March 2011 

NASA/CR–2011-217078 

Equivalent Longitudinal Area Distributions of 

the B-58 and XB-70-1 Airplanes for Use in 

Wave Drag and Sonic Boom Calculations 

 

Ana F. Tinetti 

NCI Information Systems, Inc., Hampton, Virginia 

Domenic J. Maglieri, Cornelius Driver, and Percy J. Bobbitt 

Eagle Aeronautics, Inc., Hampton, Virginia 
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March 2011 

NASA/CR–2011-217078 

Equivalent Longitudinal Area Distributions of 

the B-58 and XB-70-1 Airplanes for Use in 

Wave Drag and Sonic Boom Calculations 

 

Ana F. Tinetti 

NCI Information Systems, Inc., Hampton, Virginia 

Domenic J. Maglieri, Cornelius Driver, and Percy J. Bobbitt 

Eagle Aeronautics, Inc., Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

 

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National Aeronautics and Space Administration. 
1

Equivalent Longitudinal Area Distributions of the B58

and XB-70-1 Airplanes for Use in Wave Drag

and Sonic Boom Calculations

Ana F. Tinetti, Domenic J. Maglieri,

Cornelius Driver, and Percy J. Bobbitt

ABSTRACT

A detailed geometric description, in wave drag format, has been developed for the Convair B-58 

and North American XB-70-1 delta-wing airplanes. These descriptions have been placed on electronic files at the NASA Langley Research Center, the contents of which are described in the 

present paper. They are intended for use in wave drag and sonic boom calculations.

Included on the electronic file and in the present paper are photographs and three-view drawings 

of the two airplanes, tabulated geometric descriptions of each vehicle and its components, and 

comparisons of the electronic file outputs with existing data. The comparison includes a pictorial 

of the two airplanes based on the present geometric descriptions contained on the electronic files 

and a comparison of the cross-sectional area distributions for both the normal Mach cuts and 

oblique Mach above and below the vehicles. Good correlation exists between the area distributions generated in the late 1950s and 1960s and the present files.

The availability of the present electronic files allows for further validation of existing sonic boom 

prediction codes through the use of two existing experimental data bases on these two airplanes. 

These data bases were acquired in the early and mid 1960s time period and, to date, have not been 

fully exploited. These two data bases consist of in-flight measurements of the supersonic flowfields above and below the B-58 and XB-70-1 airplanes, acquired in 1963 and 1966, respectively, 

at distances of from about 10 to 95 body lengths.
2

TABLE OF CONTENTS

ABSTRACT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

TABLE OF CONTENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

LIST OF TABLES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

LIST OF FIGURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

SYMBOLS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

TEST AIRCRAFT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

INPUTS TO ELECTRONICS FILES DESCRIBING AIRPLANE GEOMETRIES . . . . . . . . . . 6

COMPARISON OF ELECTRONIC FILE OUTPUTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

Airplane Description. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

B-58 Cross-Sectional Area Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

Wind Tunnel Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7

Full-Scale Airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

Oblique Mach Cut. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

XB-70-1 Cross-Sectional Area Distributions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

Wind Tunnel Model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

Full-Scale Airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

Oblique Mach Cut. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

DESCRIPTION OF ELECTRONIC FILE CONTAINING AIRCRAFT GEOMETRY

AND WAVE-DRAG DEFINITIONS OF B-58 AND XB-70-1 AIRPLANES . . . . . . . . . . . . . . 10

SUMMARY REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

TABLES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

FIGURES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
3

LIST OF TABLES

TABLE I.- GEOMETRIC CHARACTERISTICS OF THE CONVAIR B-58 AIRPLANE . . . . 13

TABLE II.- GEOMETRIC CHARACTERISTICS OF THE XB-70-1 AIRPLANE . . . . . . . . . . 16
4

LIST OF FIGURES

Figure 1.- Photographs of delta-wing airplanes to be geometrically described. . . . . . . . . . . . . . 21

Figure 2.- Three-view drawings of delta-wing airplanes to be geometrically described. . . . . . . 22

Figure 3.- Pictorials of B-58 and XB-70 as generated from current electronic files 

containing geometric descriptions of both airplanes. . . . . . . . . . . . . . . . . . . . . . . . . . 23

Figure 4.- Comparison of normal cross-sectional area distributions of B-58 wind tunnel

model and components (inlet capture area not included). . . . . . . . . . . . . . . . . . . . . . 24

Figure 5.- Comparison of normal cross-sectional area distributions of B-58 with and 

without MB-1 fuselage pod (inlet capture area not included). . . . . . . . . . . . . . . . . . . 25

Figure 6.- Comparison of total area distributions above and below B-58 airplane with 

MB-1 pod at M=1.65 (inlet capture area not included). . . . . . . . . . . . . . . . . . . . . . . . 26

Figure 7.- Comparison of non-dimensional total normal cross-sectional area distributions

of XB-70-1 airplane. Inlet capture area not included. Wing tips at 65o

. . . . . . . . . . . 27

Figure 8.- Comparison of total normal cross-sectional area distributions of XB-70-1 airplane.

Inlet capture area included. Wing tips at 0o.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

Figure 9.- Comparison of normal cross-sectional area distributions of XB-70-1 airplane

and components. Wing tips at 0o. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

Figure 10.- 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 at 65o down). . 29
5

INTRODUCTION

In 1963, the USAF and NASA conducted flight tests to define the supersonic flow field above and 

below a B-58 delta-wing bomber airplane. A specially instrumented F-106 aircraft was used to 

“probe” the B-58 flow field at distances of about 14 to 95 body lengths from the B-58 (ref. 1). 

