April 2011 

NASA/CR–2011-217077 

Measured Sonic Boom Signatures Above and 

Below the XB-70 Airplane Flying at Mach 1.5 

and 37,000 Feet 

 

Domenic J. Maglieri and Herbert R. Henderson 

Eagle Aeronautics, Inc., Hampton, Virginia 

Ana F. Tinetti 

NCI Information Systems, Inc., Hampton, Virginia 
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April 2011 

NASA/CR–2011-217077 

Measured Sonic Boom Signatures Above and 

Below the XB-70 Airplane Flying at Mach 1.5 

and 37,000 Feet 

 

Domenic J. Maglieri and Herbert R. Henderson 

Eagle Aeronautics, Inc., Hampton, Virginia 

Ana F. Tinetti 

NCI Information Systems, Inc., Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

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Hanover, MD 21076-1320 

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National Aeronautics and Space Administration. 
 1

Measured Sonic Boom Signatures Above and Below the XB-70 Airplane

Flying at Mach Number 1.5 and 37,000 Feet

By Domenic J. Maglieri and

Herbert R. Henderson and Ana F. Tinetti

ABSTRACT

During the 1966-67 Edwards Air Force Base (EAFB) National Sonic Boom Evaluation Program,

a series of in-flight flow-field measurements were made above and below the USAF XB-70 using

an instrumented NASA F-104 aircraft with a specially designed nose probe. These were accomplished on three XB-70 flights conducted at a Mach number of about 1.5 at an altitude of about

37,000 feet and at a gross weight of about 350,000 pounds. A total of six supersonic passes were

made with the F-104 probe aircraft through the XB-70 shock flow-field; one above the XB-70 on

the first flight, two below the XB-70 on the second flight, and three below the XB-70 on the third

flight. Separation distances ranged from about 3000 feet above and 7000 feet to the side of the

generating aircraft and about 2000 feet and 5000 feet below the generating aircraft. Complex

near-field “sawtooth-type” signatures were observed in all cases. In fact, at ground level, the XB70 shock waves had not coalesced into the two-shock classical sonic boom N-wave signature, but

contained three shocks.

The purpose of these in-flight measurements was to gather an additional database on a very large

and heavy aircraft to be used in providing a check on and improvement to the generalized theory

for predicting sonic boom signatures. Although the tests were successfully completed, the results

were never formally documented appearing, only briefly, in a few reports to reflect the nature of

the flight tests.

The present report documents the results of the XB-70/F-104 probe flight tests and is based upon

file copies of most of the original information and database developed in the 1966-67 time period.

Included in this report is a description of the generating and probe airplanes, details of the inflight and ground pressure measuring instrumentation, the flight test procedure and the aircraft inflight and ground pressure measuring instrumentation, the flight test procedure and aircraft positioning, surface and upper air weather observations and the six measured in-flight pressure time

histories from the three XB-70 flights along with the corresponding ground measured signatures.
2 

THIS PAGE INTENTIONALLY LEFT BLANK 
 3

TABLE OF CONTENTS

INTRODUCTION - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 7

SYMBOLS - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 8

APPARATUS AND METHODS- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 9

Generating and Probe Airplanes- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 9

Pressure Measuring Instrumentation - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -10

Flight Test Procedures - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -10

Aircraft Positioning - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 11

Weather Observations - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 12

RESULTS - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 12

Inflight Measurements - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 12

Wave Shapes - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 12

Peak Positive Overpressures - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 14

Signature lengths- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 15

Ground Measurements - - - -- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 15

Microphone Set-Up and Characteristics - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 15

Wave Shapes - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 16

Signature Characteristics - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 16

Correlation with Airplane Geometry - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 16

SUMMARY REMARKS- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 17

APPENDIX A

DESCRIPTION AND CALIBRATION OF PRESSURE INSTRUMENTATION - - - - - - 18

APPENDIX B

DESIGN AND AERODYNAMICS CALIBRATION OF PRESSURE PROBE - - - - - - - 20

Basic Considerations - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 20

Present Applications - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 20

WIND TUNNEL TESTS - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 21

Introduction - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 21

Symbols - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 21

Apparatus and Tests - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 22

Test facility and conditions- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 22

Test apparatus and procedures- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 22

Measurements - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 23

Data and Precision - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 23

Probe Calibration - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 23

Pressure measurements in vicinity of oblique shock wave - - - - - - - - - - - - - - - - - - - - 23

Results and Discussion - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 23

Probe calibration at angle of attack- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 23

Probe capability for sensing pressure changes across an oblique shock wave - - - - - - - 24

Probe reflection factors for correcting in-flight measurements - - - - - - - - - - - - - - - - - 25

General comments- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 25

REFERENCES - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 26
 4

LIST OF TABLES

Table 1 Geometric characteristics of XB-70 airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .29

Table II Summary of XB-70 and F-104 in-flight probe test conditions . . . . . . . . . . . . . . . . . .34

Table III XB-70 flight conditions overhead of ground cruciform microphone array . . . . . . . .35

Table IV Upper air atmospheric data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .36

Table V Surface weather observations at EAFB runway 22/04 . . . . . . . . . . . . . . . . . . . . . . . .39

Table VI Summary of XB-70 maximum positive overpressures, signature lengths and

periods for in-flight probe tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .40

Table VII XB-70 sonic boom signature characteristics measured at ground level with

the six microphone cruciform array . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .41
 5

LIST OF FIGURES

Figure 1 Photograph of XB-70 delta wing supersonic aircraft used as the sonic

boom generator in the present studies. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .42

Figure 2 Three view drawing of XB-70 airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .43

Figure 3 Photograph of XB-70 showing windshield-nose ramp positions . . . . . . . . . . . . . . . .44

Figure 4 Area distribution of XB-70-1 vehicle used as shock-wave generation airplane. . . . .45

Figure 5 F-104 airplane with nose-boom probe installation for measuring the shock

flow-field in the vicinity of the disturbance generating XB-70-1 airplane . . . . . . . . .46

Figure 6 Photographs of the in-flight recording instrumentation mounted in the F-104

access bay . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .47

Figure 7 Principal details and dimensions of full-scale probe used for in-flight measurements

and for wind tunnel tests at a Mach number of about 2.01 . . . . . . . . . . . . . . . . . . . . .48

Figure 8 Adaptor required to mate specially instrumented noise-boom pressure probe

used originally on an F-106 to the NASA F-104 used in present tests . . . . . . . . . . . .49

Figure 9 Schematic of XB-70 and F104 nominal ground tracks showing area in which

probe missions were flown and ground measurements were acquired . . . . . . . . . . . .50

Figure 10 Sketches illustrating general position of probe aircraft and generating aircraft . . . . .51

Figure 11 Copy of November 23, 1966 film trace showing in-flight time histories of

differential pressures measured in flow-field above XB-70 aircraft. . . . . . . . . . . . . .52

Figure 12 Copy of December 12, 1966 film trace showing in-flight time histories of

differential pressures measured in flow-field below the XB-70 aircraft. . . . . . . . . . .53

Figure 13 Copy of December 16, 1966 film trace showing in-flight time histories of

differential pressures measured in flow-field below XB-70 aircraft . . . . . . . . . . . . .54

Figure 14 XB-70 flow-field shock-wave signature overpressures . . . . . . . . . . . . . . . . . . . . 55

Figure 15 Comparison of measured and calculated distance between bow-wave and

tail-wave of generating airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .57

Figure 16 XB-70 measured sonic boom signatures at ground level following in-flight

probe tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .58
 6

Figure 17 Diagrams of waveforms and signature characteristics which represent the

various categories of measured sonic booms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .59

Figure 18 Planform and side views of bomber airplane with time history of pressure signature as

measured above and below airplane. Signature length has been adjusted to made distance between nose and tail shocks approximately the same as the airplane lengths .60

Figure 19 Response characteristics of pressure instrumentation use for in-flight

measurements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .61

Figure 20 Steady-state calibration of flight probe at angles-of-attack from -2.5o to 9.5o, as

obtained from tests in the Langley 4-foot supersonic pressure tunnel (pressure

orifices at bottom of probe). M  p  185 lbs/ft2. . . . . . . . . . . . . . . . . . . . . . . . . .62

Figure 21 Wind-tunnel apparatus and test arrangement for generating and determining

the strength of an axisymmetrical disturbance used in obtaining experimental

evidence concerning the reflection characteristics of the flight probe . . . . . . . . . 63

Figure 22 Flight-probe capability for sensing static-pressure changes across an axisymmetrical

disturbance (bow wave generated by body of revolution), as evidenced by

comparison of probe-indicated, survey-indicated, and estimated pressure

changes across bow-wave, M  5
 1Pawlowski, J. W.; Graham, D. H.; Boccadoro, C.H.; Coen, P.G. and Maglieri, D.J.: “Origins 

and Overview of the Shaped Sonic Boom Demonstration Program,” AIAA 2005-0005, January 

2005. 7 

INTRODUCTION 

Verification of existing and newly-developed sonic boom prediction codes required an experimental 

database consisting of sonic boom pressure signatures measured at ground level and also by probing 

the supersonic flow-field above, below, and to the side of the generating airplane. A majority of the 

ground measurements are acquired at airplane altitude, Mach conditions such that all of the shocks 

within the supersonic flow field have coalesced into the classical N-wave sonic boom signature (ref. 

1-3). In addition, the atmospheric through which these shock waves propagate, especially the lower 

layers, can strongly alter this N-wave shape such that peaked or rounded waveforms are observed 

with the resulting large variations in sonic boom overpressures as compared to “clean” N-waves 

observed under “quiescent” atmospheric conditions (refs. 1-4). Complex near-field sonic boom 

signatures have been measured at ground level for flights of aircraft at very low altitudes where 

multiple shocks are experienced and the signature takes on a “sawtooth” appearance (refs. 5 and 6). 

In-flight flow-field measurements, on the other hand, are not significantly influenced by the 

atmosphere through which they propagate, especially if the separation distances between the 

generating and probing aircraft are small. In addition, these near-field signatures are more complex 

in that they indicate the shock patterns well before they have coalesced into the single bow and tail 

shock typical of the classical far-field N-wave. Thus, they provide for a critical test of the predictive 

codes. 

The in-flight supersonic flow-field database has been gathered over four decades. In 1956 the 

United States Air Force (USAF) conducted flight tests of an F-100 probing below and to the side of 

an F-100 generating airplane at distances of from about 2 to 41 body lengths (ref. 7). In 1960 this 

was followed by a series of in-flight measurements by NASA using an F-100 to probe very close to 

the side of an F-100, F-104 and B-58 airplane at separation distances of from 1 to 8 body lengths 

(ref. 8). In 1963 the USAF and NASA extended this in-flight database by probing the flow-field 

above and below the B-58 with an F-106 which incorporated a specially designed and instrumented 

nose probe at distances of from about 14 to 95 body lengths (ref. 9). During the 1966-1967 Edwards 

Air Force Base (EAFB) National Sonic Boom Evaluation Program, the opportunity was taken to 

acquire near-field signatures above and below the very large and heavy USAF XB-70 airplane using 

a NASA F-104 (ref. 10) airplane equipped with the same specially instrumented nose probe used on 

the F-106/B-58 tests of 1963. Separation distances of from about 10 to 42 body lengths were 

experienced. During the 1995 time period, NASA conducted an extensive series of in-flight probe 

measurements using the SR-71 as the generating airplane and the F-16XL as the probing airplane. A 

significant number of flow-field penetrations were made at distances of from about 5 to 76 body 

lengths below the SR-71 with the F-16XL and the distances to about 200-300 body lengths below 

the SR-71 using an instrumented YO3A subsonic airplane and at ground level (ref. 11). The SR-71 

flight tests were part of the NASA High Speed Research (HSR) Program aimed at establishing the 

technology base for any future High Speed Civil Transport (HSCT). The most recent in-flight probe 

tests were conducted using a USN F-5E aircraft which was reshaped to produce a flat-top sonic 

boom signature at the ground. These measurements were part of the 2003 DARPA Shaped Sonic 

Boom Demonstrator Program (SSBD) and the 2004 NASA Shaped Sonic Boom Experiments 

Program (SSBE).1
 8

With the exception of the 1960 NASA probe tests (ref. 8), the NASA SR-71 in-flight flow-field

data (ref. 11) and the SSBD and SSBE measurements, little use has been made of the 1963 and

1966 probe measurements of the B-58 and XB-70 flow-field signature data in terms of theory validation. This is due, in part, to the lack of sufficient details of the B-58 and XB-70 geometric and

aerodynamic description which has recently been generated (ref. 12) and the availability of the

details of the XB-70 measurements (the B-58 tests are reported in full detail in ref. 9). Although

the XB-70/F104 in-flight probe measurement effort was successfully completed, the results were

never formally documented, appearing only briefly in a few reports in preliminary form (refs. 10,

13-15) to reflect the general nature of the flight test results.

The present report documents the results of the XB-70 and F-104 flight tests and is based upon

file copies of most of the original information and database developed in the 1966-1967 time

period. Included in this report is a description of the generating and probe airplanes, details of the

in-flight and ground pressure measurement instrumentation, the flight tests procedures and aircraft positioning, surface and upper air weather observations, and the six measured in-flight pressure time histories from the three XB-70 flights along with the corresponding ground measured

signatures.

Since the nose probe pressure instrumentation used on the NASA F-104 aircraft is the same as

that used on the USAF F-106 for the NASA-USAF B-58 probe tests (ref. 9), Appendices A and B,

taken from reference 9, are also included in this report. Appendix A, by John F. Bryant, Jr., provides a description and static calibration of the pressure instrumentation, and Appendix B, by Virgil S. Ritchie, gives a detailed description of the unique instrumentation probe used to obtain the

pressure measurements along with the corresponding static and wind tunnel calibrations.

Since the flight experiments, objectives and flight-test techniques of both the 1966 XB-70/F-104

and the 1963 B-58/F-106 probe tests were similar, the report format utilized in the B-58/F-106

report (ref. 9) will be used for the present report. In addition, calculations of the XB-70 sonic

boom overpressures, period, and signature lengths will be made using the predictive scheme

available at the time of the flight tests.

