ADA387404 Effects of Nuclear Weapons Tests on Aircraft in Flight Upshot Knothole

•Hi:-

■I

L. i

20010312 112

UPSHOT-KNOTHOLE

T£ST" WT-748 /

Copy No. 158 A

•i- j

NEVADA PROVING GROUNDS

March-June 1953 DISTRIBUTION STATEMENT A APPilES-PER NTPR REVIEW.

DATE //^/-^gga /

AGENCY TECHNICAL LIBRARY

Project 54 ATOMIC WEAPON EFFECTS ON AD TYPE AIRCRAFT IN FLIGHT to)

GR°U? V.-W?.*

ThiB docuMôntains rest defin^PêAMmic E Its^flnsmittal oRwApi&losür£ coflnts in any manner to an unauthorized

DISTRIBUTION STATEMMr ÄP"*ibited. Approved for Public Release 7: /hffCAJ/'

Distribution Unlimited *£™Zr&*** JP~\ HEADQUARTERS FIELD COMMAND, ARMEO FORCES SPECIAL WEAPONS PROJECT ,

SANDIA BASE, ALBUQUERQUE, NEW MEXICO Reproduced From

' . ■ ..■' ,a,;,:;;:.r=v.;'J Best Available Copy

; w**-z**>i

^4^^ a V/VCtAK,.

j-f—

54VJC-31009

w"t

Reproduced Direct from Manuscript Copy by AEC Technical Information Service

Oak Ridge, Tennessee

Inquiries relative to this report may be made to

Chief, Armed Forces Special Weapons Project Washington, D. C.

If this report is no longer needed, return to

AEC Technical Information Service P. O. Box 401

Oak Ridge, Tennessee

Defense Threat Reduction Agency 8725 John J Kingman Road MS 6201

Ft Belvoir, VA 22060-6201

TDANP/TRC March 2, 2001

MEMORANDUM TO THE DEFENSE TECHNICAL INFORMATION CENTER ATTN: OCQ

SUBJECT: DOCUMENT UPDATES

The Defense Threat Reduction Agency Security Office has performed a classification/distribution statement review of the following documents. The documents should be changed to read as follows:

WT-1628, AD-357954, OPERATION HARDTACK, PROJECT 3.4, LOADING AND RESPONSE OF SURFACE-SHIP HULL STRUCTURES FROM UNDERWATER BURSTS, UNCLASSIFIED, DISTRIUBTION STATEMENT A.

WT-1301, AD-341065, OPERATION REDWING, PROJECT 1.1, GROUND SURFACE AIR BLAST PRESSURE VERSUS DISTANCE, UNCLASSIFIED, DISTRIBUTION STATEMENT A.

WT-748, OPERATION UPSHOT KNOTHOLE, PROJECT 5.1, ATOMIC WEAPON EFFECTS ON AD TYPE AIRCRAFT IN FLIGHT. UNCLASSIFIED, DISTRIBUTION STATEMENT A. FORWARD TO YOU FOR YOUR COLLECTION

WT-9001-SAN, GENERAL REPORT ON WEAPONS TESTS, UNCLASSIFIED, DISTRIBUTION STATEMENT A. FORWARD TO YOU FOR YOUR COLLECTION.

POR-2260-SAN, OPERATION SUN BEAM, SHOTS LITTLE FELLER 1 AND 2, PROJECT 1.1, AIRBLAST PHENOMENA FROM SMALL YIELD DEVICES, SANITIZED VERSION. UNCLASSIFIED, DISTRIBUTION STATEMENT A. FORWARD TO YOU FOR YOUR COLLECTION.

If you have any questions, please call me at 703-325-1034.

OCO CUU-X-Jw Q

ARDITH JARRETT Chief, Technical Resource Center

WT-748

This document consists of 198 pages

No. -LÔ of 260 copies, Series A

OPERATION UPSHOT-KNOTHOLE

Project 5.1

ATOMIC WEAPON EFFECTS ON AD TYPE AIRCRAFT IN FLIGHT

REPORT TO THE TEST DIRECTOR

by

Leo Rogin Alden C. DuPont Christian G. Weeber

DISTRIBUTION STATEMENT A Approved for Public Release

Distribution Unlimited

March 1954

This docum defined Its tjdmsmittal

tins restricted data as Energy

unauthorized Irson is prohibited

Naval Air Material Center Philadelphia 12, Pennsylvania

DNA-76-051 18

541VC-31009

ABSTRACT

This report presents measured, observed and calculated data associated with atomic weapon effects upon the structure of Model AD aircraft in the vicinity of an atomic explosion. Data covering weapons effects and airplane structural response to these effects are presented for aircraft in level flight attitude, tail toward the blast in a vertical plane containing the burst point. This orientation represents an escape position of an AD type aircraft following delivery of an atomic weapon.

During Operation UPSHOT-KNOTHOLE, this project participated in a total of five shots. One or the other of two Navy Model AD aircraft converted to drone configuration was flown in Shots 1, 2, J, 8, and 9. The slant ranges at burst time involved in these shots varied from 14,400 ft for the AD-2 piloted flight of Shot 1 to 6200 ft for the AD-2 pilotless flight of Shot 7. In Shot 7, the actual yield exceeded the planned yield by greater than 30 per cent. The drone aircraft was positioned for near critical weapon effects and the higher thermal radiation severely weakened all the blue painted skin on the underside of the wing. Both the port and starboard outboard wing panels were torn off at the time of shock arrival as a result of the weakened skin and combined overpressure and gust effects. A considerable amount of valuable information on thermal damage to aircraft in flight was obtained from these panels which were recovered after the test. Neither panel incurred any significant additional damage due to the free fall and subsequent ground Impact. Visual analysis of the structural failures indicated that the aircraft might have survived had the bottom skin of the wing been bare aluminum or painted heat resistant white instead of standard blue.

In addition to the above flight tests, aluminum alloy panels of various thicknesses and paint finishes were exposed at three different stations on the ground during Shot 9 to obtain supplemental information on the effects of thermal radiation. Effective thermal absorptivity coefficients obtained ranged in value from 0.12 to 0.16. The results of these tests are reported in Appendix 3.

Measured overpressures were in agreement with the theoretical values. Measured thermal radiation was seen to be appreciably greater than predicted as a direct result of ground reflectivity. Thermal calculations using ß = 0.55, (albedo) provided good correlation with test measurements. Peak aircraft accelerations as

measured were approximately double the calculated values; however, the measured wing and tail loads were in close agreement ^^/ith the loads calculated using rigid body relationships. Aircraft elasticity effects, even on this comparatively rigid airplane, were readily seen. No direct correlation between measured and calculated aircraft skin temperature rise has been established, although the effects of heat received, skin thickness, and surface finish, are indicated. The arbitrary assumption of turbulent or laminar airflow, and corresponding cooling rates, resulted in agreement with the rates as measured during time histories of skin temperature rise. Results of metallurgical studies on aircraft skin specimens, begun in an attempt to determine skin temperature rise in Shot 7, indicated effects normally associated with temperatures far in excess of those recorded. This is believed to be due to microscopic thermal concentrations in the grain structure of the material brought about by the instantaneous application of the thermal pulse. The effects are so localized that no serious structural consequences, exceeding those indicated by the thermal data presented in this report, are expected. Appendix A presents these metallurgical results. Appendix B presents data on ground panels which may prove useful in further analysis of the temperature problem.

The use of the data presented in this report for the purpose of improving analytical methods for predicting the effect of atomic weapons on Naval aircraft is recommended. Recommendations for future tests include thermal radiation measurements in flight to further evaluate ground refleetivity and measurement of time histories of aircraft skin temperature rise in flight, followed by metallographic study of the structure.

FORSWORD

This report is one of the reports presenting the results of the 78 projects participating in the Military Effects Tests Program of Operation UPSHOT-KNOTHOLE, which included 11 test detonations. For readers interested in other pertinent test information, reference is made to WT-782, Summary Report of the Technical Director. Military Effects Program. This summary report includes the following information of possible general interest.

a. An over-all description of each detonation, including yield, height of burst, ground zero location, time of detonation, ambient atmospheric conditions at detonation, etc., for the 11 shots.

b. Compilation and correlation of all project results on the basic measurements of blast and shock, thermal radiation, and nuclear radiation.

c. Compilation and correlation of the various project results on weapons effects.

d. A summary of each project, including, objectives and results.

e. A complete listing of all reports covering the Military Effects Test Program.

ACKNOWLEDGMENTS

The test program reported herein was successfully accomplished only through the combined efforts of many individuals, both military and civilian, representing many different agencies. Although individual acknowledgments cannot be made here, the following is a list of the organizations who contributed to the success of this program:

Bureau of Aeronautics of the Navy

Directorate of Weapons Effects Tests, Armed Forces Special Weapons Projects

Douglas Aircraft Company, El Segundo Division

Electronics Associates. Long Branch, N. J.

Massachusetts Institute of Technology, Dept. of Aeronautical Engineering

Naval Air Development Squadron Five

Naval Materials Laboratory

Naval Ordnance Test Station

Naval Radiological Defense Laboratory

Signal Corps Engineering Laboratory

Special Weapons Center of the Air Force

Wright Air Development Center

$ f ?'*■ ü-\ f ;.J"':*n

CONTENTS

ABSTRACT 3 FOREWORD • 5 ACKNOWLEDGMENTS '

ILLUSTRATIONS 11

TABLES l6

CHAPTER 1 INTRODUCTION , 17

1.1 Objective 17

1.1.1 Purpose of Project 17 1.1.2 Need for Project 17 1.1.3 Report Objective 17

1.2 Experiment Design 18

1.2.1 Background 18 1.2.2 Instrumentation 18 1.2.3 Analytical Methods 21 1.2.4 Opsrational Procedures 24-

CHAPTER 2 RESULTS AND OBSERVATIONS 25

2.1 Results 25 2.2 Data and Observations 25

2.2.1 Shot 1 25 2.2.2 Shot 2 28 2.2.3 Shot 7 31 2.2.4 Shot 8 36 2.2.5 Shot 9 41

CHAPTER 3 DISCUSSION 75

3.1 General 75 3.2 Thermal Effects 75

9

3.2.1 General 75 3.2.2 Thermal Radiation 76 3.2.3 Aircraft Skin Temperature Rise ... 76

3.3 Nuclear Radiation 80

3.3.1 Gamma Radiation SO 3.3.2 Neutron Radiation 81

3.4 Overpressure and Gust Effects 81

3.4.1 Overpressure 81 3.4.2 Gust Effects 81

CHAPTER 4 CONCLUSIONS AND RECOMMENDATIONS 120

4.1 Conclusions 120

4.1.1 Thermal Effects 120 4.1.2 Overpressure and Gust Effects . . . 121 4.1.3 General 1Z1

4.2 Recommendations 121

4.2.1 Thermal Effects 121 4.2.2 Overpressure and Gust Effects . . . 122

SYMBOLS AND DEFINITIONS i23

APPENDIX A Metallurgical Tests of Skin Specimens Taken frcm AD Type Aircraft Exposed in Flight to Thermal Radiation from an Atomic Explosion 125

A.l OBJECTIVES 126

A.2 METHOD 126

A.2.1 Strength Tests 126 A.2.2 Metallographic Examinations 126

A.3 RESULTS 127

A.3.1 Strength Tests 127 A.3.2 Metallographic Examinations 127

APPENDIX B Effects of Thermal Radiation on Thin Aluminum^ Alloy Panels Exposed on the Ground to an Atomic Explosion 136

B.l OBJECTIVE 137

10

B.2 BACKGROUND 137

B.3 TEST DESCRIPTION 138

B.4 INSTRUMENTATION 139

B.5 RESULTS 14.0

B.6 CONCLUSIONS HI

B.7 RECOMMENDATIONS LU

BIBLIOGRAPHY 195

ILLUSTRATIONS

1.1 Model AD Type Drone Aircraft (AD-2) 19 1.2 Bottom View of AD Drone Showing Temperature and

Gamma Measurement Locations 20

2.1 AD-2, Starboard Aileron Showing Scorched Blue Paint on 0.016 Skin - Shot 2 , 45

2.2 AD-2, Starboard Aileron and Wing Tip Showing Scorched Blue Paint on 0.016 Skin - Shot 2 46

2.3 AD-2, Port Elevator and Horizontal Stabilizer Showing Scorched Blue Paint on 0.016 Skin - Shot 2 ... 47

2.4 AD-2, Port Wing Flap Showing Scorched Blue Paint on 0.020 Skin and on 0.025 Skin Forming Curved Portion Aft of Rear Shear Web Facing the Burst Shot 2 48

2.5 AD-2, Port Wing Flap Showing Scorched Blue Paint on 0.025 Skin Forming Curved Portion Aft of Rear Shear Web Facing the Burst - Shot 2 49

2.6 AD-2, Port Wing Showing Scorched Blue Paint on 0.025 Skin on Hinged Aft Edge of Wing Panel - Shot 2 50

2.7 AD-2, Port Wing Panel and Tip Showing Scorched Blue Paint on 0.032 Skin - Shot 2 51

2.8 AD-2, Aperture Between Elevator and Horizontal Stabilizer Showing Destroyed Rubberized Fabric Aerodynamic Seal - Shot 2 52

2.9 AD-2, Aft of Rear Spar of Horizontal Stabilizer Showing Burned Fabric Lightening Hole Covers - Shot 2 . . 53

2.10 AD-2, Center of Starboard Aileron Showing Burned Out 0.016 and 0.025 Panels - Shot 7 . . 54

2.11 AD-2, Port Aileron Showing Aileron Seal Intact - Shot 7 55

2.12 AD-2, Port Aileron - Showing Relative Damage to Various Combinations of Surface Finish and Skin Thickness - Shot 7 56

11

2.13 AD-2, Starboard Aileron - Showing Thermal Damage to Aluminized Lacquered 0.025 Skin Compared with Stand- ard Blue Painted 0.025 Skin - Shot 7 57

2.14 AD-2, Port Aileron - Showing Thermal Damage to Aluminized Lacquered 0.025 Skin Compared with Stand- ard Blue Painted 0.025 Skin - Shot 7 58