During the 1966-1967 EAFB National Sonic Boom Evaluation Program (ref. 2), the USAF and 

NASA conducted flight tests of an F-104 probing above and below the supersonic flow-field of 

the much larger XB-70-1 delta-wing bomber at distances of about 10 to 42 body lengths. The purpose of these in-flight measurements was to add to the sonic boom data base being used to validate existing sonic boom prediction codes.

Little use was made of the 1963 and 1966 probe measurements of the B-58 and XB-70 flow-field 

signatures in terms of sonic boom theory validation. This was due, primarily, to the lack of sufficient details of the B-58 and XB-70-1 geometric and aerodynamic descriptions and, in part, to the 

availability of the details of the XB-70-1 probe measurements. The B-58 probe tests, however, 

were reported in full detail in reference 1. Although the XB-70-1/F-104 in-flight probe measurement effort was successfully completed, the results were never formally documented. They 

appeared only briefly in a few reports in preliminary form to reflect the general nature of the flight 

test results.

The need to formally document the 1966 XB-70-1 probe flight tests and to provide geometric 

details of the B-58 and XB-70-1 airplanes was identified within the NASA High Speed Research 

(HSR) Program and funds were made available to accomplish these two tasks. Formal documentation of the 1966 XB-70-1 probe tests has been completed and reported in reference 3.

Detailed geometric descriptions of the B-58 and XB-70-1 airplanes have been completed and are 

presently available in a wave drag format on electronic file at the NASA Langley Reseaarch Center (LaRC). The purpose of this report is to provide an overview and description of the information contained on these electronic files. Included are photographs and three-view drawings of the 

two airplanes, tabulated geometric descriptions of each vehicle and its components, and comparisons of the electronic file outputs with existing data. These comparisons include a pictorial of the 

two aircraft as generated by the present geometric description file, normal cross-sectional areas of 

the complete airplane and each component, and total cross-sectional areas above and below the 

vehicles at oblique Mach cuts corresponding to the flight test data. These results are compared 

with those generated in the late 1950s and early 1960s time period prior to the existence of the 

present computational capability.

SYMBOLS

A Cross-sectional area of airplane obtained by normal or oblique cuts, sq. ft.

l Airplane reference length, ft

M Airplane Mach number

X Cylindrical coordinate measured along body axis, ft

θ Angle measured from horizontal (-90o under airplane, + 90o above)
6

TEST AIRCRAFT

Photographs of the USAF Convair B-58 and North American XB-70-1 delta-wing airplanes are 

presented in figure 1; three-view drawings of each aircraft are shown in figure 2. Detailed geometric characteristics of the B-58 airplane based upon the 1/5- and 1/40-scale wind tunnel models 

described in references 4 to 7 are provided in Table I. The geometric characteristics of the XB-70-

1 airplane taken from reference 8 are presented in Table II. Information contained in Tables I and 

II, along with aerodynamic dimensional data contained in references 8 to 10 allow for an accurate 

and detailed geometric description of the B-58 and XB-70-1 airplanes.

The Convair B-58 delta-wing airplane (figs. 1a and 2a) has a length of 96.8 feet (from nose to tip 

of tail), a wing span of 56.8 feet, and a total wing area of 1542 square feet. Aircraft weight at 

brake release for the in-flight probe tests of reference 1 ranged from about 135,000 pounds to 

145,000 pounds. During the actual probe runs, the B-58 gross weight ranged from 84,000 pounds 

to 115,000 pounds. For all the probe flights of reference 1, the aircraft was configured with the 

MB-1 fuselage pod as shown in figures 1a and 2a. Engines were at 104 percent RPM and exhaust 

nozzles were in partial afterburner. The aircraft was powered by four GE-J-79 turbojet engines, 

each producing 15,600 pounds of thrust with full afterburner.

The North American XB-70-1 delta-wing airplane (figs. 1b and 2b) has a length of 189 feet 

(including noseboom), 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 of reference 3 ranged from about 529,000 

pounds to 536,000 pounds. During the actual probe runs, the XB-70-1 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. The bypass was set at 400 square inches, all six 

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

was powered by six YJ93-GE-3 turbojet engines, each producing 31,000 pounds thrust with full 

afterburning. 

INPUTS TO ELECTRONIC FILES DESCRIBING AIRPLANE GEOMETRIES

Data from references 4 to 7 and 8 to 10 were used, respectively, to describe the geometries of the 

B-58 and XB-70-1 aircraft in the wave drag format of reference 11. In the case of the B-58, the 

details of the airplane geometry are based upon 1/15-scale wind tunnel (ref. 4), free-flight models 

(ref. 5), and 1/40-scale wind tunnel models (refs. 6 and 7). Geometric descriptions were obtained 

from three-view dimensional drawings, cross-sections, and from tabulations contained in the 

reports. 

With the exception of the MB-1 pod and the tail, only the left half of the aircraft has been 

described. The fuselage was described using 23 radial locations per side at 20 longitudinal stations; the wing using 17 chord stations at 7 spanwise stations; the nacelles using 37 radial locations at 14 longitudinal stations; the MB-1 pod using 37 radial locations at 21 longitudinal 

stations; pod fins using 23 chord stations at 2 spanwise locations; and the airplane vertical tail 

using 23 chord stations at 4 spanwise locations. Nacelle pylons and main landing gear fairings 

were also defined.
7

In the case of the XB-70-1, the details of the airplane geometry are based upon the documentation 

of the full-scale vehicle in references 8 to 10. Geometric descriptions were obtained from the 

three-view dimensional drawings, cross-sections, and from tabulations contained in the reports. 