SYMBOLS

A area of XB-70-1 airplane section obtained by oblique cut for a nominal 

Mach number of 1.5, ft.2

h vertical distance from ground to airplane, ft.

l length of bomber airplane, ft.

M airplane Mach number

M differential Mach number between generating and probe airplanes

pmax peak positive overpressures, lbs/ft2
 9

r slant range separation distance between generating airplane and probe airplane,

measured perpendicular to generating-airplane flight track (positive when probe

airplane is below generating airplane), , ft.

S horizontal distance of probe airplane behind generating airplane at penetration, ft.

T time interval between bow and tail shock waves of generating airplane in horizontal

plane, msec

V airplane ground velocity, ft/sec

V differential ground velocity between generating and probe airplane, ft/sec.

W gross weight of XB-70 airplane, lbs.

X distance between bow and tail shock waves of generating airplane in horizontal

plane (signature length), ft.

x axial distance from nose of airplane, ft.

Y lateral separation distance between generating and probe airplanes, ft.

Z vertical separation distance between generating and probe airplanes (positive when

probe airplane is below generating airplane), ft.

 bow shock rise time to maximum overpressure, msec.

1/2 bow shock rise time to one-half maximum overpressure, msec.

 azimuthal position about generating aircraft (defined in fig. 4).

 experimentally determined shock-wave angle of ground level.

APPARATUS AND METHOD

Generating and Probe Airplanes

The USAF XB-70-1 delta-wing airplane, shown on figure 1, was used as the generating vehicle.

A three-view drawing of this airplane is shown in figure 2 and detailed geometric characteristics

are provided in Table I that is taken from reference 16. The aircraft has a length of 189 feet

(including nose boom), a wing span of 105 feet, and a total wing area of 6297.8 square feet. Aircraft weight at brake release for the three probe flights ranged from about 529,000 pounds to

536,000 pounds. During the actual probe runs, the XB-70 gross weight ranged from about

320,000 pounds to 350,000 pounds; wing tips were full down at 65 degrees and the nose ramp

windshield was in the down position (see fig. 3). The bypass was set at 400 square inches and all

engines were at 100 percent RPM and exhaust nozzles were in partial afterburner. The aircraft is

powered by six YJ9-GE-turbojet engines, each producing 31,000 pounds thrust with full after-

  Z

2

  Y

2

+
 10

burner. Calculated area distributions, based on a Mach 1.5 oblique cut for a position above and to

the side of the XB-70 and also positions below the XB-70, are given in figure 4. These area developments correspond to the probe flight measurements contained in this report and were generated

using the vehicle geometric description given in reference 12.

The NASA F-104 airplane, shown in figure 5(a), was used with a specially instrumented nose

boom probe for sensing pressure changes during flight through the flow-field of the XB-70-1 airplane. The special nose-boom pressure probe is shown in figure 5(b) and is the same probe that

was mounted on the USAF F-106 to survey the flow-field above and below the B-58 airplane in

1963 (see ref. 9). Photographs of the in-flight recording instrumentation mounted in the F-104 are

shown in figure 6. Included are a carrier amplifier, NASA recording oscillograph, a temperature

control box, and a NASA timer. Both the USAF XB-70 and NASA F-104 were based at Edwards

Air Force Base, California, the XB-70 being operated by the personnel of the Air Force Flight

Test Center (AFFTC) and the F-104 by the NASA Flight Research Center (FRC), now NASA

Dryden Flight Research Center (DFRC).

Pressure Measuring Instrumentation

The specially instrumented nose-boom probe was designed, fabricated and calibrated by NASA

Langley Research Center personnel. Details of the pressure probe and wind tunnel tests to determine the pressure-sensing characteristics of the probe are described in Appendices A and B of reference 9 and are reproduced in this report for completeness. The general arrangement of main

dimensions of the probe components are illustrated schematically in figure 7. (Symbols in figure 7

are defined in Appendix B.) Two NASA inductive-type miniature pressure gages were contained

in the probe at locations near pressure-sensing orifices. The probe was laboratory-checked, once

again, before installation on the NASA F-104 airplane to reestablish its sensitivity to a vibration

environment. For the present flight tests, the pressure probe was fitted with a conical tip (fig. 7).

An adapter, shown in figure 8, was required in mating the rear portion of the instrumented probe

to the F-104 such that the angle-of-attack of the probe would be near zero degrees for the

expected flight conditions.

Flight Test Procedures

As mentioned previously, the XB-70 probe tests were an attachment to the Phase II part of the

1966-1967 EAFB National Sonic Boom Evaluation Program. The general arrangement of the

probe flight test plan is provided in figure 9. Probing flights were conducted in “piggy-back” fashion during the XB-70’s sonic boom run over the main test area near the west-end of Rogers Dry

Lake where sonic boom ground pressure measurements, building response measurements, and

subjective response studies were being conducted (see ref. 17).

To accommodate both the in-flight probe measurements and the ground test site measurements,

the XB-70 established a steady flight condition of about Mach 1.5 and 37,000 feet MSL about 100

nautical miles east of the ground test site (just east of Soda Lake) and maintained these conditions

flying a nominal 245 degrees magnetic heading (261 deg.true) that brought it over the main test

area. Probe flights were to be conducted above and below the XB-70 during its run-in to the main
 11

test area (see fig. 9). If the probe flights were not completed by the time both aircraft were abeam

of four corners, the F-104 probe aircraft would break off so as not to boom the test area.

The probe flight plan required the F-104 to be at 2000 to 5000 feet above or below the XB-70

flight altitude and at a Mach 1.3 over Soda Lake awaiting the inbound XB-70 flying at about

Mach 1.5 to overtake it. Once the aircraft slipped back through the XB-70 flow field, the F-104

probe aircraft would accelerate to about Mach 1.7 and hold this speed as it penetrated the XB-70

flow-field from rear to front. After completing the penetration, the F-104 reduces to about Mach

1.3 and slides back through the XB-70 flow-field. This sequencing of three surveys of the XB-70

flow field takes about 3 to 4 minutes and some 60 to 70 nautical miles. The speed and altitude of

the XB-70 and F-104 were held as steady as possible during penetration; however, the F-104 did

experience small variations in velocity which ranged about +5ft./sec. A maximum of three and a

minimum of one F-104 probe penetrations were possible depending upon aircraft coordination

positioning and fuel remaining.

The probe pressure-measurement system on the instrumented airplane was kept inert from the

time of takeoff until steady flight conditions were established (see Appendix A for details). Just

prior to penetration of the pressure field of the generating airplane, the pilot of the F-104 probe

airplane was instructed by radio to activate the pressure measurement system. In addition, The F104 pilot transmitted a timing signal to the ground tracking station both prior to and subsequent to

penetration. This timing signal was superimposed on the tracking data and the data record of the

flight recorder.

Aircraft Positioning

Positioning and guidance of both the XB-70 and F-104 aircraft was accomplished using two

USAF-FPS-16 precision radars operated by AFFTC personnel located at the EAFB radar facility

(SPORT). A radar transponder was located on the bottom of the XB-70 and 102 feet behind the

nose. The F-104 transponder was located about 10 feet behind the tip of the probe. Accuracies in

range, velocity and acceleration were quoted as 40 feet, 10 ft/sec, and 5 ft/sec2, respectively.

All six (6) probe runs were conducted with the XB-70 flying at nominal conditions of about Mach

1.5 and 37,000 feet MSL and a gross weight of about 340,000 pounds. Figure 10 provides

sketches illustrating the general position of the probe aircraft and generating aircraft for the six

probe runs. The XB-70 and F-104 coordinate systems at time of penetration is shown in figure

10(a) where S, Y and Z are the distances that the F-104 is behind, to the side, and above or below

the XB-70. The distance, r, is the slant range distance of the F-104 from the XB-70 flight axis (for

y = 0, r =z). Figure 19(b) provides a rear view, looking in the flight direction, showing the six F104 probe runs relative to the XB-70 flow field. Note that although the intent was to position the

F-104 directly above and directly below the XB-70, only four of the six penetrations were within

less than 1000 feet of the vertical. Run 1 of the flight test date, December 16, 1966, was about

3000 feet to the side and the single penetration of flight test date, November 23, 1966, was about

7000 feet to the side of the XB-70 flight track.

A summary of the XB-70 and F-104 in-flight probe test conditions is given in Table II for each of

the three test dates for all six probe penetrations. Included is the time of penetration, the XB-70
 12

Mach, altitude, heading, weight, ground speed, canard and elevon positions, the F-104 Mach and

altitude, heading and ground speed. Also shown are the S, Y, and Z and r distances of the F-104

from the XB-70 at time of shock penetrating and the closure rate between the XB-70 and F-104

(M and V). The XB-70 pilots experienced light to moderate turbulence on all three probe test

dates, and only during one of the six penetrations (Pass 1 on Dec. 16, 1966) did the XB-70 pilot

and co-pilot experience a slight bump from the F-104 shock field.

Table III provides a summary of the XB-70 flight conditions overhead of the ground cruciform

microphone array at the main test site for the three probe flight test days. The time at overhead of

the ground measurement site is some 2 to 5 minutes after the probe flight penetrations. Comparison of the XB-70 flight parameters of Tables I and III indicate the vehicle was essentially at

steady-level conditions for the entire run in from Soda Lake some 100 miles east to the main test

site at EAFB. The XB-70 ground velocities given in Table III were obtained from the ground

microphone cruciform array.

Weather Observations

Both surface and upper air weather observations were made during the 1966-67 EAFB National

Sonic Boom Evaluation Program. Rawinsonde observations from the EAFB weather facility

include measured values of pressure, temperature, relative humidity and wind speed and direction

from near the surface to altitudes well in excess of the aircraft flight altitudes. The Edwards upper

air atmospheric data, as archived by the National Climatic Data Center in Asheville, North Carolina, for the days (Nov. 23, Dec. 12 and Dec. 16, 1966) on which the XB-70 probe measurements

were conducted is presented in Table IV. The aircraft headings were such that slight to moderate

headwinds were encountered at flight altitude on all three XB-70 probe flights.

A summary of the surface weather conditions observed at the Edwards runway 22/04 on the three

test days and at the approximate time the XB-70 was overhead of the main test area cruciform

microphone array is presented in Table V. Over the three test day mornings, cloud cover ranged

from clear to overcast with no precipitation. Winds were calm and temperatures ranged from 27oF

to 47oF.

RESULTS

Both the in-flight probe and ground measurements of the shock wave pressure signatures of the

XB-70 are presented. Figures 11-15 and Table VI relate to the measured signatures obtained from

the six in-flight probe runs and figures 16-17 and Table VII address the measured signatures at the

two ground sites for the three XB-70 flights. A correlation of the in-flight and ground pressure

time histories with airplane geometry is shown in figure 18.

In-Flight Measurements

Wave shapes. - A copy of the F-104 probe aircraft film trace showing the in-flight time histories

of differential pressures measured on all six penetrations on the three XB-70 flights of November

23, December 12, and December 16, 1966 are presented in figures 11, 12, and 13, respectively. In

each case, the top pressure trace was obtained with gage 1, which was connected to the forward
 13

orifices on the measuring boom, whereas the bottom trace was obtained with gage 2, which was

connected to the rearward orifices (see fig. 7). The two pressure traces are not directly comparable

in amplitude because of differences in the sensitivities of the gages and in the reflection factors for

the probe at the orifice locations and, possibly, because of effects of boundary layer and airplane

angle-of-attack.

True time on the film records of figures 11-13 is represented by right to left direction, thus when

the XB-70 was overtaking the F-104 probe airplane, the XB-70 bow-shock is presented first and

its tail shock last. At the top of the film were a series of dots at 0.5 sec time intervals (shown

where available). However the 0.5 second time interval is indicated for each run. Because of the

fore and aft displacement of the two sets of orifices on the probe (about 10-inches), penetration of

the rear gage (gage 2) is indicated a very short time ahead of the indication by the front gage (gage

1) during the XB-70 overtake of the F-104, and the reverse is true when the F-104 overtakes the

XB-70. The absence of any noticeable oscillations in these pressure time histories, prior to the

entrance and following the exit of the probe airplane from the XB-70 flow-field, indicates the

probe tip vibrations, experienced on some of the close-in runs on the B-58/F-106 probe tests of

1963 (ref. 9) was not experienced on these flights. Also indicated in figures 11-13 are the estimated lengths of each signature based upon the film timing marks and the closure rate between

the XB-70 and F-104 as determined from the radar data and listed in Table II.

On the first probe flight of November 23, the F-104 was able to complete only one penetration

which was at a position 3290 feet above and 7100 feet to the left-side of the XB-70. The pressure

time histories are shown in figure 11. Examination of the pressure traces indicate the presence of

three main shocks with another weaker shock appearing just prior to the tail shock. Note, too, that

the gradual downward shift of the ambient levels on both pressure gages prior to the F-104 sliding

back through the XB-70 bow shock and continuing throughout the flow-field traverse, indicating

a slight change in F-104 altitude during the probe run.

A signature length X of 529 feet is established based upon a closure rate between the two aircraft

of 310 ft/sec (see Table II). It should be noted that this signature length is significantly longer than

those observed at ground level for the XB-70 flying at the same operating conditions of Mach 1.5

and 35,000 feet altitude. The signature length of the 529-feet results primarily from the 14 degrees

difference in headings between the XB-70 (259 deg.true) and F-104 (273 deg.true) aircraft during

the time of penetration (see Table II). Thus, the F-104 aircraft probed the XB-70 flow-field on a

“skewed” rather than a “parallel” path 3290 feet above and 7100 feet to the left of the XB-70

flight track resulting in a longer signature. In addition to the heading differences, the slight change

in the F-104 altitude from the beginning to end of the probe penetration (as indicated by the

downward shift of the ambient level signature traces in figure 11 would also contribute to the

observed increased signature length.