2.15 AD-2, Port Aileron - Showing Relative Damage to Various Combinations of Surface Finish and Skin Thickness - Shot 7 ' 59

2.16 AD-2, Outboard Port Wing Panel and Wing Tip Showing Wing Tip Intact. Note how White Star Prote.: r.ed 0.032 Skin - Shot 7 60

2.17 AD-2, Outboard Port Wing Panel Showing Heat Resistant White Painted 0.032 Plate - Shot 7 61

2.18 AD-2, Outboard Port Wing Panel Showing Aluminized Lac- quered 0.040 Plate - Shot 7 62

2.19 AD-2, Tip of Elevator and Horizontal Stabilizer Show- ing Fabric Lightening Hole Covers and Elevator Tip Intact - Shot 7 63

2.20 AD-2, Topside of Port Wing Panel Showing Scorched Blue Paint on 0.016 Skin - Shot 7 64

2.21 XBT2D-1, Starboard Aileron Showing Scorched Aluminized Lacquer on 0.016 and 0.025 Skin - Shot 8 65

2.22 XBT2D-1, Port Aileron Showing Scorched Standard White Paint on 0.016, 0.025 and 0.040 Skin - Shot 8 66

2.23 X3T2D-1, Port Elevator Showing Scorched Painted Sur- faces. Note Skin Ripples on 0.016 Standard Blue Paint- ed and Aluminized Lacquered Surfaces - Shot 8 67

2.24 XBT2D-1, Starboard Wing Flap with Thermal Test Panels Installed, Outboard Portion - Shot 8 68

2.25 XBT2D-1, Starboard Wing Flap with Thermal Test Panels Installed, Inboard Portion - Shot 8 69

2.26 XBT2D-1, Port Wing Flap with Thermal Test Panels In- stalled, Inboard Portion - Shot 8 7°

2.27 XBT2D-1, Port Wing Flap with Thermal Test Panels In- stalled, Outboard Portion- Shot 8 71

2.28 XBT2D-1, Center of Wing Showing Scorched Blue Paint on 0.032 Skin - Shot 8 ?2

2.29 XBT2D-1, Standard Aileron Wing Tip Showing Scorched Standard Blue Paint on 0.016 and 0.040 Skin. Note Ripples on Standard Blue Painted 0.016 Skin - Shot 8 . . 73

2.30 XBT2D-1, Starboard Stabilizer and Elevator Showing Scorched Blue Paint on 0.032 and 0.040 Skin - Shot 8 . . 74

3.1 Aircraft and Calorimeter Orientation Relative to Burst . 77 3.2 Time History of Measured Thermal Radiation 78 3.3 Reflected Thermal Intensity - Tower Shot 82 3.4 Reflected Thermal Unit Source - Air Drop 83 3.5 Comparison of Measured and Calculated Thermal Radiation. 84 3.6 Measured and Calculated Thermal Radiation Reduced to

1 KT 35 3.7 Time History of Measured Aircraft Skin Temperature Rise. 36

12

3.8 Aircraft Skin Temperature Time History for Measured and Calculated Cooling Rates 87

^.9 Incremental Temperatures Measured in Aircraft Skin with Heat Resistant White Paint Finish - Shot 8 88

3.10 Incremental Temperatures Measured in Standard White Painted Aircraft Skin - Shot 8 89

3.11 Incremental Temperatures Measured in Unpainted Air- craft Skin - Shot 8 90

3.12 Incremental Temperatures Measured in Aircraft Skin with an Aluminized Lacquer Finish - Shot 8 91

J>.lj Incremental Temperatures Measured in Standard Blue Painted Aircraft Skin - Shot 8 92

3.14 Temperature Rise in Aircraft Skin VS the Reciprocal of Skin Thickness for Different Surface Finish - Shot 8 93

3.15 Temperature Rise in Aircraft Skin VS the Reciprocal of Skin Thickness for Different Surface Finish - Shot 2 94-

3.16 Comparison of Equivalent Absorptivity for 0.016 and 0.064 Aircraft Skin with Different Surface Finish ... 95

3.17 Comparison of Measured and Calculated Gamma Radiation, Indicating Structural Shielding Effects 96

3.18 Measured and Calculated Gamma Radiation Reduced to 1 KT in Sea Level Homogeneous Atmosphere 97

3.19 Triple Point Trajectory Curves Indicating Aircraft Position at Shock Arrival 98

3.20 Measured and Calculated Shock Arrival Times VS Slant Range 99

3.21 Comparison of Measured and Calculated Overpressure . . . 100 3.22 Measured and Calculated Overpressure Reduced to 1 KT

in a Sea Level Homogeneous Atmosphere 101 3.23 Pressure Ratio VS Density Ratio 102 3.24 Gust Velocity as a Function of Pressure Ratio and

Altitude 103 3.25 Time History of Tail Acceleration and Tail Loads -

Shot 8 104. 3.26 Time History of e.g. Acceleration, Wing Loads and

Overpressure - Shot 8 105 ^.27 Measured Accelerations and Tail Loads - Shot 8 106 3.28 Measured Overpressure and Wing Loads - Shot 8 107 3.29 Time History of e.g. Acceleration, Tail Acceleration,

Tail Loads and Overpressure - Shot 2 108 3.30 Time History of Wing Loads Indicating Shot 2

Estimated Peaks Based on Wing Dynamic Response 109 3.31 Time History of Blast Data - Shot 1 110 3.32 Time History of Blast Data - Shot 9a (First Shock,

Shot 9) Ill J.^3 Time History of Blast Data - Shot 9c (Third Shock,

Shot 9) 112

3.34 Comparison of Measured and Calculated Aircraft Normal Acceleration 113

13

3.35 Comparison of Measured and Calculated Horizontal Stabilizer Shear i:L4

3.36 Comparison of Measured and Calculated Horizontal Stabilizer Bending 115

3.37 Comparison of Measured and Calculated Wing Shear .... Ho 3.38 Comparison of Measured and Calculated Wing Bending,

W.S. 57 117 3.39 Comparison of Measured and Calculated Wing Bending,

W.S. 107 and W.S. 162 11° 3.40 Time History of Airplane Pitching Motion iiy

A.l Photomicrographs Showing Unaffected and Slightly Over- heated Aluminum Alloy Sheets 134

A.2 Photomicrographs Showing the Effects of Overheating 24S Aluminum Alloy Sheets 135

B.l Panel Construction V+z B.2 Frame Construction }fy B.3 Arrangement of Panels in Frames 144 B.4 Tensile Test Specimen ^45 B.5 Thermal Energy Indicator Strip Before Exposure 145 B.6 Thermal Energy Indicators After Exposure 146 B.7 Frame with Mounted Panels and Thermal Energy Indicators. 147 B.8 Mounted Panel Showing Temp Tape Installation 14-8 B.9 Frame with Temporary Protective Covering 149 B.10 Frame at Test Site Before Exposure 149 B.ll Temp Tape 150 B.12 Frame Number 1 After Exposure - Slant Range 7150 ft,

Theimal Energy 18.3 cal/cm2 150 B.13 Frame Number 2 After Exposure - Slant Range 5330 ft,

Thermal Energy 30 cal/cnr 151 B.14 Frame Number 3 After Exposure - Slant Range 3800 ft,

Thermal Energy 43 cal/cm2 152 B.15 Panel 1 After Exposure - 0.012 In. Thick - Bare Metal -

Station 1 153 B.16 Panel 2 After Exposure - 0.012 In. Thick - Aluminized

Lacquer - Station 1 153 B.17 Panel 3 After Exposure - 0.012 In. Thick - White Lacquer

Station 1 154 B.18 Panel 8 lifter Exposure - 0.016 In. Thick - Sea Blue

Lacquer - Station 1 154 B.19 Panel 10 After Exposure - 0.016 In. Thick - Sea Blue

Lacquer - Station 1 155 B.20 Panel 11 After Exposure - 0.020 In. Thick - Bare

Metal - Station 1 155 B.21 Panel 13 After Exposure - 0.020 In. Thick - Alumi-

nized Lacquer - Station 1 156

B.22 Panel 15 After Exposure - 0.020 In. Thick - White ^ Lacquer - Station 1

B.23 Panel 17 After Exposure - 0.020 In. Thick - Sea Blue „ Lacquer - Station 1

14

B.24 Panel 23 After Exposure - 0.025 In. Thick - Sea Blue Lacquer - Station 1 157


B.26 Panel 34- After Exposure - 0.051 In. Thick - Sea Blue Lacquer - Station 1 158

B.27 Panel 5 - After Exposure - 0.016 In. Thick - Bare Metal - Station 2 159

B.28 Panel 6 After Exposure - 0.016 In. Thick - Aluminized Lacquer - Station 2 159

B.29 Panel 7 After Exposure - 0.016 In. Thick - White Lacquer - Station 2 160

3.30 Panel 9 After Exposure - 0.016 In. Thick - Sea Blue Lacquer - Station 2 160


B.32 Panel 20 After Exposure - 0.025 In. Thick - Bare Metal - Station 2 161

















B.49 Panel 36 After Exposure - 0.051 In. Thick - Sea Blue' Lacquer - Station 3 170

15

B.50 B.51

B.52 B.53 B.54 B.55 B.56 B.57 B.58 B.59 B.60 B.61 B.62 B.63 B.64 B.65 B.66 B.67 B.68

B.69

B.70

B.71

B.72

1.1

2.1 2.2

3.1

A.l

A.2

B.l B.2 B.3 B.4

B.5 B.6 B.7

Orientation of Elevator Sections for Test Exposure . . 171 Elevator Sections Showing Locations of Test Specimens and Temp Tapes 172 Elevator Section 1 - Exposed Side - Station 2 173 Elevator Section 1 - Unexposed Side - Station 2 . . . . 173 Elevator Section 2 - Exposed Side - Station 2 174. Elevator Section 2 - Unexposed Side - Station 2 . . . . 174. Elevator Section 3 - Exposed Side - Station 1 175 Elevator Section 3 - Unexposed Side - Station 1 . . . . 175 Elevator Section 4 - Exposed Side - Station 1 176 Elevator Section 4 - Unexposed Side - Station 1 . . . . 176 Elevator Section 5 - Exposed Side - Station 2 177 Elevator Section 5 - Unexposed Side - Station 2 . . . . 177 Elevator Section 6 - Exposed Side - Station 2 178 Elevator Section 6 - Unexposed Side - Station 2 . . . . 173 Elevator Section 7 - Exposed Side - Station 1 179 Elevator Section 7 - Unexposed Side - Station 1 . . . . 179 Elevator Section 8 - Exposed Side - Station 1 180 Elevator Section 8 - Unexposed Side - Station 1 . . . . 130 Equivalent Thermal Absorptivity Coefficient VS Peak Temperature Attained - Sea Blue Lacquer 181 Equivalent Thermal Absorptivity Coefficient VS Peak Temperature Attained - White Lacquer 182 Equivalent Thermal Absorptivity Coefficient VS Peak Temperature Attained - Aluminized Lacquer 183 Equivalent Thermal Absorptivity Coefficient VS Peak Temperature Attained - Bare Metal 184 Equivalent Thermal Absorptivity Coefficient VS Panel Thickness - Sea Blue Lacquer 185

TABLES

Calculated Wing and Horizontal Stabilizer Load Constants ^

Thermal Damage to Aircraft Skin - Shot 7 34 Summary of Primary Data for All Shots 43

Calculated Absorptivity Values - Shot 8 80

Summary of Metallurgical Test Results on Skin Specimens from Model XBT2D-1 Airplane 123 Summary of Metallurgical Test Results on Skin Specimens from Model AD-2 Drone Airplane 131

Frame 1 - Panel Test Data 186 Frame 2 - Panel Test Data 187 Frame 3 - Panel Test Data 188 Ultimate Strength and Yield Strength of Unexposed 24 ST Aluminum Alloy Panel Material 189 Elevator Section Data - Station 1 190 Elevator Section Data - Station 2 192 Radiant Energy Measurements 194

16

CHAPTER 1

INTRODUCTION

1.1 OBJECTIVE

1.1.1 Purpose of Project

Project 5.1 was established for the specific purpose of obtaining flight test data and information which can be used in defining those regions in space which are unsafe for Model AD type naval aircraft following the delivery of"an"atomic air burst weapon. Much of the flight test information gathered, however, was to be generally applicable to the study of weapon effects on aircraft in the vicinity of an atomic weapon explosion.

1.1.2 Need for Project '

To date, most of the information available for defining unsafe regions has been either static ground test data or of a theoretical nature, based to some extent on limited flight data obtained at comparatively low levels of weapons effects. To reach optimum effectiveness in atomic weapon delivery it is essential that the raps in these data be filled.

Specifically it is anticipated that the data obtained by this project will be used as follows:

a. To experimentally verify or redefine the structurally safe regions ôr a Model AD airplane flying in the vicinity of an atomic explosion.

b. To improve methods for analytically determining the effects of atomic weapons on naval aircraft structures.

1.1.3 Report Objective

The objective of this report is to present the measured, observed, and calculated data associated with atomic weapon effects upon the structure of Model AD aircraft in flight in the vicinity of an atomic explosion. Five separate sets of data,

17

corresponding to Shots 1, 2, 7, 8, and 9, Operation UPSHOT-KNOTHOLE are presented. Direct correlation between certain items of data is presented as observed; however, no attempt has been made to include in this report all the theoretical studies necessary to correlate the data presented.

1.2 EXPERIMENT DESIGN

1.2.1 Background

Reference (l) presents procedures for determining the regions in space which are unsafe for naval aircraft following the explosion of an air burst atomic weapon. In order to investigate these regions and their specific application to AD aircraft, two standard blue AD type aircraft, a Model AD-2, Bureau Number 12236j, and a Model XBT2D-1, Bureau Number 09103, were converted to drone aircraft and instrumented for the measurement of atomic weapon effects. (See Fig. 1.1.)