Only the left half of the aircraft has been described. Because of the lack of cross-sectional information, the fuselage was described using elliptical contours with 19 radial locations at 30 longitudinal stations. The wing was described using 20 chord stations at 7 spanwise stations (9 spanwise 

stations for the wing with tips drooped 65o), the canard using 24 chordwise stations at exposed 

root and tip; the vertical tail using 23 chord stations at 3 span stations; and the duct body using 33 

radial locations at 11 longitudinal stations. The lower wedge was also described.

COMPARISONS OF ELECTRONIC FILE OUTPUTS

Figures 3 through 10 have been generated from the present electronic files that describe the 

detailed geometries of the B-58 and XB-70-1 airplanes. These figures are intended to illustrate the 

capability and accuracy of the present electronic files in providing the geometric inputs required 

to perform wave drag or sonic boom calculations.

Airplane Description

Pictorials of the B-58 and XB-70-1, as generated from the current electronic files containing their 

geometric descriptions, are presented in figures 3a and 3b, respectively. The NASA Langley 

“Viewer” program (ref. 12) was used to generate these isometrics. “Viewer” is an Open Windows 

based XView application that displays and prints geometries from multiple formats. Good comparison is noted between the airplane pictorials of figure 3 and the photographs and three-views of 

these aircraft shown in figures 1 and 2, respectively. Note, too, the details of the various airplane 

components such as inlets and nacelles, the fuselage pod, and the vertical tails.

B-58 Cross-Sectional Area Distributions

Wind tunnel model. - A comparison of the normal cross-sectional area distribution for the 1/40 

scale wind tunnel model of reference 6, using the present vehicle geometric description, is given 

in figure 4. Shown on the two plots are the area distributions of the wind tunnel model components (fuselage, wing, nacelles, pod, etc.). The uppermost curve represents the total cross-sectional area distribution of all the components. In figure 4a, from reference 6, the total area curve is 

a sequential buildup of each of the model components, beginning with the fuselage, wing, 

nacelles, vertical tail, landing gear fairings, and the MB-1 pod. Prior to the availability of computational means, the cross-sectional area distributions of aircraft configurations were generated by 

immersing each of the model components and then the complete model into a tank and measuring 

the liquid displacement. Another method was to build the model and components out of balsa 

wood and then make normal or oblique saw cuts and measure the resulting cross-sectional areas.

In figure 4b, generated using the present electronic file of the vehicle geometric description, normal cross-sections are shown for each wind tunnel model component. The current wave drag program (ref. 11), however, does not provide for the sequential buildup of the components in forming 

the total area distribution. Note, too, that the area developments of the fuselage and nacelles were 
8

truncated at their end termination points to simulate the base areas associated with the fuselage 

sting support and nacelle exits of the wind tunnel model. 

Good correlation is seen to exist between the area development generated in 1956 (fig. 4a) and the 

present electronic file data base (fig. 4b). This can be readily seen by directly comparing the area 

distributions for the fuselage, wing, and the total area curves. In fact, if an overlay is made of the 

two data sets, nearly complete correlation exists when the curves of figure 4b are shifted to the left 

by about one-half inch, in the abscissa.

Full-scale airplane. - Normal cross-sectional areas distributions for the full-scale B-58 airplane, 

with and without the MB-1 fuselage pod, are presented in non-dimensional form in figure 5. The 

curves shown in figure 5a, taken from reference 13, were generated in 1961 at LaRC. The curves 

of figure 5b were generated using the present electronic files. Total cross-sectional area distributions for the airplane with and without the MB-1 pod are provided since sonic boom signatures 

have been obtained on each configuration. The in-flight flow-field pressure signature measurements reported in reference 1 are taken with the pod on the airplane and the ground level sonic 

boom signatures reported in reference 13 are with the pod off. Adding the pod to the airplane 

increases the total cross-sectional area; however, it also results in a much smoother curve of the 

total area buildup. Good agreement exists between the curves of figure 5a, calculated in 1961 and 

representing the normal cross-sectional areas for the B-58 airplane with and without the MB-1 

pod, and those shown in figure 5b, obtained using the present electronic files of the B-58 geometric description.

Oblique Mach cut. - Area distributions based on an oblique cut for positions above and below the 

B-58 airplane with the MB-1 pod are presented in figure 6. The oblique cut, made for a Mach 

number of 1.65, is representative of the flight conditions of the B-58 during the in-flight probe 

experiments of reference 1. Figure 6a and 6b, respectively, represent the total area distributions 

for positions directly above (θ = 90o) and below (θ = -90o) the airplane. The solid curves on each 

plot were calculated in 1963 and are taken from reference 1. The dashed curves are based on the 

electronic files of the present report.

It should be noted that these oblique cut area distribution plots are in non-dimensional form. This 

format is usually applied to “normal” Mach cuts, where both the “physical” and “effective” 

length of the aircraft are the same. For oblique cuts, the “physical” and “effective” aircraft lengths 

will be different. In fact, for the Mach 1.65 cut on the B-58, the “effective” length is larger than 

the “physical” length for positions below the aircraft and shorter for positions above the aircraft. 

In order to make comparisons with the 1963 probe flight measurements (ref. 1), the physical 

length of the airplane was used. Another feature to be observed in the 1963 curves of figure 5 is 

that the area distribution goes to zero at X/l = 1.0. This results from the fact that the inlet capture 

area was not included.