On the second probe flight of December 12, 1966, the F-104 was able to complete two penetrations of the XB-70 flow-field. The first pass with the XB-70 overtaking the F-104 occurred at a

distance of 4727 feet below and 220 feet to the side, and on the second pass the F-104 overtook

the XB-70 penetrating the flow-field at a distance of 4656 feet below and 825 feet to the side. The

difference in the XB-70 and the F-104 headings during each of these two penetrations was 2

degrees and zero degrees, respectively. Pressure time histories of these two passes are given in
 14

figure 12. It can be seen that these signatures consist of four primary shocks with an additional

two secondary shocks. Signature lengths of 198 feet and 220 feet are noted for the two passes

based upon closure rates of 155 ft/sec and 344 ft/sec, respectively (see Table II). The difference in

flow-field penetration time for the two passes is apparent from the traces.

On the third probe flight of December 16, 1966, the F-104 was able to complete three penetrations

of the XB-70 flow-field. The first pass with the XB-70 overtaking the F-104 occurred at a distance of 1870 feet below and 2900 feet to the side. On the second pass, the F-104 overtook the

XB-70 penetrating its flow-field at a distance of 2031 feet below and 980 feet to the side. On the

third pass, the F-104 slid back through the XB-70 flow-field at a distance of 1802 feet below and

590 feet to the side. The difference in the XB-70 and F-104 headings during each of these three

penetrations was 2 degrees, zero degrees, and zero degrees, respectively. The pressure time histories for the three passes are presented in figure 13. It can be seen that from four to five major

shocks are evident with as many as four additional secondary shocks. Signature lengths of 252

feet, 211 feet and 188 feet are noted for these three passes based upon closure rates of 268 ft/sec,

153 ft/sec and 103 ft/sec, respectively (see Table II).

Peak positive overpressures. - Following each of the probe flights, the NASA Flight Research

Center (now Dryden Flight Research Center) applied the appropriate calibration curves for each

of the two gages on each of the three flight test dates to the pressure time histories of figures 11-13

and established the results presented in figure 14. Each of the six flow-field signatures were

reconstructed from figures 11 to 13 so that they all begin with the bow-shock and end with the

tail-shock. The location of each of the individual shocks are presented as ratio of the total signature which is 100 percent. Gage 1 (front gage) is represented by the solid line and gage 2 (aft

gage) by the dashed line. It should be emphasized that the finite shock rise times associated with

the signatures of figure 14 are a result of the reconstruction process and, thus, are not a true indication of the actual shock rise time. The main purpose of figure 14 was to assign an absolute overpressure scale to the film trace signatures of figures 11, 12, and 13.

Examination of the signature traces of figure 14 shows that the overpressures measured by gages

1 and 2 are fairly consistent, with somewhat higher values recorded by gage 1 for locations below

the XB-70 and slightly higher values observed by gage 2 for the single pass above the aircraft.

Peak positive overpressures ranged from about 1.2 lbs/ft2 at 42 body lengths above and to the side

of the XB-70 (fig. 14a) to as high as 5.0 lbs/ft2 at 10 body lengths below the aircraft (fig. 14f). In

all but one case (fig. 14d), the overpressure associated with the tail wave shock was equal to, or

greater than, any of the positive peaks in the signature with a maximum of 6.0 lbs/ft2 observed on

pass 3 of the third flight (fig. 14f).

A summary of the maximum positive overpressures associated with each of the six signatures, as

measured by gages 1 and 2, are listed on the right hand side of Table VI along with the signature

length (X) and period (T). On the left hand side of the Table are listed the aircraft flight conditions and slant range distance (r) from the probe aircraft perpendicular to the XB-70 flight path.

Also shown on the last column on the right of the Table is the calculated maximum overpressure

obtained by the “Carlson” method (ref. 18) which was in use at NASA Langley during the time
 15

period of the probe tests. It can be seen from Table VI that good correlation exists between the

measured and predicted overpressure levels.

Signature lengths. - The signature lengths X described as the distance between the bow and tail

shocks of each of the six probe signatures (listed in Table VI) is based upon the film time interval

data of figures 11 to 13 and the aircraft velocities obtained from radar. These values (with the

exception of the probe flight of November 23, 1966) along with the average signature length of

three XB-70 flights obtained by the cruciform and site 9 ground measurements (see Table VII) are

plotted as a function of slant range distance from the XB-70 aircraft and presented in figure 15.

Also shown on the figure is a horizontal dashed line representing the length (189 ft.) of the XB-70

generating airplane. The solid line curve represents the calculated values of X from equation (3)

of reference 1, which is based upon the far-field volume theory of reference 20 and was in use at

the time of these flight experiments.

Both the simplified volume theory and the data points indicate an increasing wavelength with

increasing distance from the generating airplane similar to that observed in reference 21 for small

projectiles and in figure 12 of reference 22 for two fighter-type aircraft and a bomber aircraft. It is

interesting to note the results presented in figure 12 of reference 22 reflect very good correlation

with the ground measured signature lengths and those calculated from equation (3) of reference 19

for small aircraft, such as the 54-foot long F-104 and the 65-foot long F8U3. For the 97-foot long

B-58, the wavelengths shown in figure 12 of reference 22 obtained from the probe flights (fig. 17

of ref. 9) and ground measurements (ref. 1), the correlation with the simple volume theory begins

to deteriorate. This is because the theory assumed volume effects only (no lift effects) and that the

aircraft is a point source. In the present case, where the XB-70 length is 189-feet and the effects of

lift are quite pronounced as compared to the smaller and lighter weight fighter aircraft, the volume

theory curve shown in figure 15 is totally inappropriate.

Ground Measurements

Microphone set-up and characteristics. - Within about 2 to 5 minutes after the completion of the

F-104 probe flights, the XB-70 arrived overhead of the main test area (the elevation of EAFB is

about 2300 feet above sea level) maintaining essentially the same steady-level flight conditions of

about Mach 1.5 at 37,000 feet MSL (see fig. 9a). Two sonic boom ground measurement set-ups

were located along the 245 degrees magnetic ground track and consisted of the free-field microphone cruciform array and the Site 9 microphone ground array some 1800 feet up-track (see fig.

9b). The cruciform array consisted of 5 microphones located at ground level at 100-foot separating distance. An additional mast microphone was suspended at a distance of 20 feet directly above

the control ground microphone. This array was employed to provide information about the wave

shapes, wave angles , overpressures, durations (periods) and rise times. Aircraft ground speeds

were calculated, as were the wave angles in both the horizontal and vertical planes, based on measured shock arrival times. Site 9 involved a ground array of 42 microphones at spacings of 10 to

200 feet in order to assess the influence of the atmosphere on sonic boom waveforms.

Each channel of the measuring system used in ground measurements consisted of a specially

modified microphone, tuning unit, d.c. amplifier, oscillograph recorder and FM magnetic tape

recorder. The usable frequency range of the complete system, including data reduction, was 0.02
 16

Hz to 5000 Hz. Field calibrations by means of discrete frequency calibrations provided a system

accuracy of + 1 dB (see ref. 1)

Wave shapes. - Measured sonic boom signatures for the three XB-70 probe flights, as observed at

microphone 1 (MLC 1) in the cruciform array and at microphones FRC 1 and LAC 31 in the Site

9 array are shown in figure 16. It may be noted that all eight signatures consisted of three shocks

and included a bow and tail shock and an intermediate shock; the signature had not coalesced into

a classical N-wave. Bow-shock overpressures ranged from p = 2.3 lbs/ft2 to 3.14 lbs/ft2 for signatures that are slightly rounded and peaked, respectively. Wave period ranged from 224 ms to

240 ms. Comparison of all eight signatures showed some influence of the atmosphere effects

regarding peaking and rounding of the shock fronts.

Signature characteristics. - A summary of the XB-70 sonic boom signature characteristics measured at ground level with the six microphone cruciform array is give in Table VII based on the

data taken from reference 23. Signature characteristics included positive and negative overpressures, bow shock rise time to the one-half and maximum amplitudes, wave period and signature

length as described in the signature sketches in figure 17. The signature length is obtained from

the measured signature period and measured airplane ground speed from the cruciform array.

Also listed on the left hand side of the Table are the XB-70 operating conditions of Mach number,

altitude, heading, distance lateral to array and boom time. On the right hand side are the shock

wave angle, ground speed and the waveform category. The 10-waveform categories used to classify all of the sonic boom signatures acquired during the 1966-67 EAFB National Sonic Evaluation Program (ref. 17) and previous sonic boom flight experiments are also illustrated in figure 17.

In addition to the 10 waveform categories shown, word descriptions are also given to each of the

categories by means of a single, two, or three letter designation; for instance, a type NP was

judged to be intermediate between a type N normal waveform and a type P peaked waveform. An

SPR is a spiked-peaked-rounded signature.

Although actual signature traces are no longer available for five of the six cruciform microphones,

an examination of the signature characteristics listed in Table VII suggests that the signatures

would be very similar to those previously shown (see fig. 16). Note that the free-air microphone

(MLC 6) atop the 20-foot mast has a bow-shock overpressure about half that measured at ground

level. Also, the signature length X, based on a measured average ground shock speed for the three

flights of about 1375 ft/sec and an average wave period of about 232 ms is 320 feet.

Correlation With Airplane Geometry

One of the main objectives of the XB-70/F-104 was the same as for the B-58/F-106 probe tests of

reference 9; namely, to obtain definite information relative to the way in which lift effects and

volume effects of the much larger and heavier airplane combine in the generation of the shock

wave patterns from the generating airplane. The data of figure 18 have been reproduced from figures 11, 12, and 13 to illustrate some of these findings. It was found in references 5 through 9 and

reference 11 that the shock wave patterns about an airplane were closely related to the airplane

geometry. In the present study, pressure signatures measured above and below the XB-70 generating airplane have been adjusted in signature length to conform to the length of the airplane and

are compared with sketches showing the main components of the airplane.
 17

Two general observations can be made. Some correlation exists between the locations of the individual shock waves and the geometrical features of the airplane. It is also obvious that the pressure signatures measured above and to the side of the airplane varies markedly from that

measured below the airplane. In particular, the location, number and amplitude of the individual

shock waves are different, and furthermore, below the airplane the positive area of the signature

exceeds the negative area, whereas the reverse is true above the airplane. Such a result would be

expected for airplane operating conditions in which lift effects are significant.

Although the probe measurement that was made at about 3000 feet above the XB-70 was not

directly overhead but some 7000 feet to the side, the signature takes on significance relative to

defining the initial conditions for the “over-the-top” or “secondary” sonic boom (ref. 14) of which

little was known and few were aware of at the time that these tests were conducted.

SUMMARY REMARKS

A series of in-flight flow-field measurements have been made above and below the USAF XB70-1 airplane using an instrumented NASA F-104 aircraft with a specially designed nose probe.

These flight tests were accomplished during the 1966-1967 EAFB National Sonic Boom Evaluation Program and involved three steady-level flights of the XB-70 on three separate days. On all

three flights, the XB-70 was at a Mach number of about 1.5 at an altitude of about 37,000 feet

above sea level and at gross weights of about 350,000 pounds. A total of six supersonic passes

with the F-104 probe aircraft were made through the XB-70 flow field; one above the XB-70 on

the first flight, two below the XB-70 on the second flight, and three below the XB-70 on the third

flight. Separation distances ranged from about 3000 feet above and 7000 feet (42 body lengths) to

the side of the XB-70 and about 2000 feet and 5000 feet (10 to 26 body lengths) below the XB-70.

Measured sonic boom signatures from the three XB-70 flights were also acquired at ground level

very soon after the probe measurements were acquired (190 body lengths).

The in-flight pressure signatures measured above and below the XB-70 were complex in nature

and had the appearance of a “sawtooth” waveform. As many as five major shocks were observed

below and near the XB-70 and the number of shocks diminished as distance from the aircraft is

increased. The influence of combined lift and volume effects are quite evident when comparing

the signatures measured below the aircraft to the one measured above the aircraft. A maximum

positive overpressure of about 1.2 lbs/ft2 was measured above the XB-70 on the first flight (42

body lengths distance), about 2.7 lbs/ft2 to 5.0 lbs/ft2 were measured below the XB-70 on the second flight (26 body lengths), and from 3.5 lbs/ft2

 to 5.0 lbs/ft2 were measured below the XB-70

on the third flight (19 to 10 body lengths). At ground level, the sonic boom signature had not yet

coalesced into the classical far-field N-wave but contained an intermediate shock. The maximum

bow shock overpressure was about 2.6 lbs/ft2 and the period of the signature was about 230 msec.
 18

 APPENDIX A

(Taken from NASA TN-D 1968, October 1963)

DESCRIPTION AND STATIC CALIBRATION OF PRESSURE INSTRUMENTATION

by John F. Bryant, Jr.

 The instrumentation for measuring the pressure field about the bomber airplane consists of the

following components: Two NASA model 49-TP inductance pressure gages (ref. 17) and a resistance-type temperature pickup mounted in the special probe on the fighter airplane as shown in

figure 5; a carrier amplifier, an NASA recording oscillograph, a resistance-type temperature control box, and an NASA timer mounted in the rocket bay of the fighter airplane; and two solenoid

valves and two constant-temperature chambers mounted in the nose bay. The pressure gage converts the static pressure on the probe into impedance changes which produce an unbalance on the

inductance-resistance bridge. This output is amplified and demodulated in the carrier amplifier

and recorded on film in the oscillograph.

 The instrumentation necessary to measure this pressure field had to be suitable for flight environments. Also required was a high sensitivity and a frequency response that was flat from zero to

30 cps. To obtain the high sensitivity, a differential pressure gage was used. An absolute pressure

gage, normally used to measure static-pressure changes, would not produce the required high sensitivity. When using a differential gage for this type of measurement, it is necessary to equalize

the pressure on the gage during the time that the fighter airplane is climbing and descending. During the measuring period one side of the gage must be sealed off and used as a reference; this was

accomplished by connecting one side of the gage to the reference orifice through a solenoid valve.