1.2.2 Instrumentation

The specific items measured were as follows: a. Burst time; by modified photoelectric cell (blue box). b. Thermal radiation: by the Naval Radiological Defense

Laboratory calorimeters and by thermal cloths, both mounted on underside of wing; 0 to j>0 cal/cm .

c. Aircraft skin maximum temperatures; by temperature sensitive papers on reverse side of panels with various surface finishes;+129 to +579°F. See Fig. 1.2, locations 1-16.

d. Aircraft skin and spar temperatures; BN and PN resistance temperature gages; -100 to +400°F and -100 to +250°F, respectively. See Fig. 1.2, locations 17 and 18.

e. Free stream overpressure; by the National Advisory Committee for Aeronautics* pressure pickup on wing tip boom; 0 to 4 psi range, 1000 cps maximum frequency on starboard, 80 cps on port side.

f. Aircraft normal accelerations at center of gravity and tail by Statham accelerometers; 0 to 10g, 70 cps (also 10 cps Giannini accelerometer at center of gravity through JlN/UKR-5 telemeter).

g. Wing bending; by SR-4 strain gage bridges at one inboard station on starboard side and several stations on port side; to limit load, 80 cps.

h. Wing shear at one inboard station on port side. i. Wing torsion at one inboard station on the port side. j. Horizontal stabilizer bending at'one inboard station on

port and/or starboard side. k. Horizontal stabilizer shear at one inboard station on

port and/or starboard side. 1. Aircraft altitude; by modified SGR-584 radar,

18

Electronics Associates plot board and telemetering of Giannini pressure pickups.

m. Aircraft horizontal range relative to burst; by radar plot.

n. Aircraft velocity by radar and by telemetering of Giannini pressure pickup.

o. Aircraft pitch and rbll attitude by telemetering of Giannini pitch and roll gyro.

p. Gamma radiation; by Signal Corps Engineering Laboratories photographic films; 0-50 Roentgen's. See Fig. 1.2, locations A-G.

Fig. 1.1 Model AD Type Drone Aircraft (AD-2)

19

Q CD I- OO

ui >

2D

B-8 I-

o ID

<0\ «oDi

z o

UI 6 CO

o en O {*] D

a. ii.i u UI UI

a. a. < UI u

D _l CO o o

a. \ \ ~Dl

cc -> < r> Q ->. i .j D \\ •*- (0 UJ

<

o (0 UI Q

_i CD

s QQ

In o

CO (0

Pi o

CO o o

u

a: o

u. a CO

-J CO

In o

u ° 1 _l 0.

O o o a Q Q> «0 o d Ü o _ (M (0 ■«■ in (0 h- 00 o>< 3 Sn

UI _J UI o <

UI or CL

•D UI I- < S

< CC UI a. 2

z g 0.

I

id nr

a UI N z

a UI N z ft

UI N Z

c li r ^co z o

0.

O -1

e Q -J O u

UI CO D u or UI

UI I Q Z I

§ U

ct

§ u I

UI i- 5

u CO iii

ui

2 3

UI z> _l

2 3 UI

UI

3 2 3

UI D _l

u - < s -j: is

o z Z s its

Z o 1-

z O l-

UI CO

3 i < CQ < X cu < CO CU < C2 ä CO

UJ Q to Z

0- CL 0. a

to o

CO (0 o (VI

m ? in — c in 5 y Ü Ü

CO

O Ö d

o Ö

o o

o d

o P

o d

o o c d c

J v

3 s 8 8 8 _> o

ü Ü o — <\i (O ^ in (0 r»- oo <J> c > q < GO Ü O UI u o _i _i

n C o

•H

cd Ü

20

The information provided by the above instrumentation is both general with respect to atomic weapons radiation and blast characteristics and specific with respect to aircraft structural response to these inputs. The complete load measuring installation is described m refs (2) and (3). The modified SCH-584 radar used is described m ref (4). Reference (5) describes the airborne instrumentation and gives sensitivities for all records.

1-2.3 Analytical Methods

Aircraft test positions were established following the general procedures described in ref (6).

1«2'3.1 Aircraft Attitude and Orientation

The aircraft was to be in level flight attitude, tail towara the blast in a vertical plane containing the burst point, simulating an escape position of an AD type aircraft following delivery of an atomic weapon. This was accomplished in all tests.

1.2.3.2 Thermal Effects

■ T^}emfl effects were considered applicable at tn. The expression ,or temperature rise in the aircraft skin exposed to radiation from the burst was assumed to be:

AT_ H-eQ SIN 9 0.833t (l.l)

In positioning the aircraft for critical thermal effects an n^nt^°nt ab3°rbt,ivit^ ê. of 0.3 was taken for standard blue painted aircraft skin, considering the minimum skin gage of 0.016 in. Ground reflectivity calculations are based on the methods of ref (9) using the relationship:

2 3 ßx2ho

>xicr-H d

r

It, ' ,Rxiu^- -< 2[(Ah/h,)+l1 0 3. 5 , K ll.68h? (KAh/h,)2+2(Ah/h,)]2 4[(Ah/h,)2+2(Ah/^]fej (1'2}

This equation is valid for the case where h?> h, and as such is suitable for all shot«. (See Figs. 3.3 and23.4)' Mult'lSng lR by the thermal yield in calories (l KT thermal yield = id-12 R

calories) gives the reflected radiation in calories per centimeter2, In accordance with ref (20), a thermal yield - .wflSk was used

21

1.2.3.3 Gamma Radiation

Gamma radiation was considered to have no structural significance. Since the aircraft used in the manned flights were to receive no appreciable amount, gamma radiation was not considered as a positioning criteria; however, it was calculated and measured. The following is the method presented in ref (10) for determining gamma radiation at a given aircraft position: Find reduced range by multiplying R^ by <y . Using this reduced range and the angular position of the aircraft with respect to the burst (e.g., 12 o'clock, aircraft directly over burst; 3 o'clock, aircraft at burst level) find the reduced dosage, D/cr2 for a 1 KT weapon. (See Fig. 3.18.) Multiplying by the factor a 2 Y gives the expected dosage. Since the emission of gamma radiation is time dependent, Fig. 13, ref (8), the aircraft motion is considered by using a method similar to that given in ref (6) to obtain total nuclear radiation received on a moving aircraft.

1.2.3.4 Overpressure

Time of shock arrival (Fig. 3.20) must be known in order to establish the aircraft position at shock arrival for a given test condition. The free-stream overpressures for the test conditions were based on the data of ref (10), free air or surface burst depending on the triple point trajectory, corrected for both height of burst other than sea level and for aircraft test altitude. The relationships used to correct overpressure and slant range for burst height are:

R H-sL(VPJ '/3

(1.3)

APH = APSL (X) (1.4)

For aircraft at test altitudes other than burst altitude the relationships for overpressure and slant range are:

V Vx

AÂPHX(VPH/(VTA)/4

The slant range at a given overpressure for any yield is obtained from the known range and yield by

22

(1.5)

(1.6)

Overpressure at the test position can now be predicted. Fig. 3-22.)

1.2.3.5 Gust Effects

(1.7)

(See

With a given overpressure there is a corresponding pressure ratio at each test altitude. Density ratios and gust velocities corresponding to the pressure ratios can be predicted. (Figs. 3.23 and 3.24.) Reference (11) is a source for these data. The effect of gust velocity on aircraft is determined in terms of allowable aircraft load factors. The following equation, given in ref (6) is used to predict the incremental load factor experienced by the airplane.

SC Ag =

La 2W

whe re:

0.3+0.7 U| w sin 9

(1.8)

U, = ^(l^-w cos 9)Z+ (w sin 0);

p / \/p can be substituted directly for the expression

0.3 + 0.7 in equation 1.8 if

curve is available. For the aircraft used, SC\_a for the wing is taKen as 1840. Wing and tail loads are assumed to be directly proportional to the calculated load factor. Structural dynamic response was not considered in the analysis. The general expression for the calculated load is

AL = A9 [K,W+K2] (1.10)

For the wing, the aerodynamic constant Ki, and inertia constant I<2J are both based on the normal spanwise load distribution identical to that for the positive high angle of attack, forward e.g. condition presented in ref (12), assuming a linear lift curve. The horizontal stabilizer load coefficients were obtained using a value of

(SCLa )HOR.STAB.//(scLa ) WING =0.206

23

and a normal load distribution proportional to the horizontal stabilizer chord, based on information from ref (13). These constants are presented, in Table 1.1. Steady state pullouts performed as a flight check of the load measuring installation result in measured wing loads in direct agreement with loads calculated using this table.

TABLE 1.1 - Calculated Wing and Horizontal Stabilizer Load Constants

Station Load «I K2 x 10-» 3

W.S. 57 Shear 0.436 -1.529

N.3. 57 Bending 37.90 -98.60

W.S. 107 Bending 22.08 -44.30

W.S. 162 Bending 9.77 -17.88

S.S. 21 Shear 0.0799 -0.1255 S.S. 21 Bending 3.515 -10.24

1.2.4 Operational Procedures

Two Navy Model AD aircraft, single place, single engine attack bombers were converted to drones using the Naval Air iSxperimental Station Type I Remote Control Equipment. One^participated in Shot 1, the other in Shot 9 as manned aircraft in^ preparation for drone operation. Lower range data were obtained in this manner.

For N0L0 (no live occupant) operations, the drone take-off was under the control of a pilot at a ground station located beside the runway. Once airborne, the drone was turned over to a pilot flying a Navy Model F8F "mother" control plane. Upon arrival at the test site, control was turned over to a pilot at the radar plotting board. Using data from the radar plot and. from telemetering direct reading meters, the pilot guided the drone to its predetermined position relative to the burst at time zero. Following the test, the "mother" aircraft again took control and the drone was returned to the base where the pilot at the runway station controlled the drone landing. An operational tolerance of ± 1 sec of time in positioning was allowed at time zero.

24

CHAPTER 2

RESULTS AND OBSERVATIONS

2.1 RESULTS

Records of thermal and blast effects were obtained for all shots with the exception of Shot 7 where the drone was destroyed at shock arrival, resulting in the loss of the blast data. Figure 3.2 presents the time histories of measured thermal radiation. Figure 3.7 presents time histories of measured skin temperature rise. Figures 3.25 through 3.31 are reproductions of time histories of measured blast effects. Table 2.2 presents a summary of primary data for all shots. Numerical data, observations and photographs for each shot are presented in Section 2.2 below.

2.2 DATA AND OBSERVATIONS

2.2.1 Shot 1

2.2.1.1 General Data

Date of Test:

Test Aircraft:

Type of Test:

17 March 1953

Model AD-2 Drone; Manned Flight

300' Tower (T-3) Ground elevation 4025 MSL

2.2.1.2

Weapon Yield: 15 KT planned; 16.2 KT Radiochemical

Test Conditions

Altitude of Test Aircraft:

Temperature at Test Altitude:

17,200' MSL; 12,900' above burst

-0.4°F

25

Pressure at Test Altitudes 7.7 psi

True Airspeed: 359 fps

Slant Range from Burst: t0, 14,400'j ts, 16,800'

Time of Shock Arrival: 14.4 sec

Aircraft Gross Weight (Shock Arrival): 13,550 lb

Aircraft Angular Position: tQ, 63°; ts, 50*30' from horizontal; (0)

2.2.1.3 Effects Data - Shot 1

(1) From Weapon

Thermal Radiation (cal/cm2)

Port Wing Calorimeter: 2.7: (2.9, filter correction, B%) Normal to Horizontal Plane: 3.2

Gamma Radiation (r)

Less than 0.1 inside the aircraft

Thermal In Cockpit (cal/cm )

Less than 1.0, from temp, cloth and paper

Free-stream Overpressure (psi)

Starboard wing pressure pickup: 0.3

(2)'On Test Aircraft

Incremental Normal Acceleration (g's)

Center of Gravity Accelerometer: 2.1 Tail Accelerometer: 2.0 Center of Gravity, Calculated: 0.97

Aircraft Attitude Change (deg)

Pitch Gyro: 2.3 Nose Down Roll Gyro: 1.4 Right Wing Up

26

Incremental Structural Loads Above lg Level Flight - Shot 1

Measured Load

Starboard Wing Bending (Sta. 57) 512 x 103

in. lb.

Port Wing Bending

Port Wing Bending

Port Wing Bending

Port Wing Shear

Port Wing Torsion

(Sta. 57) 525 x 103

in. lb.

(Sta. 107) 308 x 103 in. lb.

(Sta. 162) 112 x 103

in. lb.

(Sta. 57) 4.63 x 103

lb.

(Sta. 57) -6.5 x 103

in. lb.

Port Stabilizer Bending (Sta. 21)44.3 x 103

in. lb.

% of Design Limit Load

10.6

10.9

12.7

9.3

11.6

0.7

16.7

Port Stabilizer Shear (Sta. 21) 1.22 x 103 lb. 15.2

Temperature Rise in Aircraft Skin, °F (Temperature Gages)

Color

Standard Blue

Standard Blue

2.2.1.4 Oscillograph Data

See Figure 3-31

2.2.1.5 Observations

Mo visible damage

Skin Thickness Temperature Rise

0.051 57

0.040 97

27

2.2.2 Shot 2


Date of Test: 24 March 1953

Test Aircraft: Model AD-2 Drone; No live occupant (NOLO) flight.

Type of Test: 300' Tower (T-4), Ground elevation 4308' MSL.

Weapon Yield: 40 KT Planned, 24.5 Radiochemical

2.2.2.2 Test Conditions

Altitude of Test Aircraft: 10,800' MSL; 6200'above burst.

Temperature at Test Altitude: 33°F

Pressure at Test Altitude: 9.90 psi

True Airspeed:

Slant Range from Burst:

Time for Shock Arrival:

477 fps

tQ 8100', ts 10,900»

8.0 sec

Aircraft Gross Weight (Shock Arrival):

Aircraft Angular Position:

2.2.2.3 Effects Data

13,360 lb.

tG, 50°; ts, 34° from horizontal; (9;

(1) From Weapon

Thermal Radiation (cal/cnr)

Port Wing Calorimeter:

Tail Calorimeter:

Port Wing Thermal Indi- cator Cloth:

Tail Thermal Indicator Cloth:

Normal to Horizontal Plane:

11.5; (12.4, filter correction, 8%)

9.8; (10.6, filter correction, 8%)

9.5

11.0

12.8

28

Gamma Radiation - Shot 2

Location Dosage Shielding Material Thickness, in.