XB-70 Cross-Sectional Area Distributions

Wind tunnel model. - A comparison of the normal cross-sectional area distribution for the 

0.000454 scale wind tunnel model of reference 14 to that generated using the present vehicle geometric description is given in figure 7. The plot is in non-dimensional format. Inlet capture area is 
9

not included and wing tips are drooped to 65o. Very good correlation is seen to exist between the 

total area development of the 1963 wind tunnel model (solid curve) and that resulting from the 

present electronic file data base. It should also be noted that the curves do not close to zero at X/l

= 1.0. This is because inlet capture area is not included in the area developments, and results in a 

base drag at the end of the engine exhaust pack.

Full scale airplane. - In figure 8 is presented a comparison of the total normal cross-sectional area 

distribution of the XB-70-1 airplane, as generated in 1961 and reported in reference 10, with that 

obtained using the present vehicle geometric description contained in the electronic files. The 

plots are in dimensional format. Inlet capture area is included and wing tips are not deflected. 

Good correlation is seen to exist between the shape for the total area development generated by 

North American in 1961 (ref. 10) and that resulting from the present geometry. A difference exists 

in the absolute values because of the manner in which the present method calculates the capture 

area. Also note that the curves close to zero area since the inlet capture area is included in the total 

area development. 

Figure 9 compares of the XB-70-1 normal cross-sectional area distributions for the complete airplane and each of its major components for the configuration generated in 1961 by North American (ref. 10), and the configuration from the present electronic file description. Shown on both 

plots are the area distributions of each vehicle component and a total airplane curve representing 

the summation of all these components. Note that, unlike the comparison of the normal cross-sections for the B-58 (see fig. 4), the area distributions for the wing and ducts (and thus the totals) 

obtained from reference 10 (fig. 9a) are quite different from those generated using the present 

geometric description (fig. 9b). The former include only the exposed wing in the wing cross-sectional area. The rest of the area and the inlet capture area (non-flow-through ducts) is attributed to 

the ducts. The latter includes the portion of the wing area covered by the duct body in the wing 

geometric definition instead of assigning it to the ducts, thus greatly reducing the effort needed for 

properly defining the entire aircraft. In addition, the inlet capture area was removed from the total 

duct area. The areas associated with the remaining components (fuselage, canards, tails, lower 

wedge) compare well.

Oblique Mach cut. - Area distributions based on an oblique cut for a position above and to the side 

(θ = 25o) and for positions below (θ = -90o) the XB-70-1 airplane are presented in figure 10. The 

oblique cut, made for a Mach number of 1.5, is representative of the flight conditions of the XB70-1 during the in-flight probe experiments of reference 3. Unlike the previous curves shown for 

the B-58 airplane (see fig. 6), the XB-70-1 curves are in dimensional form. The abscissa represents the “effective” length of the vehicle. 

As seen in figure 10, the area developments are quite different for a position above and to the side 

of the airplane as compared to a position below the airplane in their shape, total area, and location 

of the maximum area value. It is also of interest to note the difference in area distributions 

between a Mach 1.5 oblique cut and that associated with a normal Mach 1.0 cut (see figs. 7 and 

8).

During the generation of the oblique cut total area distributions of figure 10 it was found that, for 

the θ = -90o position (below the aircraft), a sharp discontinuity (spike) appeared on the area devel-
10

opment aft of the maximum area at an affective fuselage length of about 2000 inches. This 

“spike” is believed to occur when the area cuts become coincident with some portion of the vehicle (for example, the wing leading edge). Discussion with NASA Langley personnel who are 

familiar with the wave drag program (ref. 11) noted that such a peculiarity is not uncommon. 

When it occurs, the area curve is “faired” through the “spike”, as was done in the present case, or 

re-run at a slightly different Mach number.

The wave drag program provides the inputs required to calculate the sonic boom due to the vehicle “volume” effects. Vehicle “lift” can also play a significant role in the prediction of the sonic 

boom signature, depending upon the vehicle weight and operating conditions. Determination of 

the boom due to lift requires knowledge of the load distribution on all the vehicle lifting surfaces 

for the specific flight conditions being investigated.

DESCRIPTION OF ELECTRONIC FILE CONTAINING AIRCRAFT GEOMETRY AND 

WAVE-DRAG DEFINITIONS OF THE CONVAIR B-58 AND XB-70-1 AIRPLANES

Both geometries were originally formatted for, and tested in, the arbitrary geometry wave drag 

program of reference 11. All results presented in this report were obtained using this format. For 

compatibility with users of the Harris wave drag program, the files are also given in Hess format.

A total of three geometries (B-58, and XB-70-1 with wing tips at 0o and 65o down) are described 

in the six electronic files contained in the compact disc (CD) provided to the NASA LaRC. File 

names for geometries given in the arbitrary wave drag format are of the form xxxxgeo.arb, and 

file names for geometries given in Hess format are of the form xxxxgeo.hes. A Portable Document Format (PDF) file of the present report is also included in the compact disc.

SUMMARY REMARKS

A detailed geometric description, in wave drag format, has been developed for the Convair B-58 

and North American XB-70-1 delta-wing airplanes. These descriptions have been placed on electronic files at the NASA Langley Research Center. The contents of the files are described in the 

present paper and are intended for use in wave drag and sonic boom calculations.

Included with the electronic files, a PDF file of the present report was also made available. The 

file contains photographs and three-view drawings of the two airplanes, tabulated geometric 

descriptions of each vehicle and its components, and comparisons of the electronic file outputs 

with existing data. The comparison includes a pictorial of the two airplanes based on the present 

geometric descriptions on the electronic files, and a comparison of the cross-sectional area distributions for both the normal Mach cuts and oblique Mach cuts above and below the vehicles. Good 

correlation exists between the area distributions generated in the late 1950s and 1960s and the 

present files.