Also connected in the reference side was a constant-temperature chamber. This added volume

minimized changes in the reference pressure due to temperature changes caused by the aerodynamic heating of the long lengths of tubing that connected the reference orifice on the instrumented probe with the valve in the nose section. The volume of the tubing was about 1 percent of

the chamber volume. To obtain the required frequency response, it was necessary to minimize the

time lags by locating the measuring pressure gage very close to the orifice. The NASA type 49

gage was selected because of its high sensitivity, good accelerating characteristics, and very small

size. Since its dimensions are only 1/4 by 7/16 inch, the gage could be mounted directly in the

probe close to the orifice. All the other instrumentation was standard flight equipment.

 It was decided to use two gages: gage 1, which measured the static pressure on the needle nose

of the instrumented probe, and gage 2, which measured the static pressure on the body of the

probe. (See figs. 4 and 5.) Gage 1 had a sensitivity of approximately 10 lb/sq ft per inch of film

deflection and was recorded by a 100-cycle galvanometer. Gage 2 ad a sensitivity of approximately 20 lb/sq ft per inch of film deflection and was recorded by a 50-cycle galvanometer. Once

the reference valves are closed, the gages essentially become very sensitive altimeters. Gage 2

was used as a backup in case gage 1 was driven off scale by too large a change in altitude of the

fighter airplane after the pilot had closed the reference valve. The lower frequency galvanometer

was used to filter out any high-frequency noise that might occur.
 19

 The response of each measuring system was determined by the frequency response of the

recording galvanometer. An example of this is shown in figure 19, where a step function was

applied to the 50-cycle galvanometer and a step function was applied to the entire measuring system. It can be seen from these step functions that the response of both is the same. The time lag of

the reference system was 3 seconds. This large lag limited the rate of climb and descent of the

fighter airplane to 6,000 feet per minute and thus kept the gages and amplifiers from being overloaded.

 The accuracy of the overall system was estimated to be 3 percent of the peak positive overpressures listed in table III. The hysteresis of the gage was 1 percent, and the accuracy of the galvanometers and amplifiers was 2 percent. The change in sensitivity of the gage was 6.5 percent per

100o F change in temperature. This was correctable to 1 percent by use of the resistance temperature gage. The effect of accelerating forces along the longitudinal axis of the fighter airplane (normal to the diaphragm) was 0.05 lb/sq ft per g. The system was constantly monitored by making

static calibrations before and after each flight.
 20

APPENDIX B

DESIGN AND AERODYNAMIC CALIBRATION OF PRESSURE PROBE

By Virgil S. Ritchie

DESIGN

Basic Considerations

 The design of a flight probe for sensing static-pressure changes in the pressure field of a large

disturbance-generating supersonic airplane involved in a number of aerodynamic and structural

considerations. A probe of conical shape and relatively large dimensions was considered suitable

for a cantilever-type installation at the end of the nose boom of a probe airplane. The conical

shape afforded the advantageous features of weak tip disturbance and thin boundary layer. The

large dimensions afforded structural rigidity, suitable locations for miniature electrical pressure

gages near the pressure-sensing orifices, and relatively large Reynolds numbers. The location of

pressure gages near the sensing orifices reduced the possibility of pressure-lag errors. The large

Reynolds numbers increased the likelihood of realizing a turbulent boundary layer on the probe

without the use of artificial transition-fixing devices, which could introduce shock waves ahead of

the pressure-sensing orifices. An arrangement of two small orifices circumferentially located in

null-pressure regions about 75o apart afforded some reduction of the errors associated with

changes of flow angularity (crossflow) around the conical probe. This asymmetric arrangement

necessitated probe orientations with the pressure orifices facing the incident disturbance wave to

be measured, but it was considered superior to a symmetrical arrangement of orifices distributed

around the circumference of the probe. The asymmetric arrangement was employed for a primary

system of pressure orifices located in the conical tip portion of the probe and for a secondary system of orifices located in an enlarged conical region of the probe. For the latter system of orifices,

which was employed to supplement the primary systems, suitable calibration information was

required, because of likely effects of the probe-enlargement shock wave as well as the thicker

boundary layer at the secondary location.

Present Application

 Principal details of the flight probe and its installation on the nose boom of a “century series”

supersonic airplane are shown in figures 4 and 5. This probe employed six pressure-sensing systems including the two systems for indicating disturbance-related pressure changes, two systems

for providing reference pressures for the differential-pressure gages, and systems for providing

approximate free-stream static (ambient) pressure and pitot pressure for the airplane flight instruments. The orifices and the tube for providing approximate ambient and pitot pressures for the

flight instruments were located at the bottom of the probe for all flights. The forward end of the

probe was made rotatable in order to facilitate the required orientation with disturbance-sensing

orifices facing the incident disturbance waves from the generating airplane. The rear portion of

the probe was secured to the nose boom in such a manner that the angle of attack of the probe

would be near 0o for the expected flight conditions. The miniature pressure gages in the probe
 21

were installed with their diaphragms perpendicular to the longitudinal axis of the probe in order to

minimize possible effects of lateral acceleration.

 WIND-TUNNEL TESTS

Introduction

 Early evidence concerning the reflection characteristics of the probe was obtained from unreported preliminary tests of a 0.75-scale model of the flight probe in the Langley 4- by 4-foot

supersonic pressure tunnel at a Mach number of about 1.82. The average test Reynolds number

(per foot) was about 2.6 x 106

, and the average static pressure corresponded to a pressure altitude

of about 50,000 feet for standard atmospheric conditions. These tests involved the streamwise

movement of the probe (with natural transition) across a disturbance (bow wave) generated by a

body of revolution and the measurement of probe-sensed pressure changes in the vicinity of the

disturbance. These early tests indicated that the primary system of orifices of the probe sensed the

same maximum pressure changes (across the employed shock wave) that were estimated by theoretical methods, whereas the secondary system of orifices sensed pressure changes about 10 percent larger than the estimated values. Also, the probe-sensed pressure changes in the vicinity of

the disturbance appeared to be of the type generally associated with turbulent boundary layers

(ref. 18). On the basis of this early information, the full-scale flight probe was constructed and the

in-flight measurements were undertaken with the view of investigating the reflection characteristics of the flight probe by means of wind-tunnel tests at a later date.

 Accordingly, after in-flight measurements, tests of the flight probe were conducted in the Langley 4- by 4-foot supersonic pressure tunnel to calibrate the approximate reflection characteristics

of the probe at a Mach number near those employed for the in-flight measurements. The probe

reflection characteristics were largely determined by the same procedure as that employed for the

early tests at a Mach number of 1.82. This procedure involved streamwise movement of the probe

across a weak axisymmetrical shock wave of predetermined strength and the measurement of

probe-indicated pressure changes across the disturbance.

 Unreported additional tests of the full-scale probe across weak shock waves in the Langley 4-

by 4-foot supersonic pressure tunnel provided information concerning the effects of angle of

attack on probe reflection characteristics. Although tests have not been included in the present

report, the results were used in arriving at the approximate reflection factors reported subsequently in this appendix.

Symbols

M free-stream Mach number

p1static pressure sensed by primary system of orifices (location 1),

 lb/sq ft

p2static pressure sensed by secondary system orifices (location 2),

 lb/sq ft.

p3static pressure sensed by system of orifices (location 3) providing 

 static pressure for probe-airplane flight instruments, lb/sq ft
 22

p1,ref static pressure sensed by orifices providing reference pressure for

 gage 1, lb/sq ft

p2,ref static pressure sensed by orifices providing reference pressure for

 gage 2, lb/sq ft

pttotal pressure, lb/sq ft

pt’ pitot pressure, lb/sq ft

p free-stream static pressure, lb/sq ft

p peak or maximum pressure change across oblique shock, lb/sq ft

r radius of body of revolution, in.

x axial distance from nose of body revolution, in.

xsapproximate longitudinal (streamwise) distance from mean location of

 oblique shock (bow wave), positive when orifices are rearward of

 shock, in.

y approximate separation distance (perpendicular to airflow direction)

 between disturbance-generating body and pressure-sensing probe or

 instrument, in.

 angle of attack of probe, deg

Apparatus and Tests

 Test facility and conditions.- The present calibration tests were conducted in the Langley 4- by

4-foot supersonic pressure tunnel at a Mach number of about 2.01 (slightly larger than the average

probe-airplane Mach number of about 1.95 employed for in-flight measurements). The average

Reynolds number per foot for these tests was about 2.4 x 106, whereas the Reynolds numbers per

foot for in-flight measurements ranged from about 1.8 x 106 to 4.5 x 106. The free-stream static

pressure employed for the tests corresponded to a pressure altitude about 55,000 feet for standard

atmospheric conditions.

 Test apparatus and procedures.- The arrangement illustrated at the top of figure 20 was used in

the calibration of the flight probe at various angles of attack. This arrangement, involving the

location of all static-pressure orifices and the pitot-pressure tube on the bottom of the probe, corresponded to that employed for the probe-airplane flights over the generating airplane. Conical tip

1 (see fig. 5) was used on the probe for the calibrations tests.
 23

 The apparatus and arrangements for generating an oblique shock wave and for surveys to

determine the strength of this shock are illustrated in figure 21. The procedure employed for surveys in the vicinity of the shock was to move the survey instrument in the streamwise direction

and measure the pressures at sufficiently close intervals to define the maximum change of pressure across the shock. Two different methods, one involving a pitot-tube technique and the other a

static-pressure orifice on a plate, gave identical results in defining the maximum pressure

changes. This oblique shock wave of predetermined strength afforded a means for determining

the reflection characteristics of the probe.

 Measurements.- Absolute manometers were used for measuring tunnel total pressures as well

as reference static pressures and pitot pressures in the test section. Differential-pressure gages

with ranges of 0.25 and 0.5 pound per square foot were employed for measuring differences

between the reference static pressure and the various local static pressures sensed by the probe or

the survey instrument. A gage with a range of 1 pound per square foot was used for measuring differences between the reference pitot pressure and local pitot pressures sensed by the survey

instrument. Gages with ranges of 3 and 9 pounds per square foot were used for measuring differences between the total pressure in the tunnel and the pitot pressure sensed by the flight probe. All

gages were calibrated before and after the wind-tunnel tests.

Data and Precision

 Probe calibration.- Most of the calibration data shown in figure 20 represent averages of measurements from two separate tests. The static-pressure data are expressed in the form of ratios of

local probe-sensed static pressures to local free-stream static pressures in order to minimize possible errors associated with flow nonuniformities. Random errors in measurements during probecalibration and tunnel-calibration tests are believed to influence the static-pressure ratios, as well

as the ratios of pitot to total pressure, by no more than +0.005.

 Pressure measurements in vicinity of oblique shock wave.- Probe-indicated static pressures in

the vicinity of the body-generated oblique shock (bow wave) are expressed as ratios of probeindicated static pressure to an average (not local) free-stream static pressure. Although these

ratios are influenced by random errors in measurements in the same manner as the probe-calibration data, the possible errors in measuring pressure changes across the oblique shock wave are

considerably less than +0.005. The survey technique appears to reduce random errors in measurement to less than about 0.15 percent of the free-stream static pressure or to less than about 3.5 percent of average pressure changes across the shock wave. An experimental measurementrepeatability check, involving several traverses of the probe across the oblique shock wave, indicated scatter of less than +2 percent in the shock-wave pressure changes sensed by the primary

orifices or by the secondary orifices.

Results and Discussion

 Probe calibration at angles of attack.- Calibration tests of the probe at various angles of attack

yielded the results shown in figure 20. The primary system of orifices and the reference-pressure

orifices in the conical tip portion of the probe indicated pressures which were generally about 1

percent larger than the free-stream static pressure. These cone-surface pressures were sufficiently
 24

influenced by angle-of-attack changes to make the primary pressure-sensing arrangement fairly

sensitive to small changes in crossflow such as might be introduced by turbulence, probe oscillations, and flow-angularity changes across shock waves, that might occur in flight. The sensitivity

of alternate conical tip 2 to angle-of-attack effects was not determined from calibration tests, but

the slightly different circumferential spacings of orifices in tips 1 and 2 (fig. 5) suggest that angleof-attack effects might be somewhat larger for tip 2 than for tip 1.

 The secondary system of orifices and the reference-pressure orifices located in the conical portion of the probe behind the enlargement region indicated pressures 2 or 3 percent less than freestream static pressure. These pressures were not influenced as much by angle-of-attack changes as

were the pressures sensed by the two systems of orifices in the conical tip of the probe.

 The orifice system for the flight instruments indicated pressures about 1 or 2 percent less than

free-stream static pressure. These pressures were influenced more by angle-of-attack changes

than were the pressures indicated by the other orifice systems. This increased influence of angle

of attack was largely associated with the size and location of the orifices for the flight-instrument

system.

 The pitot pressures sensed by the tube that was offset from the bottom of the probe were somewhat larger than those expected for a tube located ahead of the interference field of the probe. The

probe-indicated pitot pressures varied consistently with angle-of-attack changes. 

 Probe capability for sensing pressure changes across an oblique shock wave.- Figure 22(a)

illustrates the approximate capability of the probe, at an angle of attack of 0o, for sensing pressures in the vicinity of a weak shock wave. It is seen that the primary system of orifices in the conical tip senses such pressure changes with small error, whereas the secondary system of orifices

senses pressure changes considerably larger than the estimated changes. These indicated probe

capabilities are supplemented by the data in figure 22(b), which compares probe-indicated, survey-indicated, and estimated maximum pressure changes across the oblique shock wave.

 Correlation of these indicated characteristics of the flight probe at an angle of attack of 0o and

a Mach number of 2.01 with unreported characteristics of a 0.75-scale model of the flight probe at

an angle of attack of 0o and a Mach number of 1.82 indicated that the primary system of orifices is

capable of accurately sensing maximum or peak pressure changes across weak shock waves at

these Mach numbers. This correlation also indicated that the secondary system of orifices sensed

pressure rises that were too large by about 10 percent at a Mach number of 1.82 and about 30 percent at a Mach number of 2.01.