Inside Cockpit:

Behind Headrest 4.2 Aluminum 0.174 Rubber and cloth 0.75 Gasoline 74.8

On floor, forward 4.2 Aluminum 0.683

On floor, port 2.5 Aluminum O.464 Rubber and cloth 0.75 Gasoline 35.6

Inside Wing Fold:

Port 6.0 Aluminum 1.55

Starboard 6.0 Aluminum 1.55

Inside Rear Fuselage:

Starboard 8.8 Aluminum 0.040

Free Stream Overpressure (psi)

Starboard Wing Pressure Pickup: 1.0

(2) On Test Aircraft


Center of Gravity Accelerometer: 3.8 Tail Accelerometer: 3.6 Center of Gravity, Calculated: 2.14


Pitch Gyro: 2.5 Nose down Roll Gyro: 0

29

Incremental Structural Loads Above 1 g Level Flight - Shot 2

Measured Load % of Design Limit Load

Starboard Wing Bending (Sta. 57)

947 x 103

in. lb. 19.7

Port Wing Bending (Sta. 57)

829 x 103

in. lb. 17.2


559 x 103

in. lb. 22.9


328 x 103

in. lb. 27.3

Port Wing Shear (Sta. 57)

9.00 x 103 lb. 22.5

Port Wing Torsion (Sta. 57)

29.0 x 10 in. lb.

2.1

Port Stabilizer Bending (Sta. 21)

72.5 x 103

in. lb. 26.7

Port Stabilizer Shear (Sta. 21)

1.86 x 103 lb. 23.2

Temperature Rise in Aircraft Skin, op _ Shot 2

Color Skin Thickness Temperature Rise

Standard Blue 0.064 >171 <190

Standard Blue 0.051 >171 < 190

Standard Blue 0.051 156

Standard Blue 0.051 *187

Standard Blue 0.040 *225


Black 0.051 > 171 4 190

Black 0.051 ■>171 <190

Aluminized Lacquer 0.064 138

Aluminized Lacquer 0.040

30

156

Aluminized Lacquer 0.040 ^ 144 ( 156

Aluminized Lacquer 0.040 138

heat Resistant White 0.040 138

Heat Resistant White 0.032 >144 < 156

*■ Temperature gage measurements. All other are temperature sensitive paper measurements.

2.2.2.4 Oscillograph Data

See Figs. 3.29 and 3.30. Values used for the approximately 0.1 sec interruption of strain gage amplifier operation were estimated, based on damping ratios established using strain gage records from Shots 1, 8 and 9.


All standard blue paint on 0.016 alclad aircraft skin on 'the under side of the control surfaces was scorched except in the Immediate vicinity of the rivets (see Figs. 2.1, 2.2 and'2.3). All standard blue on 0.020 skin was scorched except where in contact with heavier structural members (see Fig. 2.4). None of the standard blue paint on .025 skin was scorched with the exception of the curved portion aft of the rear shear web wnicn was directly facing the burst point (see Figs. 2.4 and 2.5), and the hinged aft edge of the outboard wing panel which was exposed to direct radiation from the underside as well as indirect radiation received through the aperture between this panel and the aileron (see Fig. 2.6). Slight irregular scorch on blue paint marks were noted on the underside of the outboard wing panel on skin up to and including 0.032 (see Fig. 2.7).

The rubberized fabric aerodynamic seals between the elevator and horizontal stabilizer were destroyed (see Fig. 2.8). The aileron seals were not directly exposed to thermal radiation and these remained intact. The fabric lightening hole covers on the horizontal stabilizer rear spar were burned through where direct thermal radiation was received (see Fig. 2.9). The airplane in this test had covered wheel wells so the tires were not exposed. There was no thermal damage to the tires.

2.2.3 Shot 7


Date of Test: 25 April 1953

Test Aircraft: Model AD-2 Drone; N0L0 Flight, Aircraft Destroyed.

31

Type of Test: 300» Tower (T-l), Ground Elevation 4238' MSL

Weapon Yield: 33 KT Planned, 43-4 KT Radiochenical


Altitude of Test Aircraft: 10,450' MSLj 5900' above burst




Slant Range from Burst: t0, 6,200'j t®, 6700'

Time for Shock Arrival: 3.75 sec

Aircraft Gross Weight 13,400 lb. (Shock arrival):

Aircraft Angular Position: t0 73°; ts 60°30' from horizontal; (6)


(l) From Weapon

Thermal Radiation, (cal/cm2)

Port Wing Calorimeter: 19.4; (20.9, filter correction, 8%)

32 to 35 (corrected for l/2 to 3/4 fireball miss)

Normal to Horizontal Plane: 54.6

Free Stream Overpressure (psi) - Shot 7

2.7 estimated. Based on Project 5.1 flight test data and observation, as given in this report.



Center of Gravity Acceleration: 16, estimated. Based on calculated 8.25 g, Project 5.1 flight

32

test data and observation, and assuming that thermal damage had no effect on acceleration.


Pitch Gyro: 65 Nose Down Roll Gyro: 85 Right Wing Down

Incremental Structural Loads Above 1 g Level Flight

Estimated Load

in. lb.


3.38 x 106 70

3.38 x 106 in. lb. 70

2.08 x 106 in. lb. 85

0 .93 x 106 in. lb. 80

35.6 x 103 lb. 90

50 x 103 in. lb. 5

04 x 103 in. lb. 110

7.79 x 103 lb. 100

Starboard Wing Bending (Sta. 57)


Port Wing Bending (Sta.107)

Port Wing Bending (Sta.l62)

Port Wing Shear (Sta. 57)

Port Wing Torsion (Sta. 21)

Port Stabilizer Bending (Sta. 21) 304

Port Stabilizer Shear (Sta. 21) 7

Based on Project 5.1 data and observation, as given in this report.

Temperature Rise in Aircraft Skin (°F)

Color , Skin Thickness Temperature Rise



Temperature Rise in Aircraft Skin (*F) - Shot 7

Temperature Rise

400

450*

300

800

Color Skin Thickness



Bare Aluminum 0.066

Bare Aluminum 0.016

33

Heat Resistant White


O.O64

0.016

200

600

* Temperature sensitive paper measurement. All others estimated, based on Project 5.1 flight test data and observation, as given in this report.

2.2.3.4 Oscillographic Data

All pertinent oscillographic data for the period after shock arrival was unintelligible. The time history of thermal radiation measured from burst time to time of shock arrival is included in Fig. 3.2.


In this test the actual yield exceeded the planned yield by greater than 30 per cent. Since the aircraft was positioned for near critical weapon effects based on the planned yield, the resulting thermal radiation severely damaged or weakened all the blue painted skin on the underside of the wing. The thermal damage to the aircraft skin is presented in Table 2.1.

TABLE 2.1 Thermal Damage to Aircraft Skin - Shot 7

Surface Finish Effects Figure No. 0.016 Skin Thickness

Standard Blue Panels missing except for small circular pieces around each rivet

2.10; 2.11

Standard White Panels missing except for edges 2.12; 2.19

Aluminized Lacquer Panels missing except for edges 2.13; 2.14


Panels remained intact 2.15

Bare Aluminum Pieces missing; panel broken along center line

2.15

Standard Blue 0.025 Skin Thickness

Panel missing except for edges 2.10; 2.13

Standard White Pieces missing; panel broken along center line

2.12

Aluminized Lacquer Panel broken along center line 2.13

34

TABLE 2.1 - Continued

Surface Finish

Aluminized Lacquer


Bare Aluminum

Standard Blue

Eeat Resistant White

Standard Blue

Aluminized Lacquer

Standard Blue

Effects

Pieces missing; panel broken along center line

Panels remained intact

Panels remained intact

0.032 Skin .Thickness Panel missing except for edges.

NOTE: White lettering gave added protection.

Panel remainted intact

0.040 Skin Thickness Pieces missing; panel broken

along center line, except on narrow panels. NOTE: White lettering gave added protection.

Panel remained intact

0.051 Skin Thickness (AL-3-S0) Skin apparently unaffected; paint

scorched

Fipure No. 2.14

2.15

2.15

2.16

2.17

2.18; 2.14; 2.16

2.18

2.16

The blue 0.025 skin on the hinged aft edge of the outboard wing panel in some places suffered the same damage as 0.016 skin on other parts of the wing. This panel was exposed to direct radiation from the underside as well as indirect radiation received through the aperture between this panel and the aileron (see Fig. 2.14). This is seen to be consistent with the damage experienced in Shot 2, as shown in Fig. 2.6.

The fabric lightening hole covers destroyed in Shot 2 were replaced with fabric covers coated with aluminized lacquer. These covers were slightly scorched but not burned through (see Fig. 2.19). Aileron seals remained intact (see Fig. 2.11). The top side of the wing gave no indication of thermal damage other than softening of standard blue paint on 0.016 skin on the aileron (see Fig. 2.20) where the under side skin was completely destroyed.

No effects directly attributed to overpressure were noted.

35

The above information was obtained from the wing outboard panels and sections of the elevator and horizontal stabilizer \*hich were torn off at shock arrival and received only minor additional damage due to the fall. Evidence indicates that the wing failure resulted when the thermally weakened skin was further damaged by shock effects, thereby reducing the over-all strength of the wing. The sections of the empennage were torn off when struck by the wing panels, -

2.2.4 Shot 8


Date of Test: 19 May 1953

Test Aircraft: Model XBT2D-1 Drone, NOLO Flight

Type of Test: 300' Tower (T-3a), Ground elevation 4025' MSL

Weapon Yield: 37 KT Planned; 27 KT Radiochemieal


Altitude of Test Aircraft: 11,200' MSL; 6,900' above burst



True Airspeed 386 fps

Slant Range from Burst: t0, 7200'; ts, 8100'

Time for Shock Arrival 5.6 sec

Aircraft Gross Weight (Shock Arrival): 12,800 lb.

Aircraft Angular Position: t0, 73°; ts, 58° from horizontal; (0)

36

2.2.4.3 Effects Data - Shot 8

(1) From Weapon

Thermal Radiation (cal/cm2)

Port Wing Calorimeter:

Starboard Wing Calorimeter:

17.8j (19.2, filter correction, 8% 24.9; (26.9, filter correction, &%

Port Wing Thermal Indicator Cloths 17 Starboard Wing Thermal Indicator Cloth: 27 Normal to Horizontal Plane: 25.4

Gamma Radiation

Do sage (0

Shielding Location Material Thickness

Inside Cockpit 28 Aluminum 0. 25 to 0.75 in. Rubber anc cloth 0 to 0.75 in. Gasoline 5 to 15 in.

Inside Wing Fold, Por* 38 Aluminum 0.051

Starboard 38 Aluminum 0.051

Inside Rear Fuselage 4^ Aluminum 0.040

Underside of Starboard Wing 52.5 No shielding


Free Strean Overpressure (psi)

Port Wing Pressure Pickup: 1.8


Center of Gravity Accelerometer: Tail Accelerometer: Center of Gravity Calculated:


Pitch Gyro: 8.1 nose down Roll Gyro: 4 right wing up

8.6 8.5 4.52

37

Incremental Structural Loads (Above 1 g Level Flight) - Shot 8

Measured Load

Starboard Wing Bending (Sta. 57) 1,970 x 103 in.lb.

Port Wing Bending (Sta. 57) 1,950 x 103 in.lb.

(Sta. 162) 546 x 103 in.lb.

(Sta. 57)

(Sta. 57)

Port Wing Bending

Port Wing Shear

Port Wing Torsion

15.97 x 103 lb.

35.00 x 103 in.lb.


41

40.5

45.5

40.0

2.5

Starboard Stabilizer Bending (Sta. 21)

Port Stabilizer Bending (Sta. 21)

Starboard Stabilizer Shear (Sta. 21)

Port Stabilizer Shear (Sta. 21)

154.5 x 103 in.lb.

139.6 x 103 in.lb.

3.71 x 103 lb.

3.54 x 103 lb.

56.9

51.4

46.4

43.6

Color

Standard Blue

Standard Blue

Standard Blue

Standard Blue

Standard Blue

Standard Blue

Standard Blue

Standard Blue

Standard Blue

Temperature Rise in Aircraft Skin. °F

Skin Thickness Material Temperature Rise

0.064

0.0£4

0.051

0.051

0.051

0.051

0.040

0.040

0.032

24ST

75ST

24ST

24ST

24ST

24ST

24ST

24ST

225

>250 <288

>250 <351

>250 <288

>250 <351

250

>322 < 442

>288 <442

>351 < 442

38


Standard Blue 0.032

Standard Blue 0.025

Standard Blue 0.020

Standard Blue 0.020

Standard Blue 0.016

Standard Blue 0.016

Black 0.051

Black 0.051








Standard White 0.040



Heat Resistant White 0.040





Material

52SH34

24S0

24S0

24ST

24ST

24ST

24ST

75ST

75 ST

243T

24ST

24ST

24ST

24S0

24ST

753T

24ST

24ST

52SH34

52SH34

Temperature Rise

> 322 < 442

> 401 < 4Ö2

> 482 < 579

> 401 < 462

> 442 < 579

>351 <442

351

>250 <351

> 174 < 190

225

> 225 < 250

250

>351 <442

>351 <442

> 401 < 482

> 174 < 225

250

> 225 < 250

225

156

225

>156 <174

>156 <190

39







Bare Aluminum 0.040

Bare Aluminum 0.040

Bare Aluminum 0.040

Bare Aluminum 0.032

Bare Aluminum 0.025

Bare Aluminum 0.020

Bare Aluminum 0.016

Bare Aluminum 0.016

Material Temoerature Rise

>156 <190

24S0 >171 <190

24S0 >190 <225

24ST >171 <225

24S0 >190 <225

24ST 225

24ST 190

24ST >225 <250

52SK34 >225 <291

250

24S0 >288 <322

24ST >250 <351

24ST >250 <351


Figures 5.25 and 3.26 contain direct reproductions of the oscillograph traces at time of shock arrival. These traces were separated as shown to eliminate the overlap of traces brouît about by the high level of the measured effects data.