The availability of the present electronic files allows for further validation of existing sonic boom 

prediction codes through the use of two existing experimental data bases on these two airplanes. 

The data bases were acquired in the early and mid 1960s time period and, to date, have not been 
11

fully exploited. These two data bases consist of in-flight measurements of the supersonic flowfields above and below the B-58 and XB-70-1 airplanes, acquired in 1963 and 1966 respectively, 

at distances of from about 10 to 95 body lengths.

REFERENCES

1. 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.

2. Maglieri, Domenic J.; Huckel, V.; Henderson, H. R.; and Pitman, 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.

3. Maglieri, Domenic J.; Tinetti, Ana F.; and Henderson, Herbert R.: Measured Sonic Boom 

Signatures Above and Below the XB-70 Airplane Flying at Mach Number 1.5 and 37,000 

Feet. NASA/CR-2011-217077, 2011.

4. Swihart, John M.: Transonic Aerodynamic Characteristics of a 1/15-Scale Model of the 

Convair B-58 Airplane. NACA RM SL56J22, Oct. 1956. 

5. Hopko, Russell N.; and Kinard, William H.: Drag at Model Trim Lift of a 1/15-Scale Convair B-58 Supersonic Bomber. NACA RM SL56G23, July 1956.

6. Driver, Cornelius: Wind-Tunnel Investigation of the Interference Effects During Separation 

of a 1/40-Scale Model Convair B-58 Airplane and Store at Mach Number of 1.41, 1.61, and 

2.0. NACA RM SL56L14, Dec. 1956.

7. Morris, Owen G.; and Turner, Kenneth L.: Wind-Tunnel Investigation of Mutual Interference Loads on a Supersonic Bomber Configuration and Store During Separation of Mach 

Numbers of 1.57, 1.77, and 2.01. NACA RM SL57J16a, Oct. 1957.

8. 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.

9. North American Aviation, Inc., Los Angeles, CA: Aerodynamic Dimensional Data for the 

B-70 Primary Air Vehicle B-70 Weapon System. NA-58-435,(56 pages), January 1959.

10. Pederson, Violet D.: Aerodynamics Dimensional Data for the XB-70 Air Vehicle. NA-61-

705 (Contract AF33(600)-42058), North American Aviation, Inc., Sept. 1961.

11. Craidon, Charlotte B.: User’s Guide for a Computer Program for Calculating the Zero-Lift 

Wave Drag of Complex Aircraft Configurations. NASA TM-85670, 1983.

12. Fenbert, James W.: Viewer User’s Guide. NASA Langley Research Center, 1993.

13. Hubbard, Harvey H.; Maglieri, Domenic J.; Huckel, Vera; and Hilton, David A.: Ground 

Measurements of Sonic Boom Pressures for the Altitude Range of 10,000 to 75,000 Feet. 

NASA TR R-198, 1964. (Supersedes NASA TM X-633.)
12

14. Carlson, Harry W.; and Morris, Odell A.: Wind Tunnel Investigation of the Sonic Boom 

Characteristics of a Large Supersonic Bomber Configuration. NASA TM X-898, October 

1963.
13

 TABLE I. - GEOMETRIC CHARACTERISTICS OF THE CONVAIR B-58 AIRPLANE

(Scaled from Table I of ref. 7)

[All wing dimensions defining spanwise locations or chord lengths are true dimensions

 in the chord plane unless otherwise specified. Station numbers are in feet]

Wing:

Span, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56.8

Total area, sq ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1542

Exposed area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1317

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.096

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Airfoil section parallel to root chord:

Root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACA 0003-46

Outboard of span station 0.565b/2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACA 0004-08

Camber . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0286b/2 offset tangent at 0.85b/2

Leading-edge sweepback, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

Trailing-edge sweepback, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .-10

Incidence, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

Dihedral, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.23

Tip-chord length, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Root-chord length, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54.3

Distance above parting plane at root chord:

Leading edge, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.0

Trailing edge, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.6

Hinge line, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.6

Airplane station of root chord at:

Leading edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29.4

Trailing edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83.6

Hinge line . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75.5

25 percent c . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54.8

37.5 percent c . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59.4

Length of c, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36.2

Span station of c, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.5

Elevon:

Hinge line at airplane station, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75.9

Inboard end of elevon at span station . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.7

Outboard end of elevon at span station . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19.7

Area of one elevon, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88.9

Fuselage:

Overall length, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89.6

Overall length from nose to tip of vertical tail, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 96.8

Maximum height, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.5
14

Maximum width, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.3

Maximum cross-sectional area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28.1

Vertical tail:

Span, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.5

Total area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160.0

Exposed area, sq ft 156.8

Area of control surface (rudder), sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40.0

Leading-edge sweepback, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52

Trailing-edge sweepback, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26.71

Hinge line sweepback, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35.51

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.628

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.324

Tip-chord length, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.4

Root-chord length, ft 16.7

Airplane station of root chord at leading edge, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77.9

Distance of root chord above parting plane, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2

Mean aerodynamic chord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.0

Fuselage station at leading edge of mean aerodynamic chord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . 85.6

Distance of mean aerodynamic chord above parting plane, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.2

Airfoil section parallel to root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACA 0005-64

Nacelle:

Overall length, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24.2

Maximum height above thrust plane. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.9

Maximum depth below thrust plane, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.5

Maximum width, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.7

Nacelle lip radius, in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.08

Duct inlet area including spike area (1 duct), sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.03