 Unreported tests of the flight probe in the vicinity of an oblique shock wave at a Mach number

of 2.01 indicated that reflection characteristics of the probe at angles of attack of 1o and -1o were

somewhat different from those at an angle of attack of 0o. Such differences were larger for the

secondary system or orifices than for the primary system.

 The described probe capabilities, as obtained from wind-tunnel tests, are believed to be representative of in-flight probe capabilities at comparable Mach numbers, Reynolds numbers, and
 25

angles of attack. Possible differences in turbulence and boundary-layer transition are believed to

be the principal sources of any discrepancies between probe characteristics in the wind tunnel and

in flight.

 Probe reflection factors for correcting in-flight measurements.- On the basis of the available

information, a reflection factor of 1.00 appeared to be appropriate for the primary system of orifices at Mach numbers near 1.82 and 2.01 and angles of attack near 0o

. The reported probe-airplane Mach numbers employed for in-flight measurements were between 1.85 and 1.99. The

estimated probe angles of attack for in-flight measurements ranged from -0.4o to -1.5o (not

including likely changes as the probe airplane traversed the disturbance field of the generating airplane). These negative angles of attack could possibly change the reflection factor by several percent. Angle-of-attack corrections have not been applied to the in-flight pressure measurements

obtained from the primary system of orifices.

 Reflection factors for the secondary system of orifices appeared to vary with Mach number,

probe angle of attack, and strength of the incident disturbance wave. Applicable reflection factors

for in-flight measurements obtained from the secondary system of orifices could not be accurately

determined from the available information, but the following values are believed to be reliable

within about 10 percent:

 The reported values of in-flight pressure data were obtained by dividing the actual measurements by these reflection factors.

 General comments.- The supersonic wind-tunnel tests of the probe designed for in-flight measurements yielded the following indications of probe capability for sensing pressure changes

across weak disturbances:

 (1) The primary system of orifices located in the conical tip portion of the probe appeared to be

capable of accurately sensing the maximum or peak changes of static pressure across weak shock

waves at Mach numbers near 1.82 and 2.01 when the probe axis was alined with the direction of

flight or relative free-stream airflow  =0o). The reflection characteristic of the probe were influenced somewhat by small changes in angle of attack.

Flight Approximate reflection factor

for secondary system

1

2

3

4

5

6

7

1.23

1.16

1.15

1.07

1.12

1.17

1.13
 26

 (2) The secondary system of orifices located in an enlarged conical portion of the probe

indicated shock-proximity pressure changes somewhat larger than those obtained by special surveys and by theoretical estimates. Approximate reflection factors for the conditions of the inflight measurements varied from about 1.07 to about 1.23.

 REFERENCES

1. Hubbard, H. H.; Maglieri, D. J.; Huckel, V.; and Hilton, D. A.: Ground Measurements of

Sonic Boom Pressures for the Altitude Range of 10,000 to 75,000 Feet. NASA TR R-198,

1964.

2. Maglieri, Domenic J.; Huckel, Vera; and Henderson R.: Sonic Boom Measurements for SR71 Aircraft Operating at Mach Numbers to 3.0 and Altitudes to 24,384 meters. NASA TN D6823, 1972.

3. Lee, R. A.; and Downing, J. M.: Sonic Booms Produced by United States Air Force and

United State Navy Aircraft: Measured Data, AL-TR-1991-0099, Wright Patterson AFB,

Ohio, 1990.

4. Maglieri, Domenic J.: Sonic Boom Flight Research, Some Effects of Airplane Operations

and the Atmosphere on Sonic Boom Signatures. NASA SP-147, 1967, pp. 25-48.

5. Maglieri, Domenic J.; Huckel, Vera; and Parrott, Tony L.: Ground Measurements of ShockWave Pressure for Fighter Airplanes Flying at Very Low Altitudes and Comments on Associated Response Phenomena. NASA TN D-3443, 1966. (Supersedes NASA TM X-611).

6. Nixon, C. W.; Hill, H. K.; Sommer, N. C.; and Guild, Elizabeth: Sonic Booms Resulting

From Extremely Low-Altitude Supersonic Flight: Measurements and Observations on

Houses, Livestock and People. AMRL-TR-68-52, U.S. Air Force, 1968.

7. Mullins, Marshall E.: A Flight Test Investigation of the Sonic Boom. AFFTC TN 56-20, Air

Res. and Devel, Command, U.S. Air Force, May 1956.

8. Smith, Harriet J.: Experimental and Calculated Flow Fields Produced by Airplanes Flying at

Supersonic Speeds. NASA TN D-621, 1960.

9. Maglieri, Domenic J.; Ritchie, Virgil S.; and Bryant, John F., Jr.: In-Flight Shock Wave Pressure Measurements Above and Below a Bomber Airplane at Mach Numbers from 1.42 to

1.69. NASA TN D-1968, 1963.

10. Maglieri, Domenic J.; Huckel, V.; Henderson, H. R.; and Putnam, T.: Preliminary Results of

XB-70 Sonic Boom Field Tests During National Sonic Boom Evaluation Program. NSBEO

1-67, pp C-II to C-II-17, July 28, 1967.
 27

11. Haering, Edward A.; Jr.; Ehernberger, L. J.; and Whitmore, Stephen A.: Preliminary Airborne

Measurement for the SR-71 Propagation Experiment. NASA TM-104307, 1995.

12. Tinetti, Ana F.; Maglieri, Domenic J.; Driver, Cornelius; and Bobbitt, Percy J.: Equivalent

Longitudinal Area Distributions of the B-58 and XB-70-1 Airplanes for Use in Wave Drag

and Sonic Boom Calculations. NASA/CR-2011-217078, 2011.

13. Garrick, I. E.; and Maglieri, D. J.: A Summary of Results on Sonic Boom Pressure Signature

Variations Associated with Atmospheric Conditions, NASA TN-D 4588, 1968.

14. Maglieri, Domenic J.; Carlson, Harry W.; and Hubbard, Harvey H.: Status of Knowledge of

Sonic Booms. Noise Control Eng. Vol. 15, No. 2, Sept.-Oct. 1980, pp. 57-64.

15. Powers, John O.; Sands, J. M.; and Maglieri, Domenic J.: Survey of U.S. Sonic Boom

Overflight Experimentation. AGARD Conference Proceedings No. 42, “Aircraft Engine

Noise and Sonic Boom,” May 1969.

16. Arnaiz, Henry H.; Peterson, John B., Jr.; Daugherty, James C.: Wind-Tunnel Flight

Correlation Study of Aerodynamic Characteristics of a Large Flexible Supersonic Cruise

(XB-70-1): III - A Comparison Between Characteristics Predicted From Wind-Tunnel

Measurements and Those Measured in Flight. NASA TP-1516, 1980.

17. National Sonic Boom Evaluation Office: Sonic Boom Experiments at Edwards Air Force

Base, NSBEO-1-67 (Contract AF 49 (638)-1058), Stanford Research Inst., July 28, 1967.

18. Middleton, Wilbur, D.; and Carlson, Harry W.: A Numerical Method for Calculating NearField Sonic-Boom Pressure Signatures. NASA TN-D 3082, 1965.

19. Maglieri, Domenic J.; Hubbard, Harvey H.; and Lansing, Donald L.: Ground Measurements

of the Shock-Wave Noise From Airplanes in Level Flight at Mach Numbers to 1.4 and at

Altitudes to 45,000 Feet. NASA TN D-48, 1959.

20. Whitham, G. B.: The Behavior of Supersonic Flow Past a Body of Revolution, Far From the

Axis. Proc. Roy. Soc. (London), ser. A, vol. 201, no. 1064, Mar. 7, 1950, pp. 89-109.

21. DuMond, Jesse W.M.; Cohen, E. Richard; Panofsky, W.K.H.; and Deeds, Edward: A Determination of the Wave Forms and Laws of Propagation and Dissipation of Ballistic Shock

Waves. J.Acous. Soc. of America, vol. 18, no. 1, July 1946, pp. 97-118.

22. Maglieri, Domenic J.; Parrott, Tony L.; Hilton, David A.; and Copeland, William L.: LateralSpread Sonic-Boom Ground-Pressure Measurements From Airplanes at Altitudes to 75,000

Feet at Mach Numbers to 2.0. NASA TN D-2021, 1963.

23. Hubbard, H. H.; and Maglieri, D. J.: Sonic Boom Signature Data from Cruciform Microphone Array Experiments During the 1966-67 EAFB National Sonic Boom Evaluation Program. NASA CR-182027, 1990.
 28

24. Patterson, John L.: A Miniature Electrical Pressure Gage Utilizing a Stretched Flat Diaphragm. NACA TN 2659, 1952.

25. Liepmann, H. W., Roshko, A., and Dhawan, S.: On Reflection of Shock Waves From Boundary Layers. NASA Rep. 1100, 1952. (Supersedes NACA TN 2334.)

26. Whitham, G. B.: The Flow Pattern of a Supersonic Projectile. Communications on Pure and

Appl. Math., vol. V, no. 3, Aug. 1952, pp. 301-348.

27. Carlson, Harry W.: An Investigation of Some Aspects of the Sonic Boom By Means of

Wind-Tunnel Measurements of Pressures About Several Bodies at a Mach Number of 2.01.

NASA TN D-161, 1959.
 29

Table I.- Geometric ebaractcrinics of XB-7Q.l airplane. 

(from reference 16) 

TCUI wing 

TOlal area (includes 230.62 m! (2482.34 rtl ) covered by fuselage bulllOt 

3.12 m2 (33 . .53 r(2) of tile wing ramp area), m2 (f~). . ........ . 

Span, m (fl) . . .. .. . . . . . . . . . . . . . . . . .. . . • .. . ..... . 

Aspect ratio. 

Taper ratio. . . •..•..•. 

Dihedral angle. deg . . ................ . 

Root cbord (wing slation 0). m (1"1) .••.••.••.........•••. 

Tip chorl1 (wing station 16m (630 In.)). m (ft) ............... . 

Mean aerodynamic chord (wing station 5.43 m (17.82 (ft». m (fl). 

. ..... 58.5.07(6297.8) 

. .... 32 (103) 

. ............. 1l.7!i I 

............. . 0.019 

. ........ 0 

. ....••.•. 35.89(1 17.76) 

. ...... 0.67 (219) 

. .... 23.94 (78.532) 

Fuselage 'tation of2!i-percenl wiDg mean aerodYlWllic chord. m Crt) ................. 41.18 (13!i. 10) 

Sweep bad: angle. deg: 

Leading edge .. ......................... ........ . 

2!i-percem e1emenl . 

Trailing edge ... 

Incidence angle. deg; 

Root (fuselagej\lllCtUre) ..................... . 

Tip (foklline 100 oulboard . 

Airfoil section (modified bc:iI;agonal): 

. .. 6."i . .'j7 

. .. .'j8.79 

. ............. 0 

..0 

. ....... ·2.60 

ROOIIO wing mlllioo 4.72m (186 In.) (thickness..ctud ratio, 2 percent) .....•..•....... 0.30 to 0.70 

Wing station 11.68 m (460 In.) to 16.00 m (630 in.) 

(thicWss-clKnl ratio, 2.5 percenr.) .......................••. . .0.30 10 0.70 

Inboard wing· 

Area (includes 230.62 m2 (2482.34 ft2)coveJed by fuselage bul not 

3.12 ml (33 . .'j3 ft2) wing ram area, m! (fi2) ............................... 488.28 (.'j2.56.0) 

Span, m (ft) .. . .. . .. ,. ...... ... .... .... . .. .... 19.34 (63.44) 

Aspect ratio . . ..... ,....... . .............. . .. . ................. ....... 0.766 

Taper ralio ....... ................... .. .. . .. 0.407 

Dihedral angle. deg. . ....... ' ....... 0 

Root ctod (wing sl3lioo 0)' m (ft) ...... . 

TipcllOn:l (wing station 9.67 m 080.62 in.)). m (ft) ....•..•. 

Mean aerodynamic cbon:! (wing station 4.lj m (l63 . .'j8 in.», m (in.) 

Pusclagc station of 2,.'j-pcroenI wing mean acrOOynamic cboo:!, m (in) ... 

Sweepbaelr: angle. deg: 

, ......... 35.89(117,76) 

14.61 (47,9<\) 

..... 26.15(1053) 

..39.07(I!i38.29) 

Leading edge. 

2!i-pereent element . 

. ...................... ......... .......... 65.57 

Trailing edge. 

Airfoil section (modifkd hexagonal): 

Root (tlticlr:ness-cbord ralio. 2 percent) ............... • • . 

Tip (tltickncss-ellord ratio. 2.4 percent) ...............•. 

. ........... ... .'j8.79 

. ............. 0 

. ..... ... 0.30100.70 

. .... 0.30100.70 
 30

Table 1.- Continued. 

Mean camber (leading edge), deg: 

Butt plane 0 ...... . ......... . . 

Butt plane 2.72 m (107 in.) ....... . 

Butt plane 3.89 m (153 in.) ...... . 

Butt plane 6.53 m (257 In.) .. . 

Buu plane 9.32 m (367 in.) 10 lip .. 

O • .IIboal'd wing· 

Area (one side only), m2 (n2) 

Sp.an. m (ft) ............ .. ... . 

Aspect ratio 

Taper ralio 

. . . . . . . . . . . . . . . . . . . .. 0.15 

...... . ......... .4.40 

, .............. , .. . ..... , ..... 2.75 

, .2.60 

.. .. 0 

Dihedral angle, deg ....................... , ...................... . 