In this test the complete underside of the aircraft was stripped down to bare aluminum with exception of panels on the aileron and elevator (see Figs. 2.21, 2.22, 2.23, 2.29 and 2.30)._ In addition, thermal panels consisting of 0.016 to 0.064 alclad aircraft skin with different paint finishes were installed on the underside of both flaps (Figs. 2.24, 2.25, 2.26. 2.27) as well as on several locations on the wing.

No thermal damage was visible on any unpainted aircraft surface. The-0.016 blue painted control surface tips had the paint completely burned off and the surface itself was badly warped. Due to the fact that all flap specimens were hand painted under field conditions, direct visual comparison with specimens prepared by standard procedures was not readily obtainable; however, the surface damage in all cases appeared consistent with the temperature

10

rise data given under "Results." These temperatures can be taken as indicative of the thermal damage to the surface finish.

The fabric lightening hole covers destroyed in Shot 2 were replaced with heat resistant white covers. These remained intact.

The airplane in this test had uncovered wheel wells. The tires, exposed to direct thermal radiation> were covered with heat resistant white paint. They experienced no thermal damage.

2.2.5 Shot 9


Date of Test:

Test Aircraft:

Type of Test:

Weapon Yield:


Altitude of Test Aircraft:

8 May 1953

Model XBT2D-1 Drone, Manned Flight

Air Drop, Height of Burst 2432' above ground, Ground Elevation 3708' MSL

J>1 KT Planned, 26.0 KT Radiochemical

21,900' MSL; 16,400' above burst

Temperature at Test Altitude: -17°F



Slant Range from Burst:

Time for Shock Arrival:

t0, 16,400'; ts, 18,000' at first shock; 18,700' at third shock.

15.5 sec first shock; 19.8 sec third shock r

Aircraft Gross Weight (Shock Arrival): 13,150 lb

Aircraft Angular Position: t0, 85°; ts, 64° at first shock, 60° at third shock; (e)


(1) From Weapon

Thermal Radiation (cal/cm )

Port Wing Calorimeter: 1.9; (2.0, filter correction,

8%) 2.8 to 3.3 (corrected for 1/2 to 3/4 fireball miss) Normal to Horizontal Plane: 3.6

Free-Stream Overpressure (psi)

Starboard Pressure Pick-up: 0.5 first shock Starboard Pressure Pick-up: 0.1 third shock



Center of Gravity Accelerometer: 2.3 First Shock, 1.0 Third Shock.

Tail Accelerometer: 2.6 First Shock, 1.1 Third Shock. Center of Gravity Calculated: 1.42 First Shock, 0.43

Second Shock.


Pitch Gyro: 3.2° First Shock, 2.4° Third Shock, Nose Down.

Roll Gyro: 3° First Shock, 3-7° Third Shock, Right Wing Up.

Incremental Structural Loads above 1 g Level Flight

Item and Location Measured Load % of Design Limit Load

First Shock

Third Shock

Starboard Wing Bending 584 x 103 243 x 103

(Sta. 57) in.lb. in.lb.

Port Wing Bending (Sta. 57) 524 x 103 230 x 103

in.lb. in.lb.

Port Wing Bending (Sta.l62) 147 x 103 64 x 103

in.lb. in.lb.

Port Wing Shear (Sta. 57) 4.14 x 1C3 2.7 x 103

lb. lb.

Port Wing Torsion (Sta. 57) 18 x 103 -5.5 x 103 in.lb. in.lb.

Starboard Stabilizer Bending ^3.2 x 103 18.4 x 103

(Sta. 21) in.lb. in.lb.

First Shock

12.1

11.8

12.3

1Q.3

1.3

15.9

Third Shock

5.1

4.8

5.3

6.7

0.4

6.8

42

Item and Location Measured Load First Shock

: Third j Shock


First Shock

: Third : Shock

13.1 5.7

22.0 7.5

13.9 4.5

Port Stabilizer Bending 35.7 x 10? 15.5 x 10? (Sta. 21) in.lb. in.lb.

Star Stabilizer Shear 1.76 x 10* 600 lb. (Sta. 21) lb.

Port Stabilizer Shear 1.11 x 103 360 lb. . (Sta. 21)

Temperature Rise in Aircraft Skin (°F)

Color Skin Thickness Temperature Rise

Blue 0.012 185 *

^Results of three identical readings.


Figures 3.32 and 3.33 present the oscillographic data associated with the first and third shocks. The second shock, which occurred 0.5 seconds before the third shock, was so small that no accurate effects readings could be made.


No visible effects damage.

TABLE 2.2 Summary of Primary Data for All Shots

Test Conditions Shot 1 Shot 2 Shot 7 Shot 8 Shot 9 ( 1 9 5 3 )

Date of Test 17 Mar 24 Mar 25 Apr 19 May 8 May

Test Aircraft AD-2 Manned

AD-2 N0L0

AD-2 N0L0

XBT2D-1 N0L0

XBT2D-1 Manned

Test Area T-3 Yucca

T-4 Yucca

T-l Yucca

T-3 a Yucca

Frenchman Flat

Ht. of Burst (ft MSL)

4325 4608 4538 4325 5510

Ground Eleva- tion (ft MSL)

4025 4308 4238 4025 3078

43

TABLE 2.2 (Continued)

Test Conditions Shot 1 Shot 2 Shot 7 Shot 8 Shot 9

Planned Yield (KT) (Radchan)

15 40 33 37 31

Actual Yield (KT)

16.2 24.-5 43.4 27.0 26.0

Aircraft Alti- tude (ft MSL)

17,200 10,800 10,450 11,200 21,900

Aircraft Ht above Burst (ft)

12,900 6200 5900 6900 16,400

Temp, at Test Altitude (°F)

-04 33 43 39 -17

Pressure at Test Alt.(psi)

7.7 9.9 10.1 9.8 6.3

Wind Velocity at Test Alt. (deg and fps)

270; 67

160;

27 265; 17

210;

25 250; 88

True Airspeed (fps)

359 477 417 386 420

Aircraft True Course (deg)

290 270 283 277

Slant Range at t0 (ft)

14,400 8100 6200 7200 16,400

Slant Range at 16,800 10,900 6700 8100 18.000(a) 18,700(c)


Test Conditions Shot 1 Shot 2 Shot 7 Shot 8 Shot 9

Aircraft Weight at tg (lb)

13,550 13,360 13,400 12,800 13,150

Position Relative to Burst at tQ (deg)

63 50 73 73 85

Position Relative to Burst at tg (deg)

50.5 34 60.5 58 64(a) 60(c)

44


Test Results

3.2 12.8 54.6 25.4 3.7 Thermal Radiation* (cal/cro2) First Third

Time of Shock Arrival (sec)

14.4 8.0 3.75 5.6 Shock Shock 15.5 19.8

Overpressure (psi) 0.3 1.0 2.7 1.8 0.5 0.1

e.g. Acceleration (Ag) Meas.

2.1 3.8 l6(est) 8.6 2.3 1.0

Tail Acceleration (Ag) Meas.

2.0 3.6 I6(est) 8.5 2.6 1.1

Incremental Wing Loads (% of Design Limit Load)

Inboard Bending 11 18 70 40 12 5

Inboard Shear 12 22 90 40 10 7

Incremental Tail Loads {% of Design Limit Load)

Bending 17 27 110 51 13 6

Shear 15 23 100 45 15 5

* Normal to horizontal plane.

,...4?fP¥ s##f&f^* '

_J J* * "«mr" V?

P

<.••"• m

Sr-'" I

"life

. SSxw

Fig. 2.1 AD-2, Starboard Aileron Showing Scorched Blue Paint on 0.016 Skin - Shot 2

45

vO rH o

mm.

c o -p c •H CO a.

I •d

o o Ü to

c

§ 10

&, •H F-<

bO Ö

'S cd

ö o (H Q)

rH

cfl cv o ,a -P f* o CO si +i CO

i

c\T a I -H

Q ^! «0 CO

t>0 •H

46

Fig. 2.3 AD-2, Port Elevator and Horizontal Stabilizer Showing Scorched Blue Paint on t).016 Skin - Shot 2

47

CO

5? w

u o « CM <D

• tn o O 0)

■r OH ■+>

(1) «al 3 cq

13

d o

•H -P

-c; o O P-, u O TJ O d)

C Ü •H

in

hOCV C •H-P SO ^x

•-< _, '

fcxO Jsi to E) C/l fn

•H 3 3UNPD

CM +J O Q) s-i •»£ OO-P

PL,

•<■ o c CM -H

I XI o Q Ö n) -< cri&H

CM

•H

'1»-~Iiilli|&l/!v -

48

T3

>

o taO

.5 o fc c

•H Ai co ir\ C\J O cv O p o C Ä o CO

p i Ö

•H ■p nl en

CU b <D w 3

rH 0> m ja

4> •rt a> Wl

J3 a o •H f-, o o 01 o Cn r/l

.O hi) <1) C&

* fn O 01 Ä Q) CO XI

CO P. ert JH

rH 01 U< <l>

n-, tu» C «M

•H o Is

+> P «H fn < O a, Ö o ^ •w

CM p 1 h

CJ o ■=< fU

ir\

CM

* bO

•H UH

4-9

Fig. 2.6 AD-2, Port Wing Showing Scorched Blue Paint on 0*025 Skin on Hinged Aft Edges of Wing Panel - Shot 2

50

CM

-P O Si co

i

•H

CO

CM

8 c o

a. 0)

H m

<D .c o u o o

CO

o CO

o. •H

c ed

r-f

a.

-p

o 04

7

<*», •H

51

"'"'ttft.;

bO c

•H > O xi C/l

u m N

•H rH CV •H ,0 P tfl O -p X CO w H 1 crt -P H Ö fl) O <1) N> r.o

•H ^ f> C) •H

M-» cR

n ^, ?-. O (1) -P -a! ni > 0 (1) •H

Ö S-,

«1 G I*, (I) 0) Tl £ 0) p N 0 TJ (» tt) X! *H £> 3 0

-P ft; u (1) TJ u. (1) <d >»

0 *\ U cv P

1 tn 0 <D << C_"

to 0!

W) •H U-,

52

m WijiM

O •H f-,

.£> ctf

TJ CD C

ep tic

1 en

CD N

•H r-l •H .O cti

CO

* O N CV

•H

O

CO

I

U en et) f-. ft 0

en > o

u o CO CD CD

PS rH O

O

r" -P •H C CD

•>-P

I hO Q -H

53

to •H tx,

10

H <D C

PH

cv O

•6 s H O * o

I

3 be C

o

c o CD

(4 CO

I cd

■+* CO

o u ©

£

o r-i

Of

bO •H

54

-p o

£1 co

-p o

c

■a CO

c

d)

o CO

O U 03

o a.

« CM

Pi

s!

55

o CO a c

•H -P

5 o 03

o •H

CO > O P Q)

cö O- s rf P

°.S CD CO F>

•H I P CO w H m

CD CD

A; tu ü Ö -rj

o CO -H

A! I CO

Ö -g & cd CD

rH .C ■r) CO

Ö

"£ &, O

P-. CD Ü

•s CO CV «H

I U Q 3 52 co

C\i

•H

56

T3 <D U <D

& Ü £^-

iJ -P O

•o X <D co N

I

•H G S -H P î H CO

UA O W -P O

CD o to

g -g 3 -p Q Ö

•H iH efl 03 CL,

E CD Q) P £ r-\ E-H PQ

ÖCÖ Ö U

•H CO £ TJ

CO -P en,

I

Ö -P

s-g CD

rH TJ •H 0

CO -Ci ft b s to O O O

•H -P X CO rO

•xLT\ cv cv

I o

9

rH

C\i

•H

o

57

•ö

U Q) -p

«J »1 i

•o c

r S o

o -Ö .p (D

-p a) d b0v3 cd ctf

73 s

c C -P

0)

1 CO

5

*H

e o o

+3 C ^ -H O Jrf

I O

CM

•H

58

o

3 o

•H

! .a

s

iS o -p a> bO

«sis

•H I -P

3 s

1* CO

I CO

si M CO (0

3| -P -H

O o

«CO

ir\

CM

S

59

J>- ' i

'tap«"

mr* -tor" m^jSto*!-. «»««<'*' "*•'!£**,»*». MM**

Fig. 2.16 AD72, Outboard Port Wing Panel and Wing Tip Showing Wing Tip Intact. Note how white star protected 0.032 skin. - Shot 7

60

Fig. 2.17 AD-2, Outboard Port Wing Panel Showing Heat Resistant White Painted 0.032 Plate - Shot 7

61

Fig. 2.18 AD-2, Outboard Port Wing Panel Shovd.ng Aluminized Lacquered 0.040 Plate - Shot 7

62

-8 Si en

HO 1 a t ■p o o crS x: -P en ß

M u © a. N •H

•H H H •H ^ XI O rt P P CO CO >

(I)

3 0 -P c -D o g

•H fci [0 O h K a>

> n o c o m

a) fn iH O o P rc etl > W) V ß d-d

a> CH P O X!

W) P,

•H d E-

CT- H

CM

63 iH U<

63

% A O U o o

CO

bO

o ro

P-. CO

bfl I

I« -P CO ^ „

Pk H O

«H • 0 o

•Ö o •H 05 +3 O. Ö O -H

E-i a) Pu

CV J5 1 3

Q

o cv CM

hfl •H &k

03

^rv.^.v<**.«i;^

6*

CO

-p o X CO

I

O X U CO o Ü ir\

CO CV O

bO •

5° .c to CO

o o <D O

c o ?-.

& O Co

•H •a!

CO O

h - cd »-} -P CO -Ö

Q) * N

H -H I ß

Q -rj CM 0

rH CV

CV

bO •H CTH

65

CO

■p o t* .. CO

■H ' «) C

■P -H CO .M

CO TJ 0) o O O h • 5 O o

CO -o c

M cd c

O O .C • CO o

c * o ô Q> O

rH •

31° Ö

-P O

° -if P-. £ •H

* cö rH fL,

I Q <D CV +3

cv

•H

66

to a>

I-) to a. •H O

CO

ID •

a> a» •p ü o cd

• 03 CO tu T)

& o CO

H cd

■ö

o CO

I CD

£ •H cö

P-.