Duct area at exit (1 duct), sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.73

Spike apex angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50

Location of inboard nacelle:

Longitudinal location of nacelle inlet at thrust center line:

Airplane station . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32.7

Distance from wing chord plane to thrust center line:

Nacelle station 0 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.4

Nacelle station 17.7 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.8

Nacelle station 24.2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.5

Wing span station of nacelle center line . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.2

Angle between wing chord plane and nacelle center line, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .-2

Leading-edge angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.03

Trailing-edge angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.62

TABLE I.- Continued.
15

Location of outboard nacelle:

Longitudinal location of nacelle inlet at thrust center line:

Airplane station . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 558

Distance from wing chord plane to thrust center line:

Nacelle station 0, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.6

Nacelle station 11.1 ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.9

Nacelle station, 24.2 ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.0

Wing span station of nacelle center line, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.7

Angle between wing chord plane and nacelle center line, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .-4

Pylon:

Leading-edge sweepback, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

Trailing-edge sweepback, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75

Main landing gear fairings:

Span station of fairing center line . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.7

Maximum width, upper fairing, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.9

Maximum width, lower fairing, ft 4.1

Maximum height above chord plane, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.0

Maximum depth below chord plane, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.2

Store

Overall length (from pod nose), ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50.9

Overall length (from pod station 0) ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 57.6

Maximum diameter, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.0

Maximum cross-sectional area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19.6

Pod nose at airplane station . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18.3

Distance from parting plane to pod center line, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.7

Angle between pod center line and parting plane, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Base area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.0

Fineness ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.392

Store fins:

Span, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.8

Area per fin, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22.5

Exposed area per fin, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.7

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.734

Leading-edge sweepback, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.111

Trailing-edge sweep forward, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.42

Length of fin mean aerodynamic chord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.2

Pod station at leading edge of root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41.8

Pod station at trailing edge of root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51.0

Airfoil section parallel to root chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NACA 0005-64

TABLE I.- Concluded.
16

TABLE II. - GEOMETRIC CHARACTERISTICS OF XB-70-1 AIRPLANE

(from reference 8)

Total wing 

Total area (includes 230.62 m2

 (2482.34 ft2) covered by fuselage but not 

3.12 m2

 (33.53 ft2) of the wing ramp area), m2 (ft2). . . . . . . . . . . . . . . . . . . . . . . . . . . 585.07 (6297.8)

Span, m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32 (105)

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.751

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.019

Dihedral angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Root chord (wing station 0), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35.89 (117.76)

Tip chord (wing station 16m (630 in.)), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.67 (2.19)

Mean aerodynamic chord (wing station 5.43 m (17.82. (ft)), m (ft) . . . . . . . . . . . . . . . . . . 23.94 (78.532)

Fuselage station of 25-percent wing mean aerodynamic chord, m (ft) . . . . . . . . . . . . . . . . . 41.18 (135.10)

Sweepback angle, deg:

Leading edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65.57

25-percent element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58.79

Trailing edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Incidence angle, deg:

Root (fuselage juncture) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Tip (fold line and outboard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -2.60

Airfoil section (modified hexagonal):

Root to wing station 4.72m (186 in.) (thickness-chord ratio, 2 percent) . . . . . . . . . . . . . . .0.30 to 0.70

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

 (thickness-chord ratio, 2.5 percent) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.30 to 0.70

Inboard wing -

Area (includes 230.62 m2 (2482.34 ft2)covered by fuselage but not

 3.12 m2

 (33.53 ft2) wing ram area, m2 (ft2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 488.28 (5256.0)

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

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

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.407

Dihedral angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Root chord (wing station 0), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35.89 (117.76)

Tip chord (wing station 9.67 m (380.62 in.)), m (ft). . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.61 (47.94)

Mean aerodynamic chord (wing station 4.15 m (163.58 in.)), m (in.) . . . . . . . . . . . . . . . . 26.75 (1053)

Fuselage station of 25-percent wing mean aerodynamic chord, m (in) . . . . . . . . . . . . 39.07 (1538.29)

Sweepback angle, deg:

Leading edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65.57

25-percent element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58.79

Trailing edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Airfoil section (modified hexagonal):

Root (thickness-chord ratio, 2 percent) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.30 to 0.70

Tip (thickness-chord ratio, 2.4 percent). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.30 to 0.70
17

Mean camber (leading edge), deg:

Butt plane 0 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.15

Butt plane 2.72 m (107 in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.40

Butt plane 3.89 m (153 in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.75

Butt plane 6.53 m (257 in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.60

Butt plane 9.32 m (367 in.) to tip . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Outboard wing -

Area (one side only), m2 (ft2

) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48.39 (520.90)

Span, m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.33 (20.78)

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.829

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.046

Dihedral angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

Root chord (wing station 9.67 m) (380.62 in.)), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . 14.61 (47.94)

Tip chord (wing station 16.00 m) (630 in.)), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.67 (2.19)

Mean aerodynamic chord (wing station 11.87 m) (467.37 in.)), m (in.) . . . . . . . . . . . . 9.76 (384.25)

Sweepback angle, deg:

Leading edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65.57

25-percent element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58.79

Trailing edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Airfoil section (modified hexagonal):

Root (thickness-chord ratio, 2.4 percent) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.30 to 0.70

Tip (thickness-chord ratio, 2.5 percent). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.30 to 0.70

Down deflection from wing reference plane, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0,25,65

Skewline of tip fold, deg:

Leading edge in . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5

Leading edge down . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

Wing tips

Up Down

Elevons (data for one side):