.. 48.39 (520.90) 

6.33 (20.78) 

0.829 

0.046 

.. , 

Root chord (wing station 9.67 m) (380.62 in.»), m (fI) 

'lip chord (wing Slalion 16.00 m) (630 in.»), m eft) , •. "."., 

14.61 (47.94) 

.......... , 0,67(2,19) 

Mean aerodynamic chord (wing 5I&tion 11.87 m) (467.37 in.», m (in.) ,. 9.76 (384,25) 

S w«pback: angle, deg: 

Leadingedge. "., .. ,.",., ... , .. ,.,." .. " .. "". ,.,." ..... Mol7 

25-pcrcenl element ..... .., ., .. , . ......•... , .. .. ,., ... ,. ,18.79 

Tniilingedge .......... ... . . .. ........ . ... , .. ,., 0 

Airfoil5CCtion (modified helUlgonaJ): 

Root (lhickness-chofd ratio, 2.4 peteem) ..... , .. , , ...... ' .... , " ..... ....... 0.30 10 0.70 

Tip (thicl:ness-chord ralio, 2.5 percent) •.•.•...•.. , •.•. , • ' ... , . , •. ... 0.30100.70 

Down deflection from wing reference plane, deg 

Skewline of tip fold, deg: 

Leading edge in ..... , ... .. . 

Leading edge down. 

ElevollS (data for one side): 

Total area an of hinge line. m' (nl) 

Sp.an. In (ft) 

lohoard chord (equivalent), m (in.) 

Sweepback: angle of hinge lioe, deg 

Deflection, deg: 

As eleV81()f ........ . . 

As aileron with ch:valOrs at ±ISo or less 

As aileron with elevators at .2.10 or less 

TOI:a1 ... 

C"""' . 

, ..... , . ,.,. 0.2.1,65 

, . 1.5 

". 3 

Winl tjps 

18.37 (l97.7) .... . 12.57 (135.26) 

6.23 (20.44). . .. . . . . 4.26 (13.98) 

29.1(116) " .............. (l16)2.~ 

o ............................. 0 

.. ·25 to 15 

. -15toI5 

.... • .1 10,1 

, ..... ·301030 

Area (iocludes 13.96 m2 (150.31 fl2) COVCLa:\ by fuselage), m2 (fi2). , ..... , . , ..... 38.61 (4 1.1 . .19) 

Span. m (ft) ....... ,., .. ,. , ... ,., ....... .. .. ,. ,.,., .. , ................... 8.78 (28.81) 

Aspect ratio ...•..•.•. .• ... • . •.. •...•...•...•. ...•. ••..•...•...... , •.....•...•. 1.997 
 31

--

~-.. a-.-I_· ..... GI,·lftl 

t; """ 1 __ '3111_11'1. "1~.(I:) 

_ .. uI) ' ___ lt1.r~I'Iill.II..1oJ 

-... .... ,,~-- -"-_.(\&), 

I 

_ 

,'I ....... ~ 

.. 

--. _ .. 

h- t ....,r-.,I .• , __ 1_1Ioi .... ,.-11 _I..... t_'_l~,...-I 

• , __ Ul-, a.do,,_ ..... ~ ... c-n.,_al ..... _1 ..... ..,.boI,,;l.8'l. 

V..-OoiI_"IWOj-

... "'_ ..... _- ...... (lIoch;toI 0 IJ m' II !II> ftIJ .,.' 

m'.I'1 _0" 

_ .... , 

T __ 

..... _1_ .. 1_ .... '11) 

'~ 

T't_I ___ '1'l. 

UIO.'L.Ul) --,_' __ ~_I.". 

(l3" Ol}).. (Ill) 

.......... _ofJ,j_.-...I_ "'''''' ! -"nO~) 

$w ............... 

~oI .. 

---

_ .. --1- _~'_ ...... 

I '_1~J.-l · ..... , "'_"'U~l 

~-.. .... ~ ........ ooI .. 'Oi .... --.... ",-

'" " 

, . 

,II)" 

• t.JoO(»". 

U'!'lI6, 

."'016 1) 

1.III(JlJ"1 

". 

.,". H I 

••• .,..~ .. 

0).4.0 .. 

.~ 

~OI(}<~ 

.~ 

11.'7< (Willi) 

.,J 1m 

~ 

'(I)(1JOO 

!II ,& 9.1) 

11I1 U'H /t 

"j11 (1 11UO'> 

"" " " • ". -

..... 

OlO .. O'lt 

• 'ID' 
 32

Wlt.b gear (XteOOcd ... 

Witb gearrdnlCled . 

Fuselage (includes canopy) -

l...cngth. m (ft) 

T.ble 1.. Canti[IUal. 

Maxillll.lm deptll (fuselage station 22.30 m 

(878 in.). M (In.) ..•.. . ..... .. ..• 

Maximum breadt.b (fusell ge sUltion 21.72 m 

(8j!i in.» . m (ill.) ..• 

Side area. m2 (ttl) . 

PlanfQITTI area. m2 co1) .. Ceutcr of gravity' 

"'" . 

Forwarrl limit. percell! mean Mefodynamic cbonl • 

Aft hmit. percent m<:UIl aerodynamic choro 

Lcllgm. m (fl) 

Muimum depth f~sc1age station 34.93 m 

(l37j in.». 1U (in.) .................. . 

MaximliID bl"eadt.b (fuselage . ta!ioo 53.34 III 

(2100 in.». m (in.) ............... . 

Side an:a. m! (I'll) 

l'lanform area, m2 (ftl) . . . . ... . . . . ••. . . .. ... . 

Inlet capi\"e ami (eacb). m1 ( 112) 1.6 1 ~ . .... . . 

Surface arM$ (net .... eued). m! (ft ~-

FliselA!c. e800py. bound.ary layer gUllel, and. t.alIpJ ~ .. 

D.JCt$ .. ...... . .......... .. 

Wing. ... illS li ps. 1IOd wtng ramp 

Ven:ieal la..i\.:; (t .... o) 

ClJlard 

Total .•. . . ..•.. ... .......•. .. 

En¥iocs (six) 

DoalWl angl e.. deS -

Upper ~rfllCtl 

Lowec surface ••... 

Sid..': . . . •. . .. . . . . 

Base areas. m! (fr2)-

Total (all engines on. minimwn Cxil area) 

Total (all CDgincIi 00. rnlAimum exit area) 

PJoj<:Ctoo Ihlckness (bcigbl) of base. m (in.) 

Width of propulsiQII padagc. em (in.) 

Eogi!ll:-

JC -~ l aTe8 (minimum). em2 (j02) . .... ... . • . .... . . . . .. . .. . 

.., . .... ±3 

S6.61 (l8!i7!i) 

. .... 2.72(106.92) 

. . 2.!i4 (100) 

. .. 87.30(939.72) 

11 0.07 (l 11W.78) 

. ...... .. .......... 19.0 

....... ... . . ......... 25.0 

11.96 (1 ~84) 

.. 2.31 (90.75) 

. ... 9.16 (360.70) 

66.58 (116.66) 

2 17.61 (2342.33) 

. ....... 3.6 1 (5600) 

. . . 264.77 (2850.0) 

318.71 (l<&30.6) 

. .. .. 864.71 (9307.7) 

87.1 2(937.7) 

49.4i (532.5) 

58 9{l7 0~ 5) 

.. YJ9J.GE-3 

........... 6 

. , 

. 6 

.12.7 (137) 

10 (107.2) 

... ... . U (48.5) 

.. 147 (.5 8) 

. 9 14 ()(,o) 

4613 (71 j ) 
 33

Table l.- Concluded. 

Jet-exit area (maximum). em2 (in2

) .......................•................. 13.678 (2120) 

Jet-exit diameter (minimum). em (in.) ....... ............................. ....... 77 (30.2) 

Jet-exit diameter (maximum). em (in.) .......... .•.. .•...................... ... . . 132 (~2) 
 34 Table II.- Summary of XB-70 and F-I04 in-flight probe test conditions. Flight Condition XU· 70 Generating Aircraft F-] 04 Probe A ircrllft Position of F·I04 relative to XD· 70 (ft) Probe Shock AM ,IV DlIl.e MSN ru, Probing Probing Altitude Gnl. Tnlc G"", C._ left right Altitude GnJ. True S Y Z , lime, (Z) M Velocity bdg. weight po'. M velocity V lO4 hdS_ (slant '.' Region Region (ft.mst) (IV",,) (dc,) (Ib) (deg) hand band (ft. ""I) (fV",,) (dc,) (fV=) Cbohiud) (to side) below) rRnge) 1I!23/66 1·1 I F-I04 )ffi·70 18:27:45 36,550 1.45 1300 259 324,000 1.05 11.6 11.4 39.840 LIS 1050 ·3 273 .30 310 10,750 7,100 ·3.290 7825 42 above overtake.', XD-70 F·I04 )ffi·70 overtAkes 18:27:32 37,639 1.50 1371 261 344,000 1.08 Il.S 11.3 32,912 1.32 1216 ·2 263 .18 ISS 5525 220 4727 4732 26 I F-l04 F·104 12/12/66 3-2 r-- below XB·70 F·I04 2 overtakes 18:29:18 37,584 1.49 1303 261 340,000 1.07 11.7 10.4 32.928 1.83 1707 ..0 261 34 344 5275 825 4656 4729 26 X8-70 )ffi·70 1 overtakes (5:52:06 37,870 1.47 1372 261 350,000 1.15 11.4 10.9 36,000 1.18 1104 ·4.5 263 ,29 268 2850 2900 1870 345 1 19 F-I04 

- F- I04 

12/16/66 4·2 below F-I04 

2 XIl·70 Q,'ertakcs 1$:54:06 37,900 1.46 1370 262 348,500 1.14 11.8 11.3 35,869 1.61 1523 -2.5 262 .15 153 2590 980 2031 2257 12 

XD-70 

-

XB-70 

3 overtakes 15:54:04 37,917 1.47 1300 200 344,000 l.lO 12.3 11.4 36,115 1.34 1257 ·2.04 260 .13 103 2050 590 1802 1896 10 ,F-I04 