'S ü

& ü

CO

bOTJ C ©

•^ O -H si cd co P.

o 3 -p H cd ,0 !>

■0 -P c ^ cd o +3

a-, to

•»vO H rH

I O <=> • CM O

£j c JxJ o

cv cv

faO •H

67

ippsojiigf

Oq

■

Emdamsmm ''ft* ' /■>'

%" *•'&•-• V"'

'M '.'•' "• (1) kî?^^.' : H ^VÖ"^* "*;;V

'■ 'i-'t • . ■51 ti u

H 0) X •-:'. "#,•;••••*= E-<

-p

-p

-p

I

o •H

•H -P

CUP-. 03 r-\ -Ö fc. ^

cö faO O

T3

o •8

-P

O

P n C

H I 05

Q H

><! P-.

-4-

bO •H

68

^mf^m^mm

, 4 #$$

\ä|#*

A'

|vl j- \

, V

p 03 0)

EH to

cd -P

fee? EH i

-p ° 5 -H 3= -P - U {X O til p-, H

bfl cd c s S c

M

U •> cd T3

.8,5?

cog

H

e g X OH

Oi

•rl CM

69

i^ CO

-P -P cn o V X

EH CO

H 1 nl

ß ^ Ü <1> •rl X -P M U

o .C P-. -P

Ü CTt

ti. O CÖ rQ iH +3 In <S W) a •*

ad T> ^ <D

-P 3 h et) O -p

P-> cn G

•V H H

1 co O H CM <». H C

ö

vO cv • <M

W) •H fe

70

HP mm

■g to

© o CO

C Ö a) o EH -P

*!

a) o H £>

H bO Ö *

•H -d IS (D

rH +3 r-l

O -P a. m

G •>M

H A i3 O H CV (1)

PL,

CM

■H Pn

71

<D 71 H PJ

V, Si o U o o

CO

t>0 c ■s CO o .c -P r/> O

si tJOCO ß

•H i

G «H •H O .* m fc (1) CM

■a c9 a> • o o

•* Ö iH o Q -P CV ß

e •H trt

XI Pi

to CM • CM

• bC

•H tu

7/.

72

» 0 H fi-)

"E B cd i-4 •g ,Q ß CO Tf

■P fn CO CO

TJ TJ

Ja S -P

8 03

o ß o O CO

m 60 m ß H

■s P. o •H x: U CO

a> P, -P

•H o H £3

2 « •H ß ^ •H 130

^J ß o

(0 •P

o£ 0) -d- co H O 3 •

o I

V T( ß u c •H CO CO AJ 5 vO

CO

hHÔ cd O H

-P * O CO O •

O •* ß H

1 o TJ

O +> p CV ß ß

K •H Ctl

•H cr)

>< Pk p.

o <N

CM

W)

■ö <]> Ä ü !H O O

CO

bO C

? o .c r/1

to J-i n ■P -p O

CÖ J3 > cn fi) H i W c T-f •H Ö ^ cd co U o <D tO o

•H • H o •H

-9 •n es) -p CD 0!

CV Tl <"-\ ?H o 0) • O C) ,0 SH c es! o -P CO -p

c •v •H

rH cfl 1 P-,

U CV m H 3

B H

o cn • CV

• W)

•H U<

74

CHAPTER 3

DISCUSSION

3.1 GENERAL

The data in this report permit the determination of a relatively- complete picture of weapons effects on AD type aircraft for a representative escape position following the delivery of an atomic weapon. In the presentation of these data, certain relationships are indicated by means of curves superimposed on the basic data. Limited theoretical analyses were employed in arriving at these curves and it is conceiv- able that some changes may result when more detailed analytical methods are applied. In all cases, however, the recorded data presented herein are considered accurate to within ± 5 percent of the maximum readings obtained.

3.2 THERMAL EFFECTS

3.2.1 General

Although the flight test thermal radiation and temperature data as measured on the inside face of the aircraft skin show reason- able consistency and agreement with thermal damage and with expected values, initial results of metallographic studies of the internal structure of skin specimens from these flights (Appendix A) indicate effects normally associated with temperatures far in excess of the temperatures measured. This condition is believed to be a result of microscopic thermal concentrations in the grain structure of the material.. These concentrations are associated with the extremely rapid rate of thermal radiation. Structural tests on these specimens indicate no loss in strength corresponding to the indicated metallographic damage. Accordingly the thermal data presented in this report are considered to adequately represent the thermal information necessary for aircraft structural considerations. The data presented in Appendix B should prove to be of assistance in further analysis. It is to be noted that although this phenomenon was not evidenced in the ground specimens, it is believed that local conditions, such as the dust layer, might have resulted in this discrepancy.

75

3.2.2 Theimal Radiation

The aircraft and calorimeter orientation relative to the burst point at t0 for all shots is given in Figure 3.1. Aircraft positions with respect to Ground Zero at t0 were determined from automatically plotted radar tracking data. Calorimeter orientations with respect to the burst points and their fields of view were obtained from motion picture films taken from GSAP cameras which were essentially mounted on the axes of the calorimeters. Figure 3.2 gives the value of thermal radiation as measured by the various calorimeters. These two figures indicate certain facts with respect to ground reflection. In Shot 8 two calorimeters were used. The port wing calorimeter was pointed back toward the burst point but aimed above it by approximately 30*. The starboard wing calorimeter was mounted to measure thermal radiation normal to the wing and was aimed approximately 15° below the burst point. The direct readings of these two calorimeters were 17.8 and 24.9 cal/cm^ respectively. These values, increased to 19.2 and 26.9 when corrected by + 8 per cent for the quartz filter on the calorimeter; were verified by corresponding readings of 17 and 27 on identically oriented cloth thermal indicators. The calculated direct radiation received by these calorimeters is 12.3 and 13.6. Using ground reflectivity calculations with a least squares value of albedo of 0.55 based on data from Shots 1, 2 and 8, the calculated total thermal radiation for these two installations is 20.8 and 25.7. These values are seen to correspond closely to the true calorimeter readings. The fact that the starboard wing calorimeter would have been exposed to maximum effects of ground reflection in addition to direct radiation from the fireball appears to justify its reading. The effect of reduced ground reflection on the port calorimeter similarly explains its lower reading. In Shot 9 the position of the fireball is approaching the ± 45° field of view limit of the calorimeter and as such it is expected that a large portion of the direct thermal radiation from the burst would be missed. This is more strongly evidenced in Shot 7 where the fireball is partially outside of the calorimeter field of view. Almost the entire thermal reading for this condition would be due to reflected thermal radiation. Considering all such factors, estimates for thermal radiation received normal to the wing can be made for the shots in which it was not directly measured. Figures 3.3 and 3.4, based on ref (9), present the reflectivity information used for this work. A comparison of calculated and measured calorimeter readings is given in Figures 3.5 and 3.6. The results appear to verify the ground reflection effect indicated in ref (7), where reflection from the ground was estimated to be 50 per cent of the direct radiation.

3.2.3 Aircraft Skin Temperature Rise

Temperature rise in aircraft skin, was initially assumed to be directly proportional to the heat received and to the reciprocal of skin thickness. The time histories of temperature rise shown on Figure 3.7 indicate that the cooling rate is an important factor in this problem. Figure 3.8 shows the general agreement between measured cooling

76

9P^^*-

«i o +3 ^ CO X f-.

1- ^ L_ o

< -p

(D ÜJ > o •H

z -P CO

< H CC 0)

_J c < o

•H H +3

z o

CO

N •H

o O

I CD -P <D

5 U o 1 o "3 fi 0)

+> <H tfl U O JH

•H «Jj

s rvi rvi a) to ^ w O co <o ^ cvi

c0lx±J~13A31 V3S NV3W 3A03V NOI1VA313

O (VJ to

•rl

77

CM ö CM

if) Ö it)

2W0/Jvo~NOI±ViaVti -1VWYGHJ.

78

^^m&

rates and those calculated following the methods of ref (14), if the proper assumption of turbulent or laminar flow is made.

A lowering of effective absorptivity values for thinner skin independently of surface finish is shown in Figures 3.9 through 3.13 for Shot 8 specimens. The curves through the test points were drawn with an attempt to minimize this effect. Figure 3.14 summarizes these data in the form of a family of curves based on readings from temperature sensitive papers. Two temperature sensitive gage readings are included as an indication of the general agreement of data measured by two different devices. Figure 3.15 is a summary of temperature rise data from Shot 2. Although the data were limited, it can be seen that curves of the same characteristic shape fit the points. An indication of the range of equivalent absorptivities for different surface finish and skin thickness for Shot 8 and Shot 2 is given in Figure 3.16. The effect of aerodynamic cooling does not appear to fully explain the lowering of the effective absorptivity values for thinner skins. This can be seen in Table 3.1 which compares the equivalent absorptivity values for representative cases of surface finish and skin thickness in Shot 8 with absorptivities considering the methods of ref (14). The same general trend of lower absorptivity for thinner skin is indicated here. In a discussion of an equivalent absorptivity value it is important to remember that it is dependent on the amount of heat assumed to be received by the material. Any amount of heat added or lost as a result of thermal damage to surface coatings will have a tendency to influence this value. Using average values for these absorptivities and the assumed normal thermal radiation of 55 cal/cm^, the temperature rise in 0.016 and 0.064 in. skin for Shot 7 was estimated. The two estimated values for skin with an aluminized lacquer finish appear to substantiate the 450°F temperature rise in the 0.040 in. skin. The predicted temperature rise appears consistent with the observed damage for the aircraft skin in this shot. References (15) and (l6) present the results of preliminary metallurgical studies conducted in an attempt to determine the approximate temperature attained by the aircraft skin in Shot 7. The data in Appendix A are the results of additional work of a similar nature.

The extreme importance of aircraft skin surface finish is readily seen in the results from Shot 7. Heat resistant white painted panels remained intact where similar panels with standard blue paint, subjected to identical thermal radiation, were completely destroyed. Bare aluminum held up almost as well as heat resistant white but in all cases experienced higher temperatures. It is to be noted here that heat resistant white and standard white paint afford roughly the same protection as long as the finish is not charred. At thermal levels high enough to seriously affect the standard white surface, it proves to be no better than aluminized lacquer. Field analysis of the structural failures which caused loss of the drone in Shot 7 were conducted by project personnel and representatives of the Structures Department, Douglas Aircraft Co. Indications are that the aircraft might have survived had the under skin been bare aluminum or painted heat resistant white instead of standard blue.

79

TABLE 3.1 - Calculated Absorptivity Values - Shot 8 (Turbulent Flow Assumed)

S 1 3 E C I M E N A B S 0 R P T I V I T Y Color Skin Thickness (in.) Reference (11) Figure 3.13

Standard Blue

0.016

0.025

0.232 - 0.398

0.415 - 0.505

0.262

0.040 0.436 - 0.710 —

0.064 0.483 0.464

Bare Aluminum

0.016

0.025

0.159 - 0.232

0.245

0.172

0.040 0.274 —

0.064 — 0.256


0.016

0.025

0.103 - 0.141

0.178 - 0.280

0.146

0.040 0.215 —

0.064 — 0.244

The coating of other thermally vulnerable items, such as fabric seals and rubber tires, \n.th heat resistant white paint permitted them to withstand approximately four times the thermal radiation that they could in their natural state.

3.3 NUCLEAR RADIATION

3.3.1 Gamma Radiation

Gamma radiation was considered to be the most important of the nuclear radiations. Gamma measurements were made at several locations on the aircraft. These measurements are presented in Figure 3.17 as compared to calculated gamma radiation where aircraft motion, density at point of burst, and aircraft clock position were considered. The reading underneath the wing, where there was no protection from direct radiation, was the highest reading. The lower readings, inside the rear fuselage, inside the wing folds and inside the cockpit for both Shot 8 and Shot 2, were roughly proportional to the amount of structural shielding associated with these locations. The materials provid-

80

^ «?^^^^^^^^^TOB

ing shielding are listed under Results for these two shots. The data from Shot 1 and Shot 9 were too low to be of significance, being less than O.lr. All pertinent gar-ma infomiation is summarized on the basis of a 1 KT yield in Figure 3.18.

3.3.2 Neutron Radiation

The effects of neutron radiation were assumed to be negligible for all test positions.

3.4 OVERPRESSURE AND GUST EFFECTS

3.4.1 Overpressure

Shots 1, 2, 7 and 8 had 300 ft burst heights while Shot 9 had a burst height of 2400 ft. On the basis of the triple point trajectory curves in Figure 3.19 it can be seen that the aircraft position for Shot 9 was in the free air region. For all other shots the aircraft was in the region of reflection and the calculated pressures were considered to be produced by a weapon with a yield of 1.8 times the given radio-chemical yield.

For these conditions, measured overpressures were found to be in close agreement with calculated when the method of ref (17) was employed in reducing the data. This consisted of taking the best straight line through that portion of the overpressure record which exhibited conventional smooth behavior, and extending it back to time of shock arrival.

Time of shock arrival for ground burst data was found to be in close agreement with shock arrival times in the region of reflection. Converting the free air yield of Shot 9 to an equivalent ground burst yield using the 1.8 reflection factor, brings this point exactly on the (Figure 3.20) shock arrival curve. A comparison of measured and

calculated overpressure is given in Figure 3.21. Figure 3.22 presents the measured overpressures in terms of a standard 1 KT free air burst. The range of pressure and density ratios investigated is indicated in Figure 3.23. The associated gust velocities are given in Figure 3.24. The curves included here are good only for standard atmospheric conditions. Since temperature is the factor governing the relationship between gust velocity and altitude, an equivalent temperature altitude based on measured temperature must be used for non-standard atmospheric conditions.

No overpressure damage was noted up to approximately 2 psi. In no case was any adverse overpressure effect on the structure noted except when it occurred in conjunction with extremely high skin temperatures.