Total area aft of hinge line, m2

 (ft2) 18.37 (197.7) . . . . . . . . . 12.57 (135.26)

Span, m (ft) 6.23 (20.44) . . . . . . . . . . . 4.26 (13.98)

Inboard chord (equivalent), m (in.) 295 (116). . . . . . . . . . . . . . . . (116) 2.95

Sweepback angle of hinge line, deg 0 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Deflection, deg:

As elevator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -25 to 15

As aileron with elevators at +15ο οr less . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .−15 tο 15

As aileron with elevators at -25o or less . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -5 to 5

Total . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -30 to 30

Canard -

Area (includes 13.96 m2

 (150.31 ft2) covered by fuselage), m2 (ft2) . . . . . . . . . . . . . . . 38.61 (415.59)

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

TABLE II.- Continued.
18

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.997

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.388

Dihedral angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

Root chord (canard station 0), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.34 (20.79)

Tip chord (canard station 4.39 m (172.86 in.)), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.46 (8.06)

Mean aerodynamic chord (canard station 1.87 m (73.71 in.)), m (in.) . . . . . . . . . . . . . 4.68 (184.3)

Fuselage station of 25-percent canard mean

aerodynamic chord, m (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.06 (553.73)

Sweepback angle, deg:

leading edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31.70

25-percent element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.64

trailing edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -14.91

Incidence angle (nose up), deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0 to 6

Airfoil section (modified hexagonal):

root (thickness-chord ratio 2.5 percent) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.34 to 0.66

tip (thickness-chord ratio 2.52 percent) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.34 to 0.66

Ratio of canard area to wing area 0.066

Canard flap (one of two):

Area (aft of hinge line), m2

 (ft2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.08 (54.69)

Ratio of flap area to canard semiarea . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.263

Vertical tail (one of two) -

Area (includes 0.83 m2 (8.96 ft2) blanketed area),

m2 (ft2

) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.74 (233.96)

Span, m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.75 (15)

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.30

Root chord (vertical-tail station 0), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.03 (23.08)

Tip chord (vertical-tail station 4.57 m

(180 in.)), m (ft) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211. (6.92)

Mean aerodynamic chord (vertical-tail station 1.88 m

(73.85 in.)), m (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.01 (197.40)

Fuselage station of 25-percent vertical-tail mean

aerodynamic chord, m (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55.59 (2188.50)

Sweepback, angle, deg:

Leading edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51.77

25-percent element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45

Trailing edge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.89

Airfoil section (modified hexagonal):

Root (thickness-chord ratio 3.75 percent) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.30 to 0.70

Tip (thickness-chord ratio 2.5 percent) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.30 to 0.70

Cant angle, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0

TABLE II.- Continued.
19

Ratio of vertical tail to wing area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.037

Rudder travel, deg:

With gear extended . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±12

With gear retracted . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ±3

Fuselage (includes canopy) -

Length, m (ft) 56.62 (185.75)

Maximum depth (fuselage station 22.30 m

(878 in.)), M (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.72 (106.92)

Maximum breadth (fuselage station 21.72 m

(855 in.)), m (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.54 (100)

Side area, m2 (ft2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87.30 (939.72)

Planform area, m2 (ft2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 110.07 (1184.78)

Center of gravity:

Forward limit, percent mean aerodynamic chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19.0

Aft limit, percent mean aerodynamic chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25.0

Duct -

Length, m (ft) 31.96 (104.84)

Maximum depth (fuselage station 34.93 m

(1375 in.)), m (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.31 (90.75)

Maximum breadth (fuselage station 53.34 m

(2100 in.)), m (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9.16 (360.70)

Side area, m2 (ft2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 66.58 (716.66)

Planform area, m2 (ft2

) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217.61 (2342.33)

Inlet captive area (each), m2 (in2) 3.61 (5600) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.61 (5600)

Surface areas (net wetted), m2 (ft2) -

Fuselage, canopy, boundary layer gutter, and tailpipes . . . . . . . . . . . . . . . . . . . . . . . 264.77 (2850.0)

Ducts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 318.71 (3430.6)

Wing, wing tips, and wing ramp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 864.71 (9307.7)

Vertical tails (two) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87.12 (937.7)

Canard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49.47 (532.5)

Total . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1584.79 (17,058.5)

Engines (six) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . YJ93-GE-3

Boattail angle, deg -

Upper surface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

Lower surface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

Side . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

Base areas, m2 (ft2) -

Total . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.7 (137)

Total (all engines on, minimum exit area) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 (107.2)

Total (all engines on, maximum exit area) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.5 (48.5)

Projected thickness (height) of base, m (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.47 (58)

TABLE II.- Continued.
20

Width of propulsion package, cm (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 914 (360)

Engine -

Jet-exit area (minimum), cm2 (in2

) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4613 (715)

Jet-exit area (maximum), cm2 (in2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13,678 (2120)

Jet-exit diameter (minimum), cm (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77 (30.2)

Jet-exit diameter (maximum), cm (in.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132 (52) 

TABLE II.- Concluded.
21

(Courtesy of U.S. Air Force)

(Courtesy of NASA Flight Research Center)

(a) Convair B-58

(b) North American XB-70-1

Figure 1.- Photographs of delta-wing airplanes to be geometrically described.
22

Figure 2.- Three-view drawings of delta-wing airplanes to be geometrically described.

189’

105’

30.0’

(a) Convair B-58 (total wing area = 1542 sq. ft.)