. 
 35 Table III.- XB-70 flight conditions overhead of ground cruciform microphone array . . Average elevon Grd. Lateral Gross Canard position True Date MSN Time Altitude M velocity Offset, weight, W (deg) heading (Z) (ft.msl) pos. (ft/sec) (n.mi.) (lb) (deg) left right (deg) hand hand 11/23/66 1-1 18:31:00 37,200' 1.46 1342 1.69S 315,000 0.92 13.0 11.9 259 12/12/66 3-2 18:31:00 37,600 1.50 1384 0.17S 339,000 1.00 11.7 10.4 262 12/16/66 4-2 15:57:00 38,600 1.46 1399 0 339,000 1.11 11.9 11.3 262 
 36 Table N. - Upper air atmospheric data. (a) rawinsondes for November 23, 1966 (launch times PST) 1966 11 23 1200:00 1966 11 23 1800:00 Wind Wind Alt Press Temp RH Dir Speed Alt Press Temp RH Dir Speed m kPa C % deg mJs m kPa C % deg mJs 724. 93.10 3.0 82 240 1 724. 93.40 8.1 69 0 0 560. 95.00 ** *** *** *** 582. 95.00 *** *** *** *** 1002. 90.00 5.8 65 255 5 1027. 90.00 4.3 54 34 2 1467. 85.00 2.7 49 225 4 1489. 85.00 1.6 59 40 4 1955. 80.00 -0.1 38 214 4 1976. 80.00 -1.2 63 62 4 2469. 75.00 -1.9 15 332 8 2487. 75.00 -3.8 34 137 1 3014. 70.00 -6.4 18 340 11 3028. 70.00 -7.6 36 239 4 3587. 65.00 -11.5 17 277 8 3600. 65.00 -11.7 40 250 5 4195. 60.00 -15.3 17 280 9 4208. 60.00 -15.7 34 231 10 4848. 55.00 -19.5 17 264 10 4858. 55.00 -20.1 28 223 11 5550. 50.00 -23.3 18 237 18 5562. 50.00 -23.6 23 229 15 6316. 45.00 -27.6 16 228 28 6322. 45.00 .28.8 24 225 21 7152. 40.00 -33.7 15 227 33 7155. 40.00 -34.6 26 223 26 8078. 35.00 -39.3 17 224 38 8077. 35.00 -40.1 26 225 34 9118. 30.00 -45.4 19 223 42 9118. 30.00 -45.1 26 226 41 10318. 25.00 -50.3 20 224 45 10318. 25.00 -50.1 26 224 42 11763. 20.00 -52.6 21 229 38 11768. 20.00 -52.0 25 222 35 12626. 17.50 -53.3 22 231 31 12631. 17.50 -52.7 24 226 29 13614. 15.00 -54.8 20 232 27 13623. 15.00 -53.5 24 227 27 14771. 12.50 -58.9 22 229 27 14789. 12.50 -57.0 25 224 25 16170. 10.00 -59.8 22 225 18 16195. 10.00 -58.7 23 215 16 17567. 8.00 -60.0 22 227 12 18395. 7.00 -60.2 22 230 9 17598. 8.00 -59.3 22 216 12 19363. 6.00 -60.5 22 232 1 18440. 7.00 -57.9 22 212 9 20501. 5.00 -59.5 22 206 11 19405. 6.00 -59.2 22 171 5 21902. 4.00 -59.9 22 152 4 20550. 5.00 -57.5 22 121 2 23708. 3.00 -56.2 22 78 6 21966. 4.00 -58.2 22 104 3 24845. 2.50 -58.1 22 71 6 23787. 3.00 -55.4 22 74 9 26273. 2.00 -55.1 22 68 8 24946. 2.50 -55.2 22 72 13 28134. 1.50 -51.2 22 75 16 26377. 2.00 -51.0 22 82 13 30787. 1.00 -45.7 22 73 14 28270. 1.50 -45.7 22 70 12 33151. 0.70 -41.3 22 43 14 31014. 1.00 -42.3 22 85 14 
 37 Table IV. - Continued (b) rawinsondes for December 12, 1966 (launch times PST) 1966 12 12 1200:00 1966 12 12 1800:00 Wind Wind Alt Press Temp RH Dir Speed Alt Press Temp RH Dir Speed m kPs C % deg m/s m kPa C % deg m/s 724. 94.10 -0.2 83 250 724. 94.20 6.7 68 120 646. 95.00 *** *** *** *** 651. 95.00 *** *** *** *** 1090. 90.00 9.2 61 30 3 1098. 90.00 8.6 52 55 2 1564. 85.00 8.9 35 56 5 1570. 85.00 8.7 31 49 5 2065. 80.00 8.8 12 69 4 2073. 80.00 10.7 13 74 5 2598. 75.00 7.9 10 75 2 2608. 75.00 7.9 14 92 2 3163. 70.00 3.8 14 41 2 3173. 70.00 4.2 16 87 1 3759. 65.00 -1.0 20 336 3 3770. 65.00 -0.9 22 278 5 4392. 60.00 -5.9 17 290 5 4403. 60.00 -6.1 22 275 6 5066. 55.00 -ILl 10 304 5 5076. 55.00 -11.5 18 265 2 5789. 50.00 -16.8 11 303 11 5798. 50.00 -17.0 16 300 6 6571. 45.00 -23.1 11 305 10 6580. 45.00 -23.3 30 285 12 7421. 40.00 -29.9 11 300 9 7430. 40.00 -30.0 36 273 12 8357. 35.00 -37.5 17 291 9 8367. 35.00 -37.1 62 264 11 9400. 30.00 -46.4 28 294 15 9413. 30.00 -46.1 58 261 15 10584. 25.00 -55.9 38 287 22 10595. 25.00 -56.8 55 266 19 11984. 20.00 -60.4 35 270 32 11987. 20.00 -61.2 48 276 22 12812. 17.50 -60.4 32 264 29 12815. 17.50 -61.3 42 274 22 13779. 15.00 -59.4 26 259 28 13769. 15.00 -61.5 34 266 23 14910. 12.50 -62.5 24 257 29 14895. 12.50 -62.5 30 258 23 16274. 10.00 -66.7 22 260 24 16268. 10.00 -63.9 24 260 19 17614. 8.00 -69.8 22 268 16 17622. 8.00 -67.9 22 271 15 18417. 7.00 -67.9 22 274 12 18434. 7.00 -64.0 22 274 12 19346. 6.00 -66.3 22 277 7 19386. 6.00 -62.8 22 278 8 20457. 5.00 -62.8 22 288 6 20514. 5.00 -61.3 22 295 6 21840. 4.00 -60.0 22 313 6 21902. 4.00 -59.6 22 290 3 23646. 3.00 -57.2 22 275 8 23708. 3.00 -57.3 22 281 9 24795. 2,50 -56.8 22 252 9 24849. 2.50 -55.9 22 269 9 26226. 2.00 -55.8 22 250 10 26286. 2.00 -55.2 22 254 7 28134. 1.50 -54.3 22 260 12 30712. 1.00 -50.7 22 254 23 33050. 0.70 -40.1 22 *** *** 
 38 Table IV. - Concluded (c) rawinsondes for December 16, 1966 (launch times PST) 1966 12 16 0900:00 1966 12 16 1800:00 Wind Wind Alt Press Temp RH Dir Speed Alt Press Temp RH Dir Speed m kPa C % deg rn/s m kPa C % deg rn/s 724. 94.10 1.7 60 0 0 724. 94.20 4.7 36 180 1 645. 95.00 *** *** *** *** 654. 95.00 *** *** *** *** 1093. 90.00 10.3 20 59 5 1100. 90.00 9.2 25 70 2 1566. 85.00 8.3 17 73 8 1572. 85.00 7.2 24 64 4 2064. 80.00 6.5 13 86 7 2069. 80.00 6.3 20 104 4 2595. 75.00 7.2 19 126 6 2600. 75.00 6.9 20 135 3 3159. 70.00 3.8 25 127 3 3163. 70.00 3.3 26 156 3 3756. 65.00 0.0 35 148 3 3759. 65.00 -0.7 32 178 5 4401. 60.00 -5.7 42 102 4 4393. 60.00 -5.9 35 190 7 5085. 55.00 -8. 1 15 32 6 5068. 55.00 -9.0 19 187 7 5820. 50.00 -12.6 16 352 6 5802. 50.00 -12.4 16 168 2 6617. 45.00 -17.8 17 359 2 6598. 45.00 -18.4 14 185 3 7484. 40.00 -25.2 19 145 1 7464. 40.00 -25.4 16 169 7 8436. 35.00 -34.3 20 235 3 8417. 35.00 -33.8 18 180 11 9494. 30.00 -43.1 21 227 8 9476. 30.00 -43.1 20 206 14 10695. 25.00 -53.1 23 235 13 10676. 25.00 -53.0 22 228 19 12099. 20.00 -62.8 23 229 16 12086. 20.00 -61.7 23 239 17 12914. 17.50 -65.9 23 232 14 12900. 17.50 -67.2 23 233 16 13846. 15.00 -67.3 23 245 10 13832. 15.00 -66.8 23 229 16 14944. 12.50 -69.8 22 253 10 14939. 12.50 -65.6 22 233 12 16281. 10.00 -69.7 22 274 7 16295. 10.00 -64.7 22 258 8 17614. 8.00 -67.1 22 306 5 17652. 8.00 -66.7 22 292 5 18417. 7.00 -67.1 22 321 5 18460. 7.00 -65.2 22 301 3 19346. 6.00 -65.2 22 339 4 19407. 6.00 -63.3 22 300 3 20465. 5.00 -63.3 22 339 4 20526. 5.00 -62.6 22 345 4 21840. 4.00 -60.7 22 296 7 21902. 4.00 -60.6 22 334 4 23625. 3.00 -60.4 22 258 7 23678. 3.00 -58.8 22 293 7 24770. 2.50 -61.1 22 280 11 24845. 2.50 -58.0 22 288 13 26139. 2.00 -58.4 22 273 15 26273. 2.00 -55.9 22 281 17 27961. 1.50 -54.7 22 263 26 28091. 1.50 -54.7 22 279 18 30636. 1.00 -47.4 22 *** *** 30757. 1.00 -46.9 22 *** *** 
 39 Table v.- Surface weather observations at EAFB runway 22/04 Test date Time, Z Temp, "F Wind, deg/kts Cloud cover Precipitation 11-23-66 17:56:00 44 000/00 clear none 18:56:00 47 000/00 clear none 12-12-66 17:55:00 42 000/00 overcast none 18:55:00 42 110/02 overcast none 12-16-66 14:57:00 27 310/02 scattered none 15:55:00 30 000/00 broken none 
 40 Table VI.- Summary of XB-70 maximum positive overpressures, signature lengths, and periods for in-flight probe tests. Measured Calculated XB-70 F-104 Date MSN Probe r Gage Gage run (ft) 1 2 X t.T t.Pmax Altitude Altitude (ft) (msec) (lb/sq.ft) M (ft. MSL) (ft. MSL) t.P,max, Ibs/sq. ft. 11/23/66 1-1 1 1.45 36,550 39,840 7825 1.02 1.31 529 389 1.05 1 1.50 37,639 32,912 4732 2.80 2.76 198 144 12/12/66 3-2 2 1.49 37,584 32,928 4729 2.80 2.27 220 162 2.60 1 1.47 37,870 36,000 3451 3.40 3.50 252 184 12/16/66 4-2 2 1.46 37,900 35,869 2257 4.12 211 154 4.40 - 3 1.47 37,917 36,115 1896 5.10 4.11 188 138 
 41 Date 11-23-66 12-12-66 12-16-66 Table VII.- XB-70 sonic boom signature characteristics measured at ground level with six microphone cruciform array (Calculated f>P max - 2.40 lb./sq. ft.) Peak amplitudes. Peak Shock positive Amplitudes 'I Wave Ground Alt. Head Boom Micro- Mission Mach Offset (lb/sq.ft.) Negative 2 , t.T t.X angle speed Number It No. deg (n.mi.) time phone (Ib/sqJl) . $ MSL mag (Zulu) No. pt P2* P3* pt pf (msec) (msec) (msec) (ft) (deg) (ft/sec) I-I 37.200 1.46 243 16951 1831:43 MLCI ** ** 3.06 2.29 2.50 0.5 1.5 229 307 58.9 1342 MLC2   2.91 2.34 2.67 0.5 6.5 229 307 MLC3 **  .* 2.78 2.22 2.38 0.2 5.0 229 307 MLC4 ** ** 2.86 2.18 2.64 0.6 2.0 230 309 MLC5   2.94 2.37 2.84 OJ 3.0 231 310 MLC6 •• I 1.44 **  **   **  ** 3-2 37.600 1.50 246 0.15S 1 1831:42 MLCI   2.57 2.27 2.50 OJ 5.0 234 324 56.7 1384 , MLC2 **  ** 2.42 1.79 2039 0.4 5.0 233 322 MLC3   2.39 2.10 2.25 OJ 5.0 234 324 MLC4 **  ** 2.47 2.13 2.51 0.5 4.5 234 324 MLC5   2.53 2.23 2.63 0.5 4.0 234 324 MLC6 •• •• 1.30 • • •• •• **  ** ** 4-2 38.600 1.46 246 0 1557:49 MLCI   2.57 2.24 2.27 0.2 5.0 232 321 55.0 1399 MLC2 **  ** 2.57 2.32 2.29 0.5 5.0 232 321 MLC3   2.49 2.23 2.02 0.5 5.0 232 321 MLC4 **  ** 2.39 2.11 2.29 0.8 6.0 232 321 MLC5   2.43 2.12 2.22 0.6 5.0 232 321 MLC6 **  ** 1.17   **  ** ** ** * See figure 7 for definition. Waveform Category NP NP NP NP NP ** NP NP NP NP NP •• NP NP NP NP NP ** 
 42 (Courtesy of NASA Flight Research Center) Figure 1.- Photograph of XB-70-1 delta wing supersonic aircraft used as the sonic boom generator in the present studies. 
 43

1------- 32 (IDS) ----------i 

,_ '- 9.1 (30) 

\ -~ 

'\ '- I 

I 

57.6 (189) 

Figure 2.- Three-view drawing of XB-70-1 airplane. Dimensions are in meters (feet). Total wing 

area is 6297.8 feet (from ref. 16) 
 44 (a) windshield-nose ramp down (b) windshield-nose ramp up Figure 3.- Photographs of XB-70 showing windshield-nose ramp positions. 
 45

300 

6_900 

C\I 

250 ..... 

-

cci' 

(l) 200 ..... 

al 

al 

c 150 0 

. .,. 

U 

(l) 

en 100 I 

en 

en 

0 

..... 

u 50 

0 

0 500 1000 1500 2000 2500 

fuselage effective length, in. 

(a) Total area distribution based on oblique cuts for a position above and to the side of the aircraft (8 = 25°) 

C\I ..... 

-

cci' 

(l) 

..... 

al 

al 

c 

0 

..... 

U 

(l) 

en 

I 

en 

en 

0 

..... 

u 

300 

6_900 

250 

200 

150 

100 

50 

0 

0 500 1000 1500 2000 2500 

fuselage effective length, in. 

(b) Total area distribution based on oblique cuts for a position below the aircraft (8 = -90°) 

Figure 4.- Area distributions of XB-70-1 vehicle used as shock-wave generating airplane. Oblique 

cuts at Mach 1.5 (inlet capture area not included. Wing tips down at 65° down). 
 46 (a) Probe airplane (b) Probe used for in-flight pressure measurements Figure 5.- F-I04 fighter airplane with nose-boom probe installation for measuring the shock-flow-field in the vicinity of the disturbance-generating XB-70-1 airplane. 
 47 Figure 6.- Photographs of the in-flight recording instrumentation mounted in the F-I04 access bays. 
 48 !------ -------ROIGlabll port jon of IIrobt-----------------! _ __________ F'i •• d pori/on 01 piobt ______ ______ ___ c.'ifleel (1I)ColIal2) Alt"l"CIIe r--obe lip Z Orifices prOvidinoj PI ! \"provl6rQ prIm ... P2 fOf 'lor rel!rtnce "de' of <J09II1 .'" IIdt 0100l}ll2 Clrculor Cant ,I Inctudtd Cl'MjI' ..... . ....... " Orillees pr.lwrl 

. ... \lOfrtltrl!l'lCe side 

\, \ ,. -GOQII 2 ..... 

';' " \C,· ... O'· E, .,/ - ",,,,',., <l> "TTl I 

. . , - I." -r 'Y 

-,-. U-\Uil .. -! .,.... =:T' .'-""\- . . '000 

=t!=8 01 '1 "'- ±l 

','" eifeuklr CQ'ICI ,,- included arq. 

- 3" 

IO.lO ".00 BoIlOm prj/'etS , L I :Ui for IllrplGllll II1SlrlllT'ott\l, 2.50 Pi lot. V'"' .... (side 0I\d fop orihcu cl(md) I PlIUI.II'I 2501.0." ... ;:. ,' ., ' 

. Mll .... / / .' 1' •.•• • BoHom I 

o .0> .om. ,-•. '-8.CQ-0; .0'" Q' . "'. 

_

_

_

_ __

_ __ .j. 

,,' 

Z".). 

Z ".). >-.(".> Sec.B-B Sec.c-e Sec.O-O SIt.£-E 

Ir'b-qI!mtflll 

. . , 

0.' 0.1811 '" O.4!1O:1: ., "" ,., ..... ,. '" , .. .4988 

_______________ 8, 

f , . . ", '.7 .Ue2 • n .U01 .. .4"2 .. ,.• .. """ 

4 !,:, 

!ite. A-A Sec. a-a 

Figure 7. - Principal details and dimensions of full scale probe used for in-flight measurements and for wind tunnel tests at a 

Mach number of about 2.01 (rotatable portion of probe is positioned for probe-airplane flight over generating airplane). 

Dimensions are in inches. 
 49 /. 375 VJ2ILL .(,/N£ L-- Z>£I.LL, TAP &- C'80R£ AS S,L/(HMiV .,c=-oR #6 - 40 No,L-; S .tiD. Sc. Z P.GRCES Figure 8.- Adaptor required to mate specially instrumented nose-boom pressure probe used originally on an F-106 to the NASA F-I04 used in present tests. 
t .. -" 

" " " o , '0 

$C.,[-",".' 