3.4.2 Gust Effects

Gust effect's were recorded in terms of accelerations and structural loads resulting from measured overpressures. These records are reproduced in Figures 3.25 through 3.33. In all cases wing gust

81

effects were noted approximately 0.01 sec after the tail as indicated in Figures 3 .25 and 3.26.

On the basis of a peak e.g. acceleration reading at a time associated with peak wing loads, and an average of two peak tail accelerations occurring in the time interval between peak tail and peak wing loads, excellent correlation was obtained between wing and tail acceleration. A comparison between these measured accelerations and corresponding calculated accelerations is given in Figure 3.34. The measured accelerations are seen to be almost twice the calculated values for all tests. How much of this is due to elastic structural

Id" 9 8 7.

0 /-h 2=3,000 \ \ /

6

5 A'

^-\

SHOT 7- *\er

/-h9=. 5,000 3'

2

lö"- 9 8 7 6

SHOT 1 ^"w

SHOT 2

rhi = 11,600 =7,000

T ^*"««^ \ /

/ -— O ^Ô ^ = \: J,340 2

(M 5

5 A SHOT 1 -J o H-l

XL 2. <OhSV c

-12 10 "

9" O" 7.

ft. D

5 C ) 2 I A € > S \ 1 D i; 2 14

HORIZONTAL RANGE, R^FTX I03

Fig. 3«3 Reflected Thermal Intensity from a Unit Source- Tower, hi = 300 ft.

82

accelerations affecting the overall acceleration picture and how much is due to limitations in the theoretical approach to the prediction of these accelerations must be evaluated by future laboratory tests and theoretical study. The fact that measured acceleration can be predicted consistently, however, is of great importance in view of the general agreement between measured and calculated structural loads and the relationship between load and acceleration (see Equation 1.10). The comparison between measured and calculated loads is made in Figures 3.35 through 3.39. Although in this particular airplane the dynamic overstress appears to approximately compensate for the gust alleviation, the fact that the recorded loads exhibit frequencies corresponding to first symmetrical bending illustrates the importance of

LEGEND

HEIGHT OF BURST ABOVE GROUND h,= 2,000 FT h =3,000 FT

HORIZONTAL RANGE, RH~FTXI03

Fig. 3.4 Reflected Thermal Intensity from a Unit Source— Air Drop h-^ = 2432 feet

83

considering aeroelastic effects even on rigid aircraft. Figure 3-35 shows the time-history of the pitching motion of the drone aircraft during the period of shock arrival for Shots 2 and 8, indicating that pitching motion had a negligible effect on aircraft loading during the first 0.1 sec after shock arrival.

CM

< ü

I z o

< ec

2 QC UJ X H Q U a. D (0 - U> a: 0.40 Ld z UJ 0.30 z o < 0.20 a < tc .J < 0.10 2 ft 0.08 UJ I 0.06 _J < fc 0.04 H

0.03

0.02

0.01-

0= CALCULATED TO POINT I "^ A = CALCULATED TO HORIZONTAL

PLANE AT POINT

O = MEASURED CORRECTED FOR FILTER(+ 8%) ®= ^- DIRECT RADIATION MISSED

(D = 3. DIRECT RADIATION MISSED

X= TOTAL ON HORIZONTAL PLANE (INCLUDING REFLECTIVITY)

NO REFLECTION

STBD PORT

WING WING

SHOT 7-

SHOTS-J

•PORT WING -TAIL

3.0 6J0

SHOT 2

X O-

SHOT I- SHOT 9-

-PORT WING X CH®\ PORT

9.0 12.0 ^ 15.0 SLANT RANGE ~FTx|Q

18.0 21.0

Fig. 3.6 Measured and Calculated Thermal Radiation Reduced to 1 KT

85

Cü

•H «

& -P

is EH

•H

CO

-P

f-. Ü U

•H •a;

xf CD U

CD

<D

O

o -p 03

•H

0) t:

•PI EH

Jo ~ ±N3I9WV 3A09V 3umvu3dW3±

86

CO <D

-P

c •H rH O O O

•o 0) -p CO H 3 O H cd o

C cfl

TS CO U

CO CO 0)

f-, c

CM

O -P 00

•H

^

U

-P cö U 0) (X s 0)

c •H

CO

cn cti $H o U

•H <c!

«0

•H (x,

wnwixvw iv/iv 87

600-r

* -

O Q d o

o CM in o ^t CO CM CM o O o o o O o o

(0 q d

12 24 36 48 60 RECIPROCAL SKIN THICKNESS ('/t) ~ IN"'

Fig. 3.9 Incremental Temperatures Measured in Aircraft Skin with a Heat Resistant White Paint Finish - Shot 8

88

0 12 24 36 48 RECIPROCAL SKIN THICKNESS ( '/f) ~ IN'1

Fig. 3.10 Incremental Temperatures Measured in Standard White Painted Aircraft Skin - Shot 8

89

"<t — O CVJ <o in ^- ro o o o o odd d

tO O (0 CM CM — O O o O O o

36 48 RECIPROCAL SKIN THICKNESS (1^) ~ IN"1

Fig. 3.11 Incremental Temperatures Measured in Unpainted Aircraft Skin - Shot 8

90

RECIPROCAL SKIN THICKNESS (Vt) ~ IN'

Fig. 3.12 Incremental Temperatures Measured in Aircraft Skin with an Aluminized Lacquer Finish - Shot 8

91

■■■

* , o c\j (0 «n *<t en o o o o

(\l o Ö

o O Ö

(0 O Ö

RECIPROCAL SKIN THICKNESS ('/t) ~IN -i

Fig. 3.13 Incremental Temperatures Measured in Standard Blue Painted Aircraft Skin - Shot 8

92

0 12 24 36 48 60

RECIPROCAL SKIN THICKNESS ( '/f ) ~ IN"1

Fig. 3.14 Temperature Rise in Aircraft Skin vs the Reciprocal of Skin Thickness for Different Surface Finish - Shot 8

93

LEGEND A=BLUE ©=ALUMINIZED a=WHITE(HEAT RESISTANT)

12 24 36 48 60

RECIPROCAL SKIN THICKNESS ('/t) ~ IN"

Fig 3.15 Temperature Rise in Aircraft Skin vs the Reciprocal of Skin Thickness for Different Surface Finish -

Shot 2

%

z <0

(O Q O

a. e

> H Q. DC O CO CO <

z y I s

0.64

0.56

0.48

0.40

0.32

0.24

0.08

'0 008 0-6 0.24 032 0-40

EQUIVALENT ABSORPTIVITY FOR 0.016 SKIN

Fig. 3.16 Comparison of Equivalent Absorptivity for 0.016 and 0.064 Aircraft Skin with Different Surface Finish

95

0 8 16 24

CALCULATED GAMMA RADIATION

32

ROENTGENS

Fig. 3.17 Comparison of Measured and Calculated Gamma Radiation Indicating Structural Shielding Effects

96

LEGEND

A=CALC. AIRCRAFT IN MOTION

BsCALCô POSITION

-1.0-0.01 4 5 6 7

REDUCED RANGE(d R A

Fig. 3.1Ö Measured and Calculated Gamma Radiation Reduced to 1 KT in Sea Level Homogeneous Atmosphere

97

20

16 O x H u. I a z

§

ui

§ (D 8

u

SHOT 8

SHOT 7

SHOT 9

SHOT 2

300 FT BURST HEIGHT

SHOT

2400 FT BURST HEIGHT

5EF^CT' 6 8 10

HORIZONTAL RANGE~FTXI03

Fig. 3.19 Triple Point Trajectory Curves Indicating Aircraft Position at Shock Arrival

98

EQUIVALENT GROUND BURST YIELDS

SHOT 1

YIELD 16.2

2 245 7 43.4 8 9a

270 14.5; r~f

£ FREE AIR BURST CONVERTED TO EQUIVALENT GROUND BURST USING 1.8 REFLECTION FACTOR

SLANT RANGED FT X I03

Fig. 3.20 Measured and Calculated Shock Arrival Times vs Slant Range

99

CX8 1.6 2.4 3.2 4.0

CALCULATED OVERPRESSURE ~ PS I

Fig. 3.21 Comparison of Measured and Calculated Overpressure

100

■SHOT 7

O MEASURED OVERPRESSURE /CORRECTED TO FREE AIR^ IBURST DATA )

SHOT 9 a

CALCULATED (FREE AIR BURST DATA)

I I M 111 3 5 7 10 20

SLANT RANGE ~ FT XIO3 30 50 100

Fig. 3.22 Measured and Calculated Overpressure Reduced to 1 KT in a Sea Level Homogeneous Atmosphere

101

1.00 1.04 1.08 1.12

DENSITY RATIO~ ßfe

.16 1.20

Fig. 3.23 Pressure Ratio vs Density Ratio

102

,<«^riiHu*iiUKVi»f£l& h

1.00 L08 1.16 1.24 1.32

PRESSURE RATIO'—^p

1.40

Fig. 3.24 Gust Velocity as a Function of Pressure Ratio and Altitude

103

TRACE NO. 14 TAIL ACCELERATION

SHOCK ARRIVAL,TAIL; 5.6 SECONDS FROM T0

STBD. HOR. STAB. FRONT SPAR SHEAR , STA. 21

STBD.HOR.STAB.REAR SPAR SHEAR, STA. 21

TIME ~ SECONDS X 10 r2

Fig. 3.25 Time History of Tail Acceleration and Tail Loads - Shot 8

104

TIME ~ SECONDS x 10 S-2

Fig. 3.26 Time History of e.g. Acceleration, Wing Loads and Overpressure - Shot 8

105

y»v :>//

INCREMENTAL HOR. STAB. SHEAR STA. 21

5.65 5.70 5.75 5.B0 TIME AFTER BURST~ SECONDS X I02

Fig. 3.27 Measured Aircraft Acceleration and Tail Loads - Shot 8

106

20 CO o 15

X 10 (0 m 05 _J

-0

fO 600 o X 400 <n CO _l 200 z

-0

<*)2000 -1500 £|000 CO 500 z -0

2 (f) 1 Q. —0

5. 60

INCREMENTAL WING SHEAR STA.57

INCREMENTAL WING BENDING STA. 162

INCREMENTAL WING BENDING STA. 57

OVERPRESSURE

5.65 5.70 5.75 5.80

TIME AFTER BURST~ SECONDSX|0'

Fig. 3.28 Measured Overpressure and Aircraft Wing Load - Shot 8

107

TRACE NO- A9-3.8I CG. ACCELERATION

J

18

/

PORT HORIZ.STAB. BENDING, STA. 21

PORT HORIZ. STAB. SHEAR, STA. 21

-OVERPRESSURE

SHOCK ARRIVAL, WING, 8.01 SECONDS FROM T0

-SHOCK ARRIVAL, TAIL; 8.0 SECONDS FROM T0

TIME^ SECONDS X 10 v-2

Fig. 3.29 Time History of e.g. Acceleration, Tail Acceleration, Tail Loads and Overpressure Shot 2

108

TRACE NO.

I

6

8

10

/

STARBOARD BENDING, STA. 57 PORT BENDING, STA. 57-

PORT REAR SPAR BENDING, STA. 57

PORT BENDING, STA. 107

PORT BENDING, STA. 162

/

PORT FRONT SPAR SHEAR, STA. 57

y

SHOCK ARRIVAL,WING; 8.01 SECONDS FROM T0

TIME ~ SECONDS X I0~2

Fig. 3.30 Time History of Wing Loads Indicating Shot 2 Estimated Peaks Based on Wing Dynamic Response

109

SHOCK ARRIVAL TAIL (14.40 SEC3 AFTER %) -

^-—SHÖcTARRIV/ärW'IG (14.41 SECS AFTER TQ)

--£- REFERENCE-^— • ' ■

1. Stbd. Wing Bending, Sta. 57 8. 2. Port Horiz. Stab. Shear, Sta. 21 9. 3. Fort Wing Bending, Sta. 57 10. 4. Port Horiz. Stab. Bending, Sta. 21 11. 5. Port Wing Rear Spar Bending, Sta. 57 12. 6. Port Wing Bending, Sta. 107 13. 7. Thermal Radiation 14.

18. Over Pressure

Port Wing Bending, Sta. 162 Thermal Radiation Port Wing Shear, Sta. 57 Over Pressure Indicated Airspeed C.G. Acceleration Tail Acceleration

Fig. 3.31 Time History of Blast Data - Shot 1

110

BLUE BOX.

'--is'-«-

: 17 ■' ,'X/ ^\

-^-vv^ . srCÂ^ÂO;

. w-^^-'VA

. jfc- 1st SHOCK ARRIVAL WINGfitl5.51 SECS AFTER TQ)j*jf —Wist SHOCK ARRIVAL TAIL ( 15.50 SECS AFTER T0)£^KFERENCl$j»-

FH5ÜS01 SECr~ : TREFERENCE -^—

1. Stbd. Wing Bending, Sta. 57.5 9. Thermal Radiation 2. Port H0riz. Stab. Shear, Sta. 21 10. Port Wing Shear, Sta. 57.5 3. Port Wing Bending, Sta„ 57.5 11. Over Pressure 4. Port Horiz. Stab. Bending, Sta.21 12. Stbd. Horiz. Stab.Rear Spar 5. Port Wing Rear Spar Bending Sta.575 Bending, Sta. 21 6. Stbd. Horiz. Stab. Front Spar 13. C. G. Acceleration

Bending, Sta. 21 14. Tail Acceleration 7. Thermal Radiation 16. Stbd. Horiz.Stab.Front Spar 8. Port Wing Bending, Sta. 164 Shear, Sta. 21

17. Stbd. Horiz.Stab.Rear Spar 18. Over Pressure Shear, Sta. 21

Fig. 3.32 Time History of Blast Data - Shot 9a (First Shock, Shot 9)

111

r ' •' T"

BLUE BOC iö4 14r - - -

_7

■ ~.,-v^/v"v

_6

j* .w^ i ^ro*yQ*>^zf<v^^?~s?x^tzK.

