m 1.47 –0.94 0.06 13.63 844

s 0.01 0.53 0.34 0.06 4

Distortion 1.2%

M∞ = 1.75

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.53 –2.93 1.20 17.34 1067

2 1.54 –3.15 0.07 17.48 1073

3 1.53 –3.01 0.00 17.33 1067

4 1.52 –2.07 0.35 17.15 1057

5 1.54 –3.27 0.13 17.50 1074

 6* 1.56 –1.42 1.63 17.71 1083

 7* 1.56 –3.29 0.52 17.42 1067

8 1.53 –3.06 0.42 17.03 1048

9 1.53 –2.33 0.13 17.54 1080

m 1.53 –2.81 0.18 17.34 1066

s 0.01 0.49 0.17 0.21 12

Distortion 2.9%

* Denotes bent probe
14

Table 7. Rake Airflow Gage Experiment flow-field summary (flight 2, 40,000 ft).

M∞ = 1.56

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.49 –0.94 0.11 10.82 669

2 1.49 –1.20 –0.50 10.74 665

3 1.49 –1.42 0.56 10.82 669

4 1.49 –1.43 0.01 10.87 672

5 1.49 –0.71 –0.83 10.78 667

 6* 1.48 –2.03 –1.26 10.65 660

 7* 1.49 –0.37 1.41 10.81 669

8 1.46 –0.80 0.05 10.83 671

9 1.48 –0.57 –1.68 10.78 668

m 1.48 –1.02 –0.40 10.80 669

s 0.01 0.38 0.79 0.04 3

Distortion 1.1%

M∞ = 1.65

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.51 –2.40 –0.55 12.11 747

2 1.50 –1.99 –0.27 11.93 737

3 1.52 –2.42 1.18 12.15 749

4 1.49 –3.42 0.28 12.13 750

5 1.50 –1.42 –1.00 12.04 744

 6* 1.50 –2.45 –1.23 11.87 733

 7* 1.50 –1.84 0.68 12.10 748

8 1.47 –2.16 0.26 12.13 752

9 1.50 –1.60 –1.45 12.07 746

m 1.50 –2.17 –0.17 12.08 746

s 0.02 0.71 0.95 0.08 5

Distortion 1.9%

M∞ = 1.86

Probe ML α, deg αf, deg pt, psi q , psf

 1* 1.64 –4.46 –0.18 16.27 979

2 1.60 –3.43 0.41 15.75 957

3 1.59 –4.99 –0.17 16.03 975

4 1.61 –4.46 0.14 16.33 989

5 1.62 –3.43 –0.70 16.18 977

 6* 1.57 –4.33 –1.44 15.37 939

 7* 1.61 –3.65 0.43 16.22 983

8 1.59 –3.66 0.02 16.43 1001

9 1.62 –3.81 –0.38 16.35 988

m 1.60 –3.96 –0.11 16.18 981

s 0.02 0.63 0.39 0.25 15

Distortion 4.2%

* Denotes bent probe
15

Between M∞ = 1.56 and 1.86, the mean local total pressure, pt, at the interface plane ranged from 

approximately 10.9 to 16.3 psi, with small deviations apparent between the individual probes. The total 

pressure deviations are representative of distortion in the local flow field. Equation (1) was used as a 

simple measure of the local distortion, and the distortion was calculated to have a range of 1.4–5.7 percent 

over ML = 1.48–1.60. At ML = 1.5, the distortion is less than 2 percent, which is an acceptable value for 

the CCIE. Interestingly, there appears to be a total pressure deficit of 0.55 psi at the top of the rake 

(probe 8) during the M∞ = 1.86 test point. This total pressure deficit is the primary reason for the elevated 

level of distortion at this Mach number. 