(a) ~ncral .yout of probe mi .. ior.' relative to main te" ano._ 

, 

, ' 

f ' I ' :v---',p-----%- --~ ~. , I . FRC ,Y-""-, . , , : L\C . -<>-<) ~+~----#----®- , / I I 

." -- • , , 

®-- - ---4}--- -~

. , !: I ' 

, t +' -+'" , 

, itn-t- U , =- "',. -

(b) Ground bUed microphone l. yooUl .. mlin tesl al't ... 

Figw-o 9.- Scbemalic of XB· 70 ond. F-l04 nominol groood trac4 Ihowing ..-ea In which probe 

mis,ion. were Hown Jnd ground "", .. uremenl wen: ocquiud. 

50
 51

s 

---

NOTE: S, Y and Z are distances between the radar transponders on the 

XB-70 and F-I04. The XB-70 transponder is located on the 

bottom of the airplane and 102 feet behind the nose. The F-I04 

transponder is located approximately 10 feet behind the tip of 

the probe. 

--

(a) XB-70 and F-104 coordinate system at time of penetration. 

z 

6000 

\ F-l04 probe aircraft 4000 

Flights of 11(23/66 - Run 1 ... 

2000 

y y 

2000 

Flights of 12/16/66 -Runs 

4000 

Flights of 12/12/66 -Runs 6000 

z 

(b) Rear view in-flight direction showing F-I04 probe runs relative to XB-70 flow field . 

Figure 10.- Sketches illustrating general position of probe aircraft wid generating aircraft. 
 52 0.5 seconds 

..... ..... 

. . .. .. " .......... -... •.. .. ....... .'-.. .. ,'.-

reference 

Pass 1, XB-70 overtakes F-104 probe aircraft (Z = -3290 ft, Y = 71 00 ft., rll = 42) 

Figure 11.- Copy of November 23, 1966 film trace showing in-flight time histories of differential pressures 

measured in flow-field above XB-70 aircraft. 
 53 f--- ---j 0.5 seconds time gage 1 

-

.. .. 

----.-. ---. " - .- / - gage 2 

" ' - . . -.-

/ reference 

(a) Pass I, XB-70 overtakes F - 104 probe aircraft (Z = 4727 ft., Y = 220 ft., r/l = 26) 

I .. time 

0.5 seconds 

gage 1 

/ reference 

(b) Pass 2, F-104 probe aircraft overtakes XB-70 (Z = 4656 ft., Y = 825 ft., r/l = 26) 

Figure 12.- Copy of December 12, 1966, film traces showing in-flight time histories of differential pressures measured in 

flow-field below XB-70 aircraft. 
 54

I" 

0.5 seconds 

bow shock 

/L//1/~ 1- gage J 

l/:>v/L\ ... F"'" 

' .. .. ... ""l / ,.... '.. .~ \ /' 

-' . 1 ~ reference 

X = 252 ft. ----.j '== 

(a) Pass 1, XB-70 overtakes F-104 probe aircraft (Z = 1870 ft., Y = 2900 ft., r/l = 19) 

I.. I 

0.5 seconds 

". 

1 tail shock 

; _ gageJ 

r gage 2. 

(b) Pass 2, F-104 probe aircraft overtakes XB-70 (Z = 2031 ft., Y = 980 ft., r/l = 12) 

I" I 0.5 seconds 

'- gage 2 

. "".:,' 

.. 1 .. ----_ X = 188 ft. ___ ~ j'--reference 

================._==~==f::.=-=== 

(c) Pass 3, XB· 70 overtakes F-104 probe aircraft (Z = 1802 ft., Y = 590 ft., r/l = 10) 

Figure 13.· Copy of December 16, 1966, film traces showing in-flight time histories of differential 

nrr.RRurr.~ mr.R~nrr.rl in Aow-nr.lrl hr.low XR-70 Rirr.rRft 
 55

(a) Pass 1 on Nov. 23, 1966. Penetration time = 18:27:45 Z 

(b) Pass 1 on Dec. 12, 1966. Penetration time = 18:27:32 Z 

(c) Pass 2 on Dec. 12, 1966. Penetration time = 18:29: 18 Z 

Figure 14.- XB-70 flow-field shock-wave signature overpressures. 
 56
 57 350 .t: ><:" 300 '" ..>:: u 0 ..c: 250 '" • 

-

'o;j • • "0 200 til 

0 150 .0 

- - - - - - - - - a - - - - - - - - - - - - - - • - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - . -

Generating-airplane length. I 

as 

Q) Far-field volume theory, M=1.5 

100 

Q) 

.0 

Q) 

u 50 til '" :.a 

0 

0.0 2.0 4.0 6.0 8.0 

r, ft 

Figure 15.- Comparison of measured and calculated distances between bow wave and tail wave of generating airplane. 
 58

v , 

MSN 1-1- Nov. 23,1966 

Mach No. = 1.46 

Altitude = 37,200 ft. MSL 

Offset = 1.69S 

Boom Time = 18:31:43 Z 

---,- Llp = 2.61 psf ; ---......r - ------ FRe 1 

---- --.r---------

L '-----< 224ms '1 

MSN 3-2 - Dec. 12 1966 

Mach No. = 1.5 

Altitude = 37,600 ft. MSL 

Offset = 0.15S ' 

Boom Time = 18:31:42 Z 

• 2.57 f r..-, ............... f--..... cruciform up . ps --.. ----........ MLCI ----L-J _________ 

234ms ·1 

\ i LAC31 

F---------

I 'J 

[ . 1. 234ms al 

MSN 4-2 - Nov. 23, 1966 

Mach No. = 1.46 

Altitude = 38,600 ft. MSL 

Offset = 0 

Boom Time = 15:57:49 Z 

cruciform 

MLCI 

Figure 16.- XB-70 measured sonic boom signatures at ground level following in-flight probe tests. 
 59

8T and X -------.~I 

-,---

t-~ -- P,- N 

. T 

-L 

TYPE 

2 NP 7 SPR 

4 

TT 

,;r~, ~lll I 

NR 

P 

PP 

IT 

-h-~ ~_ ri·----~ B 

Figure 17.- Diagrams of wavefonns and signature characteristics which represent the various 

categories of measured sonic booms. 

PR 

R 

co 
 60

<'>p - 1.21blft.2 T 

__ M = 1.5 @ 37000 ft. 

~ 

<'>p - 2.5 Iblft.2 ground level 

Figure 18.- Planform and side views of bomber airplane with time history of pressure signature as 

measured above and below the airplane. Signature length has been adjusted to make distance between 

nose and tail shocks approximately the same as the airplane length. 
r .10 sec --I 

(a) Entire system (pressure gage and galvanometer). 

v 

(b) Galvanometer only. 

Figure 19.- Response characteristics of pressure instrumentation used for in-flight measurements. 

61
 62 p,'p P2 P3 ',ref P, = primary system of orifices (location I) p 2,r8f P2 = secondary system of orifices (Iocotion 2) p = system of orifices (location 3) provid ing static ;3 pressure for flight instruments 1.04 1.02 p ..0- -<J "" -'l. J:Y -- , P", 1.00 .98 fI .94 :; ': 111 t11f1 H{+ttl .96 1.0 1.0 P, ,reI 2 -p;-1.0 0 .98 .76 p' t .74 If .72 ,/'.J / rV P' t ..C>- 1--- -<;J' d±ftff±±Hfffl -Q:' -- ..0- -- "'-' '" ...0- 0- -0 - - 1.92 1.94.8 E .96 E I I . 98 2.00'§ 2.02....J 2.04 -4 -2 0 2 4 6 8 10 -2 o 246 8 10 Angle of otfack,a,deg AnQle of ottock .. Q,deO Figure 20.- Steady-state calibration of flight probe at angles of attack from -2.5° to 9.5°, as obtained from tests in the Langley 4-foot supersonic pressure tunnel (pressure orifices at bottom of probe). '" 2.01; '" 185 lb./sq. ft. 
 63

Bod y ordinates 

Diom.=O.lOO x r x 

~.~> . . t 

0.00 0.000 1.00 

--.s., . 10 .038 1.25 

§1.87~ . .25 .088 1.50 

.50 .150 1.75 

2.00 .75 .188 1.87 

(a) Disturbance-generating body of revolution (same shape as model D of ref. 27) 

~- .. - Pilot -tube location 

Static-pressure orifices (dlam. of orifice 1·0·020 in. ; diem. of orifices 2,3, and 4· 0.031 in.) 

I 23 4 

, :: 

I ,':' 

..... -- - - - Plate surface smooth for distance of 3.00 in. from leodirg edge 

.----~ -----.----.----.----------.-----

Plan view 

0.200 

.188 

.150 

.088 

. 050 

./.:;.::: ~ ~ } Stotic - pressure orifices 

~'~' ..... . t Ii J.....-1 ~ #-.090 0.0. tubing 

.~ .. {i CfJlj...;,...' '__ .J! II 

1.00 062 10· 

.125~ 0-1 I 

:' .00~ 

l. ".- Pi1ot- pressure tube 

,r-- ____ Silver solder 

-L- .040 ! 

.060 .d5E~:tJj-~~::~¥\ -:;;; Side T T:U:~ .........L. 

.o!~ot~~~g Pion 

Details of pitot - tube end 

_____ ---Ij.4~toL 12.35 ~ 

Side elevotion 

(b) Survey instrument for measuring pressure changes across body-generated disturbance. 

Figure 21.- Wind-tunnel apparatus and test arrangement for generating and determining the strength 

of an axisymmetrical disturbance used in obtaining experimental evidence concerning the reflection 

characteristics of the flight probe. Dimensions are in inches. 
 64 ___ ,_ . ___ Body of revolution ;a"" O· -,: .-. - - - - - - - - Survey instrument Bow wove ---- -- Pitot - pressure tube Stotic- pressure orifices - ---./ (c) Arrangement for measuring pressure changes across disturbance by use of survey instrument oriented to minimize wave-reflection effects. Figure 21.- Concluded. 
 65

PI 

Pm 

1.08 

1.06 

-~ 

~ 

1.04 6 Estimated 

pressure 1 rise 

(refs. 26 and 27) 

1.02 

I--

1.00 

.98 

.96 -4 

1--1 1 ( 

y=7.375 In. 

-2 o 

f( 

b 

\ 

1.06 

1.04 

1.02 

PI 

Pm 

1.00 

~ . 0 

•• - ------Body of revolution; a'" 0 

~ 1.06 

~ 1.04 ~ 

~ ~ 1\ 

\ 1.02 

PI 

Pm ~ 

1.00 

1\ 

"\ 1.04 

l'q 1\ ~ 1.02 

~ 

~ -\ 1.00 

,......Q f.ioc Ib .98 .98 \ y= 13.375 in. y= 17.375 in. y:13.375 in. 

_96 .96 .96 L.. ____ _ ... __ 

2 -2 0 2 -2 0 2 -2 o Approximate longitudinal distance from bow wave. xs' in. 

1D4 

1.02 

1.00 

~ 

Pm 

.98 

.96 2 -2 

~ 

r-- '\ I~ ~ 

; ( f\ 

IOU: 0 

y= 19.375 in . 

o 

(a) Probe-indicated pressure changes across body-generated bow wave. 

.08 

I 

.06 

6p .04 Pm 

.02 

0 2 4 

- - - - Estimates based on far-field volume theory (refs. 26 and 27) 

o 

o 

o 

Surveys 

Primary system of orifices \locatio~ I) } Probe meosurements 

Secondary system of Orifices (iocallon 2) 

o 

<> 

0- --

6 8 10 12 14 16 18 20 2 2 24 26 

Approximate separation distance be_en body of revolution and flight probe or survey plate ,y,in. 28 

(b) Comparisons of estimated and measured maximum pressure rises across bow wave. 

2 

30 

Figure 22.- Flight-probe capability for sensing static pressure changes across an axisymmetrical 

disturbance (bow wave generated by body of revolution), as evidenced by comparisons of probeindicated, survey-indicated, and estimated pressure changes across bow wave. Moo:::: 2.01 
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Contractor Report

 4. TITLE AND SUBTITLE

Measured Sonic Boom Signatures Above and Below the XB-70 Airplane 

Flying at Mach 1.5 and 37,000 Feet

5a. CONTRACT NUMBER

NNL07AE38T

 6. AUTHOR(S)

Maglieri, Domenic J.; Henderson, Herbert R.; Tinetti, Ana F.

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center Eagle Aeronautics, Inc.

Hampton, VA 23681-2199 13 West Mercury Boulevard

 Hampton, VA 23669-2508

 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

Langley Technical Monitor: Peter G. Coen

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Unclassified - Unlimited

Subject Category 71 

Availability: NASA CASI (443) 757-5802

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email: help@sti.nasa.gov)

14. ABSTRACT

During the 1966-67 Edwards Air Force Base (EAFB) National Sonic Boom Evaluation Program, a series of in-flight flow-field 

measurements were made above and below the USAF XB-70 using an instrumented NASA F-104 aircraft with a specially designed nose 

probe. These were accomplished in the three XB-70 flights at about Mach 1.5 at about 37,000 ft. and gross weights of about 350,000 lbs. Six 

supersonic passes with the F-104 probe aircraft were made through the XB-70 shock flow-field; one above and five below the XB-70. 

Separation distances ranged from about 3000 ft. above and 7000 ft. to the side of the XB-70 and about 2000 ft. and 5000 ft. below the 

XB-70. Complex near-field "sawtooth-type" signatures were observed in all cases. At ground level, the XB-70 shock waves had not

coalesced into the two-shock classical sonic boom N-wave signature, but contained three shocks. Included in this report is a description of 

the generating and probe airplanes, the in-flight and ground pressure measuring instrumentation, the flight test procedure and aircraft 

positioning, surface and upper air weather observations, and the six in-flight pressure signatures from the three flights.

15. SUBJECT TERMS

Sonic boom; Sonic boom signatures; XB-70; Supersonic; Shock waves

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(443) 757-5802

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01- 04 - 2011