(b) North American XB-70-1 (total wing area = 6297.8 sq. ft.)
23

Figure 3.- Pictorials of B-58 and XB-70 as generated from current electronic files containing 

geometric descriptions of both airplanes.

(b) North American XB-70

(a) Convair B-58
24

distance from model nose, in.

Area, in.

2

0.00 5.00 10.00 15.00 20.00 25.00 30.00

0.00

1.00

2.00

3.00

4.00

5.00

6.00

7.00

8.00

9.00

distance from model nose, in.

Area, in.

2

0.00 5.00 10.00 15.00 20.00 25.00 30.00

0.00

1.00

2.00

3.00

4.00

5.00

6.00

7.00

8.00

9.00

(a) As generated in 1956 (ref. 6)

(b) As generated using present vehicle geometric description

Figure 4.- Comparison of normal cross-sectional area distributions of B-58 wind tunnel model 

and components (inlet capture area not included).

fuselage

wing

tail

inboard

nacelles

outboard

nacelles

m.l.g. fairing pod

fuselage

wing

pod

tail

inboard

nacelles

total

outboard

nacelles

i.n. pylon m.l.g. fairing
25

x/l

A/l

2

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

0.000

0.004

0.008

0.012

no stores

with stores 

x/l

A/l

2

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

0.000

0.004

0.008

0.012

no stores

with stores 

(a) As generated in 1961 (ref. 13)

(b) As generated using present vehicle geometric description

Figure 5.- Comparison of normal cross-sectional area distributions of B-58 with and without 

MB-1 fuselage pod (inlet capture area not included).

no pod

with pod

no pod

with pod
26

x/l

A/l

2

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

0.000

0.005

0.010

0.015

from reference 1 

present report

x/l

A/l

2

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

0.000

0.005

0.010

0.015

from reference 1 

present report

(a) Area distribution based on oblique cuts for positions above the airplane (θ = 90o)

(b) Area distribution based on oblique cuts for positions below the airplane (θ = -90o)

Figure 6.- Comparison of total area distributions above and below B-58 airplane with MB-1 pod

at M=1.65 (inlet capture area not included).

Direction of

area cut

Direction of

area cut
27

fuselage station, in.

cross-sectional area, ft

2

0 500 1000 1500 2000 2500

0

50

100

150

200

250

300

from reference 10 

present report

x/l

A/l

2

0.00 0.20 0.40 0.60 0.80 1.00

0.000

0.001

0.002

0.003

0.004

0.005

0.006

0.007

from reference 14 

present report 

Figure 7.- Comparison of non-dimensional total normal cross-sectional area distributions of XB-70-1 

airplane. Inlet capture area not included. Wing tips at 65o

Figure 8.- Comparison of total normal cross-sectional area distributions of XB-70-1 airplane.

Inlet capture area included. Wing tips at 0o
28

fuselage station, inches

cross-sectional area, ft

2

0 500 1000 1500 2000 2500

0

50

100

150

200

250

300

fuselage station, inches

cross-sectional area, ft

2

0 500 1000 1500 2000 2500

0

50

100

150

200

250

300

(a) As generated in 1961 (ref. 10). Inlet capture area included.

(b) As generated using present vehicle geometric description. Inlet capture area not included.

Figure 9.- Comparison of normal cross-sectional area distributions of XB-70-1 airplane and 

components. Wing tips at 0o

canard

total ducts

wing and wing ramp

fuselage and canopy

total

verticals lower protruberance

total

canard

fuselage and canopy

lower protruberance verticals

wing

total ducts
29

fuselage effective length, in.

cross-sectional area, ft

2

0 500 1000 1500 2000 2500

0

50

100

150

200

250

300

fuselage effective length, in.

cross-sectional area, ft

2

0 500 1000 1500 2000 2500

0

50

100

150

200

250

300

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

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

Figure 10.- 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 at 65o

 down).

θ = 90ο

θ = 25ο

θ = 0ο

θ = 90ο

θ = −90ο

θ = 0ο

θ = −90ο
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Contractor Report

 4. TITLE AND SUBTITLE

Equivalent Longitudinal Area Distributions of the B-58 and XB-70-1 

Airplanes for use in Wave Drag and Sonic Boom Calculations 

5a. CONTRACT NUMBER

NNL05AB74T

 6. AUTHOR(S)

Tinetti, Ana F.; Maglieri, Domenic J.; Driver, Cornelius; Bobbitt, Percy J.

 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

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NASA

13. SUPPLEMENTARY NOTES

Langley Technical Monitor: Peter G. Coen

12. DISTRIBUTION/AVAILABILITY STATEMENT

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

A detailed geometric description, in wave drag format, has been developed for the Convair B-58 and North American XB-70-1 delta wing 

airplanes. These descriptions have been placed on electronic files, the contents of which are described in this paper

They are intended for use in wave drag and sonic boom calculations. Included in the electronic file and in the present paper are photographs 

and 3-view drawings of the two airplanes, tabulated geometric descriptions of each vehicle and its components, and comparisons of the 

electronic file outputs with existing data. The comparisons include a pictorial of the two airplanes based on the present geometric 

descriptions, and cross-sectional area distributions for both the normal Mach cuts and oblique Mach cuts above and below the vehicles. 

Good correlation exists between the area distributions generated in the late 1950s and 1960s and the present files. The availability of these 

electronic files facilitates further validation of sonic boom prediction codes through the use of two existing data bases on these airplanes, 

which were acquired in the 1960s and have not been fully exploited.

15. SUBJECT TERMS

Sonic boom,; Equivalent area distributions; Wave drag; Aircraft geometry

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01- 03 - 2011