The local dynamic pressure at the 40,000-ft steady-state test points ranged from approximately 680 to 

990 psf. These values are well within the PFTF limit of 1,100 psf, suggesting that higher local Mach test 

points could be achieved at 40,000 ft, if necessary, for the CCIE or other potential future experiments. As 

previously mentioned, the target dynamic pressure for the CCIE at its design speed of ML = 1.5 is 

1,000 psf. This value was determined to be required to provide a specific mass flow to a conceptual 

RBCC engine that is no longer in development. At ML = 1.5, the measured dynamic pressure is 

approximately 760 psf. Because dynamic pressure has a second-order effect on the performance of the 

inlet, the repercussions of flying the ML = 1.5 test point at this lower dynamic pressure should be 

negligible. Alternatively, it would be possible to fly the CCIE at a greater local dynamic pressure by 

flying at the same local Mach number, but at a lower pressure altitude.

During the first flight, additional test points were flown at 45,000 and 35,000 ft to check for altitude 

effects on the local flow conditions at the interface plane. Tables 5 and 6 show the results from the 

steady-state test points for the pressure altitude runs at 45,000 and 35,000 ft, respectively. The data 

obtained at these altitudes show that the local Mach number and distortion for a given free-stream Mach 

number are effectively the same among the three altitudes flown during the flight test. In contrast to local 

Mach number and distortion, the local dynamic pressure between the different pressure altitudes will not 

be the same, as it is dependent on the free-stream pressure, and thus the pressure altitude. Because of the 

dynamic pressure limits on the PFTF, the maximum free-stream Mach number at 35,000 ft was limited to 

1.75. As shown in table 5, the average local Mach number at this condition was 1.53. The total pressure 

distortion was 2.9 percent with a mean dynamic pressure of 1,066 psf. This particular test point represents 

the highest local Mach number that experiments can be flown using the PFTF at 35,000 ft. 

The second flight consisted of one supersonic run with steady-state test points at M∞ = 1.56, 1.65, and 

1.86 at 40,000 ft. Table 7 presents the results from this flight. A comparison of tables 3 and 7 shows that 

the average local Mach numbers were in excellent agreement with those measured during the first flight. 

The total pressure distortion at the two lower Mach test points were in good agreement. At M∞ = 1.86, the 

total pressure distortion was calculated to be 4.5 percent, some 1.2 percent lower than that obtained from 

the first flight. A likely reason for this sizeable difference is that the rake was inverted for the second 

flight. Inverting the rake placed probe 6 at the top location on the rake, a region where the largest total 

pressure deficit was seen during the first flight. Since probe 6 was one of the bent probes, the data were 

not included in the distortion calculation. In table 8, the mean interface plane flow quantities from both 

flights were averaged to produce a final correlation between airplane free-stream Mach number and the 

local flow conditions at the aerodynamic interface plane at 40,000 ft. The correlation from this table 

should prove useful in the flight test planning of the CCIE and future PFTF projects.

Distortion % = p p

p

t t

t avg

,max ,min

,

£ <

¤

2

¥

¦

 ́ × 100 (1)
16

Table 8. Average interface plane quantity from flights 1 and 2.

40,000 ft

M∞= 1.56 M∞ = 1.65 M∞ = 1.86

ML 1.48 1.50 1.60

α, deg –0.94 –1.96 –3.87

αf, deg –0.17 –0.09 –0.02

pt, psi 10.87 12.19 16.24

q , psf 672 753 986

2. Level Acceleration Results

As discussed previously, the pressure lag during the acceleration portion of the supersonic runs was 

estimated at approximately 0.005 psi. This value was less than the noise level of the pressure transducers, 

making it reasonable to examine the acceleration data without correcting the measured pressures for lag. 

Figure 7 shows a comparison between flights 1 and 2 of the mean interface plane Mach number (fig. 7a) 

and local dynamic pressure (fig. 7b) as a function of airplane free-stream Mach number. The local Mach 

number and dynamic pressure generally show good agreement between the two flights. The relationship 

between local and free-stream Mach number is fairly linear with a change in slope at approximately 

M∞ = 1.74. Another item of interest from figure 7a is the sharp drop off of approximately 0.02 in the local 

Mach number as the engine power is reduced for the M∞ = 1.86 steady-state test point. An examination of 

the pressure data from the probes showed what appeared to be significant pressure rises of varying 

magnitude on all of the probes as the engine power setting was reduced to hold the condition at the 

steady-state test point. After the engine trimmed at the reduced power setting, the probe pressures fell to 

levels lower than they had previously attained before the reduction in the power setting. This reduction in 

pressure resulted in a decrease of approximately 0.02 in the local Mach number as seen on the plot. This 

behavior was observed to a smaller extent at the M∞ = 1.65 test point yet was absent at the M∞ = 1.56 test 

point. 
17

Figure 7a. Mach number.

Figure 7b. Dynamic pressure.

Figure 7. Average local Mach number and dynamic pressure during the 40,000-ft supersonic acceleration 

runs. 
18

D. Local Flow Angularity

The airplane surfaces, particularly the nose and variable position inlet ramps, generate oblique shock 

waves when the airplane is traveling faster than the speed of sound. These shock waves turn the flow and 

alter the local flow angularity underneath the airplane. Inlet spillage, which is a function of free-stream 

Mach number, also is thought to affect the flow angularity underneath the airplane. This section provides 

a detailed description of the results of the local flow angularity measurements.

1. Steady-State Results

Tables 3–7 provide the time-averaged local flow angularity measurements for the steady-state test 

points. As discussed previously, the local angle of attack was measured with respect to the experiment, 

which was inclined 5° above a horizontal plane parallel to the waterlines of the airplane. In general, the 

data showed the presence of a negative angle of attack or downwash on the rake at all flight conditions in 

which the rake was within its Mach number calibration range. This downwash is a result of the 

free-stream flow being turned away from the underbody of the airplane as it passes through the oblique 

shock waves generated by the airplane. As the free-stream Mach number changes, the angles of the 

oblique shocks change, which results in the local angle of attack being a function of the free-stream Mach 

number. Moreover, the inlet ramps are scheduled to provide the appropriate mass flow to the engines over 

a wide range of Mach numbers. Inlet ramp deflection and spillage, which are functions of free-stream 

Mach number, also help create the variable flow angularity underneath the airplane. 

Table 3 shows the flow angularity results from flight 1 at a pressure altitude of 40,000 ft. During the 

supersonic run, the downwash underneath the airplane increased with free-stream Mach number. From a 

free-stream Mach number of 1.56 to 1.86, the average local angle of attack at the interface plane varied 

from –0.87° to –3.8°, while the average local flank angle was essentially 0°. At the CCIE design Mach 

number of ML = 1.5, the local downwash is approximately –1.75°. Although not optimal, this level of 

downwash is still an acceptable level for the CCIE. Tables 5 and 6 show the time-averaged flow 

angularity for the pressure altitude points at 45,000 and 35,000 ft, respectively. Overall, the data obtained 

at the M∞ = 1.56 and 1.86 test points showed no significant differences from those obtained at the same 

free-stream Mach numbers at 40,000 ft. The 35,000-ft, M∞ = 1.75 test point resulted in a mean local angle 

of attack and flank angle of 2.8° and 0.2°, respectively. 

Table 7 provides the flow angularity data from flight 2. Compared with flight 1, for a given 

free-stream Mach number, the mean local angle of attack measured during flight 2 was more negative; 

that is, the downwash was greater. The biggest discrepancy between the two flights was at the M∞ = 1.65 

test point, in which the mean local downwash on the second flight was approximately 0.4° greater than 

that measured during the first flight. Potential sources of these discrepancies are discussed in section 

III-D-3, “Sources of Flow Angularity Discrepancies Observed Between Flights.” 

To better understand the flow behavior from over the interface plane, contour plots of both angle of 

attack and flank angle were generated. Figures 8–11 show the flow angularity contours for flight 1. In the 

plots, the unbent RAGE probes are shown as black circles, and the bent probes are represented by red 

circles. The inlet cowl of the follow on experiment, the CCIE, is shown as a dashed black line. It was 

necessary to include the data from the bent probes in the generation of these plots because of the 

algorithm that was used to extrapolate the probe data over the entire inlet area. The flow angularity in the 

region of the bent probes should be viewed with caution for the reasons discussed previously. The figures 

show that both the angle of attack and flank angle are not uniform across the interface plane, and that the 

flow angularity is mostly downwash that increases with Mach number. Interestingly, these figures show a 

reduced region of downwash at probe 3, located directly starboard to the center probe. This downwash 

deficit, compared to the other probes, was the reason for conducting the second flight with the rake 

inverted, thereby placing a different probe in that same area of interest. For comparison with the first 
19

flight, figures 12–14 show the flow angularity contours from the second flight. Although the absolute 

values of flow angularity were different between flights 1 and 2, the angularity trends were similar, 

particularly when the downwash was considered. The downwash deficit that was seen on the first flight 

was still present at the location starboard of the center probe. The variable flow angularity over the 

interface plane, particularly at the location starboard to the center probe, is not optimal, but it is not 

thought to be excessively detrimental to the future flight test of the CCIE.

Figure 8a. Angle of attack.

Figure 8b. Flank angle of attack.

Figure 8. Flow angularity contours at a free stream Mach number of 1.56 at 40,000 ft (flight 1).
20

Figure 9a. Angle of attack.

Figure 9b. Flank angle of attack.

Figure 9. Flow angularity contours at a free stream Mach number of 1.65 at 40,000 ft (flight 1).
21

Figure 10a. Angle of attack.

Figure 10b. Flank angle of attack.

Figure 10. Flow angularity contours at a free stream Mach number of 1.86 at 40,000 ft (flight 1).
22

Figure 11a. Angle of attack.

Figure 11b. Flank angle of attack.

Figure 11. Flow angularity contours at a free stream Mach number of 1.75 at 35,000 ft (flight 1).
23

Figure 12a. Angle of attack.

Figure 12b. Flank angle of attack

Figure 12. Flow angularity contours at a free stream Mach number of 1.56 at 40,000 ft (flight 2).
24

Flight 13a. Angle of attack.

Flight 13b. Flank angle of attack.

Figure 13. Flow angularity contours at a free stream Mach number of 1.65 at 40,000 ft (flight 2).
25

Flight 14a. Angle of attack.

Flight 14b. Flank angle of attack.

Figure 14. Flow angularity contours at a free stream Mach number of 1.86 at 40,000 ft (flight 2).

2. Level Acceleration Results

Figure 15 shows the mean interface plane angle of attack and flank angle from both flights as a 

function of free-stream Mach number during an entire supersonic run. In figure 15a, the plot for flight 1 

shows a linear relationship between mean local angle of attack and free-stream Mach number. Flight 2 

displays a linear relationship as well, albeit with a different slope, up to approximately M∞ = 1.7 where the 
26

slope of the curve changes. The plots of mean local flank angle for both flights (fig. 15b) almost resemble 

mirror images of each other, and the angles were both within ± 0.5° for the entire supersonic run. 

Figure 15a. Angle of attack.

Figure 15b. Flank angle of attack.

Figure 15. Average flow angularity during the 40,000-ft supersonic acceleration runs.
27

3. Sources of Flow Angularity Discrepancies Observed Between Flights

An exhaustive analysis of the flight data was performed to try to determine the cause of the flow 

angularity discrepancies between flights. Several factors were determined to have had the potential to 

alter the local flow underneath the airplane. During flight 2, the airplane flew at roughly 0.25° of 

additional sideslip compared to flight 1. It was also noted that the free-stream temperature at altitude was 

10°F cooler during flight 2. The colder, denser air allows the engines to operate more efficiently. The 

precise way in which this temperature difference affects inlet spillage is not completely understood, but it 

seems reasonable to believe that the spillage between the two flights was different, thus potentially 

altering the flow underneath the airplane. Lastly, two factors from the rake’s calibration potentially 

affected the local flow measurements between the two flights as a result of the rake being rotated 180° for 

the second flight. As described in reference 5, each probe was not calibrated individually; instead, the 

rake was calibrated as a unit in a wind tunnel. Nonuniformities in the wind tunnel flow speed and 

direction had unquantifiable effects on the calibration of all of the probes, except possibly the center 

probe, because this probe was the only one that was fixed in space in the tunnel as the calibration data 

were obtained. Additionally, the calibration data that were obtained in the wind tunnel were insufficient to 

entirely eliminate misalignment bias between the rake and the flow in the tunnel. 

IV. Concluding Remarks

The flow-field survey rake from the Rake Airflow Gage Experiment (RAGE), comprised of nine 

five-hole conical probes mounted in a cruciform configuration on the Propulsion Flight Test Fixture, was 

flown on the NASA Dryden F-15B research airplane in the latter half of 2009. The goal was to quantify 

the flow field underneath the F-15B airplane at the location of the aerodynamic interface plane of the 

Channeled Centerbody Inlet Experiment (CCIE), a novel supersonic inlet design scheduled for flight test 

at the National Aeronautics and Space Administration Dryden Flight Research Center in the near future. 

Steady-state test points were taken at pressure altitudes of 35,000, 40,000, and 45,000 ft over a range of 

free-stream Mach numbers, M∞, up to 1.86. The local flow angularity, Mach number, total pressure 

distortion, and dynamic pressure were measured and correlated to airplane free-stream conditions. 

Of primary importance to the RAGE was the determination of the free-stream Mach number that 

would produce a local Mach number, ML, of 1.5 at the aerodynamic interface plane of the inlet. The data 

obtained from these flights suggest that a free-stream Mach number of 1.65 will produce ML = 1.5. At this 

condition, the rake, already inclined 5° from horizontal, is not aligned with the flow, as there is an 

average of nearly 2° of downwash on the rake. Because axial flow is desired, it might be possible to fly 

the CCIE at a lower Mach number, perhaps ML = 1.48, corresponding to M∞ = 1.56. This condition would 

decrease the downwash to approximately 1°. A reduction in the local Mach number, however, also has the 

effect of reducing the local dynamic pressure, which is important, but not nearly as critical as the local 

Mach number to the performance of the inlet. Although the RAGE was conducted primarily in support of 

the CCIE, the data obtained in these flights have value to any future inlet experiments that utilize the 

Propulsion Flight Test Fixture. 
28

References

1. Corda, Stephen, M. Jake Vachon, Nathan Palumbo, Corey Diebler, Ting Tseng, Anthony Ginn, and 

David Richwine, “The F-15B Propulsion Flight Test Fixture: A New Flight Facility for Propulsion 

Research,” NASA/TM 2001-210395, 2005.

2. Vachon, Michael Jacob, Timothy R. Moes, and Stephen Corda, “Local Flow Conditions for Propulsion 

Experiments on the NASA F-15B Propulsion Flight Test Fixture,” NASA/TM 2005-213670, 2005.

3. Palumbo, Nathan, Tim Moes, and M. Jake Vachon, “Initial Flight Tests of the NASA F-15B 

Propulsion Flight Test Fixture,” AIAA 2002-4131, Proceedings of the 38th AIAA/ASME/SAE/ASEE 

Joint Propulsion Conference and Exhibit, Indianapolis, Indiana, July 7-10, 2002.

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Propulsion Conference and Exhibit, Indianapolis, Indiana, July 7-10, 2002.

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Sciences Meeting Including The New Horizons Forum and Aerospace Exposition, Orlando, Florida, 

January 5-8, 2009.

6. Miller, K. J., J. C. Sisco, B. L. Austin, Jr., T. N. Martin III, W. E. Anderson, and Nathan M. Palumbo, 

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2003-0390FB, 39th JANNAF Combustion Subcommittee Meeting, Colorado Springs, Colorado, 

December 1–5, 2003. 

7. Gracey, William, “Measurement of Aircraft Speed and Altitude,” NASA RP-1046, 1980.

8. Huston, Wilber B., “Accuracy of Airspeed Measurements and Flight Calibration Procedures,” 

NACA TR-919, 1948.