-W! 1 «"wV^V."-*.

r3rd SHOCK ARRIVAL WING (19.SO SECS AFTER T0>:

J$FERENC|£ — ^7

-1EFEREN3- ,«A*#UX üiijwsS——

1. Stbd. Wing Bending, Sta. 57.5 2. Port Horiz. Stab. Shear, Sta. 21 3. Port Wing Bending, Sta. 57.5 4. Port Horiz. Stab. Bending, Sta. 21 5. Port Wing Rear Spar Bending Sta .57.5 6. Stbd. Horiz. Stab. Front Spar

Bending, Sta. 21 7. Thermal Radiation 8. Port Wing Bending, Sta. 1&+ 9. Thermal Radiation

18. Over Pressure

10. 11. 12.

13. 14. 16.

17.

Port tfing Shear, Sta. 57.5 Over Pressure Stbd. Horiz. Stab. Rear Spar

Bending, Sta. 21 C. G. Acceleration Tail Acceleration Stbd. Horiz.Stab.Front Spar

Shear, Sta. 21 Stbd. Horiz.Stab.Rear Spar

Shear, Sta. 21

Fig. 3.33 Time History of Blast Data - Shot 9c (Third Shock, Shot 9)

112

SHOT 7

OCENTER OF GRAVITY ACCELERATION GKTAIL ACCELERATION

2 4 6 8

CALCULATED NORMAL ACCELERATION ~Ag

Fig. 3.34 Comparison of Measured, and Calculated Aircraft Normal Acceleration

113

HOT 8

CALCULATED INCREMENTAL TAIL SHEAR^LbXlO3

Fig. 3.35 Comparison of Measured and Calculated Horizontal Stabilizer Shear

1U

O X

O z O Z u

_J

z u

id oc o z

O ÜJ DC D «0 <

2

SHOT 8

CALCULATED^ MEASURED

CALCULATED INCREMENTAL TAIL BENDING~lrvLbx|03

Fig. 3.36 Comparison of Measured and Calculated Horizontal Stabilizer Bending

115

0 4.0 80 12.0 160 20O

CALCULATED INCREMENTAL WING SHEAR~LbX|03

Fig. 3.37 Comparison of Measured and Calculated Wing Shear - ¥. S. 57

116

2.8

2.4

to O x 2.0 _l

Id flu o z

1.6

1.2 _l

£ Z u 2 u a: u 2 0.8 a u a: CO < ÜJ 5 0.4

50% DES GN LIMIT BENDING STA 57

MEASURED = CALCULATED

SHOT 2

-SHOT 9 a

-SHOT I

SHOT 8

0=STARBOARD EXPORT

SHOT 9c

J 1 U- -1 I 1_ _l L -J » u.

0 0.4 08 1.2 1.6 2.0

CALCULATED INCREMENTAL WING BENDING ~ In. Lb x I06

Fig. 3.38 Comparison of Measured and Calculated Wing Bending - W. S. 57

117

O X

_l

O Z Q Z u m o z

P z

u £E O Z

o U £E

to < UJ 5

CALCULATED INCREMENTAL WING BENDING ^|n. Lb X 10

Fig. 3.39 Comparison of Measured and Calculated Wing Bending - W. S. 107 and W. S. 162

118

if)

CO 5 Q -p Z Q £ O M LJ C f) •H

i ü -P

_l •H P-i <

> s ■° g H

ft < JU •H

* <C

u <H o O I to r?

o 2 -p

o 0C

•rl

U. (D

5 E-i

O H 3 cn

M •H Pn

Oi if)

dfl 390N NMOa 3SON

S33y03CI~3DNVhO 310NV HDlId 3NVlddlV

119

CHAPTER 4

CONCLUSIONS AND RECOMMENDATIONS

4.1 CONCLUSIONS

4.1.1 Thermal Effects

a. Thermal radiation on aircraft in flight in clear air is a result of ground reflection as well as direct radiation from the fireball. Additional thermal radiation may also be received from clouds through reflection, depending upon cloud location with respect to both aircraft and burst point. As such, refs (1) and (8) appear to be inadequate for predicting thermal effects on aircraft in flight unless reflectivity calculations, such as are outlined in ref (9), are included .

b. Heat resistant white paint provides aircraft skin with a greater degree of protection from thermal radiation than any^ other surface finish tested. Next to heat resistant white, bare aluminum appears to offer the best protection against thermal damage. Aircraft skin painted standard blue is extremely vulnerable to thermal damage, being able to stand less than one-third of the thermal radiation with- stood by skin painted with heat resistant white. Use of white or ligtt colored paint on tires, aerodynamic seals and other dark colored materials provides similar protection against thermal radiation.

c. Microscopic thermal damage as indicated by metallographic tests on aircraft skin specimens does not appear to be an important factor where structural considerations associated with thermal radiation are involved. Temperature rise measured by temperature sensitive papers and strain gages on the back of these specimens appear consistent with structural thermal damage and as such are believed to be applicable to the investigation of aircraft thermal limitations as treated in this report.

d. Thermal damage to paint finishes, such as standard white, may modify its absorptivity characteristics. Equivalent absorptivity for a given finish appears to be affected by skin thickness and aerodynamic cooling. Reference (14) provides a satisfactory means of cal- culating the cooling rate if the aerodynamic flow conditions are known.

120

4.1.2 Overpressure and Gust Effects

a. When the yield of the weapon and the position of the airplane are known, the time of shock arrival may be predicted with negligible error.

b. The methods of predicting overpressures encountered by aircraft in flight in the vicinity of an atomic blast produce results which are in agreement with measured values to within ± 0.1 psi.

c. Overpressure up to at least 2 psi has no adverse effect on the structure of AD type aircraft. Overpressure effects in Shot 7, for approximately 3 psi, were masked by the associated thermal damage.

d. The methods of ref (1) used in predicting the acceleration introduced to AD type aircraft in flight by blast, appear to provide agreement with a sustained or effective acceleration rather than with the peak acceleration measured in flight. This is seen as a result of the close correlation between measured loads and those calculated using the accelerations determined by these methods.

4.1.3 General

a. Orientation of aircraft in flight with the longitudinal axis pointed directly away from the burst should result in reduced thermal and gust effects when compared with the positions investigated in this report.

b. The installations for the measurement of blast effects, as well as those used for the measurement of thermal effects, had certain limitations in that the effects as measured may have been modified by dynamic response characteristics associated with the airplane. Addi- tional information concerning thermal and shock inputs might be obtained were comprehensive laboratory tests to be conducted to establish the dynamic characteristics of the complete measuring systems as installed in the test aircraft.

4.2 RECOMMENDATIONS

The following recommendations are made, based on data presented in this report.

4.2.1 Thermal Effects

It is recommended that: a. Additional thermal radiation measurements in flight be

made when opportunity permits, in order to further evaluate the effect of ground reflectivity. Time histories of aircraft skin temperature rise for various thicknesses and surface finishes should be included. Metallographical studies should be made on all specimens to provide additional data on measured temperatures and thermal damage.

b. Theoretical methods for predicting temperature rise and thermal effects on aircraft skin be furthör investigated in order to provide closer correlation with thermal phenomena as measured and observed.

121

c. In those atonic weapon delivery problems where critical thermal conditions exist, consideration be given to the use of a skin finish of naval aircraft with better thermal characteristics than that of the standard blue paint. Tires, aerodynamic seals, and other exposed materials should be similarly protected.

4.2.2 Overpressure and Gust Effects

It is recommended that: a. Dynamic analysis methods for predicting wing and tail loads

on AD type aircraft be investigated in an attempt to provide additional correlation with flight test data.

122

SYMBOLS AND DEFINITIONS

a« Speed of sound at aircraft altitude, feet per second.

du Speed of sound at burst height, feet per second.

Cps Instrument operational frequency, cycles per second.

C|_Q Slope of lift curve, per radian.

D Gamma dosage, roentgens.

g Normal accelerometer indication, non-dimensional.

Ag Incremental normal accelerometer indication, non-dimensional.

hi Burst height above ground, feet.

hQ Aircraft height above ground, feet.

Ah Aircraft height above burst, feet,hp—h .

IR Unit reflected thermal radiation from ground, per centimeter .

MSL Mean sea level.

NOLO No live occupant in drone aircraft.

p Atmospheric pressure at aircraft altitude, pounds per A square inch.

p Atmospheric pressure at burst height, pounds per square H inch.

Pi Peak instantaneous pressure, pounds per square inch.

AP Overpressure, pounds per square inch.

Q Total thermal radiation, calories per square centimeter,

R Range from point of burst to point in space, feet.

R/\ Range from point of burst to aircraft, feet.

R Range from point of burst to point in space at burst H altitude, feet

123

S Aircraft wing area, square feet.

t Aircraft skin thickness, inches.

t0 Time of burst, time zero.

t Time of shock arrival, seconds from 1Q .

T. Absolute temperature of aircraft altitude, °R.

Tu Absolute temperature at burst altitude, °R. H

AT Aircraft skin temperature rise, °F.

Ui Peak instantaneous aircraft velocity, feet per second.

I). Initial aircraft velocity, feet per second. A w Gust velocity, feet per second.

W Weight of aircraft, pounds.

Y Weapon yield, kilotons.

Q Angle between horizontal and line joining burst point with aircraft, degrees.

ß Albedo, percent of thermal radiation reflected by a point, non-dimensional.

X Fuchs factor, non-dimensional; -Z"(^D~*^T ^) X~e v H H '

fi Equivalent thermal absorptivity factor, incorporating all heat loss considerations, non-dimensional.

p Initial ambient density at aircraft altitude, slugs per A cubic foot.

p Ambient density at burst altitude, slugs per cubic foot.

p Peak air density, slugs per cubic foot,

r / Altitude density ratio at burst altitude; H/p

SL

12A

APPENDIX A

METALLURGICAL TESTS OF SKIN SPECIMENS TAKEN FROM MODEL AD TYPE AIRCRAFT EXPOSED IN FLIGHT

TO THERMAL RADIATION FROM AN ATOMIC EXPLOSION

by

RICHARD SCHMIDT

JOHN ERTHAL

NAVAL AIR MATERIAL CENTER

PHILADELPHIA 12, PENNSYLVANIA

125

APPENDIX A

METALLURGICAL TESTS OF SKIN SPECIMENS TAKEN FROM MODEL AD TYPE AIRCRAFT EXPOSED IN FLIGHT

TO THERMAL RADIATION FROM AN ATOMIC EXPLOSION

A.l OBJECTIVES

The primary objectives of these metallurgical tests were: a. to determine the strength properties of the various aircraft

skin specimens after exposure in flight to the thermal radiation from an atomic explosion.

b. to determine the approximate magnitude of the maximum temperature rise in the various aircraft skin specimens by ascertain- ing whether or not melting of the clad and/or base metal occurred.

A.2 METHOD

A.2.1 Strength Tests

Specimens for tensile tests were manufactured from samples of the aluminum alloy skin cut from various portions of the underside of the test aircraft which participated in Shots 2, 7, and 8. Skin samples were obtained having different thicknesses and surface finishes. Due to the restricted size and/or condition (thermal damage, etc.) of the various skin panels from which samples were cut, sub- standard size tensile specimens had to be manufactured.

SR-4 type strain gages were mounted on the various test specimens and tensile tests conducted to determine ultimate strength, yield strength, and per cent elongation. Where practical, at least three tensile specimens were obtained and tested from the same skin sample.

A.2.2 Metallographic Examinations

Small specimens cut from each of the various skin samples were mounted, polished, and subjected to microscopic examination.

By comparing the data obtained from the metallographic examination of these specimens with existing data on the solidus and liquidus

126

temperature of the alloys, an estimate of the maximum temperature reached by each specimen could be obtained. For example, the solidus temperature for 24S aluminum alloy is 935°F and the solidus of the cladding is 1190°F. If melting has occurred in the base metal and not in the cladding, the temperature reached was between these values.

A.3 RESULTS

A.3.1 Strength Tests

The physical properties of ultimate strength, yield strength, and per cent elongation of the specimens tested are listed in Table A.1 for Shot 8 and in Table A.2 for Shots 2 and 7. No attempt has been made to compare these results with "standard" values due to the fact that sub-standard size test specimens were used (elongation data not comparable) and no test specimen material in its original condition was available for establishing realistic reference standards. However, the following handbook values as obtained from ref (18) are listed for the different types of aluminum alloy specimens tested:

Material

24S T3 Alclad

24S0

61S T6

75S T6 Alclad

AL3S0

52SH34

52SO

Ultimate Strength psi

64,000

27,000

45,000

76,000

16,000

37,000

27,000

Yield Strength, % psi ELongation/2 in.

44,000

11,000

40,000

67,000

6,000

31,000

12,000

18.0

19.0

12.0

11.0

30.0

10.0

25.0

A.3.2 Metallographic Tests

Results of the metallographic examinations indicating the maximum temperatures reached by the test specimens and their microstruc- tural characteristics are listed in Table A.l for Shot 8 and in Table A.2 for Shots 2 and 7. For purposes of comparison, the maximum temperatures attained by the different test specimens as measured during Shot 8 are also listed in Table A.l. These measurements were made by means of maximum indicating temperature sensitive papers covering the range from 130°F to 580°F in specified increments. The temperature papers were mounted on the interior surface of the aircraft skin. Similar data are not available for comparison with the metallographic test results listed in Table A.2.

127

CO

ö o W 0

l-H

IS

«0 I- ü

1 •H c D CO

CO ■P ü (D o

•rl

-p ü Lü cc

< bO o c ü C Ig*

•H -P M ü C n

•H

1? £F r- 2 •H iH •H »H •H fl> iH •H -H 5P ä. <ü -H </> -p -p -P -P © -P -P -P -P +3 C -H

•H cö -p f^

+3 -P

r- OH © a> © <u a* ü pH

© © ^■ä -P s -P S E" g s -p S d ^ s

Ü s s © CO

o f§ © o

X

< rr u O 3 2

° < < Q

8 o H H

O

8 H

^8 rH Q o

H rH

1° © c£

ld DC 2 a

*

° 2uOP rt * & to CV

to CV

to CV

to «r» O c^ vr\CV ir\cv e*- cv vO CV vO cv

<l-j<OCl cv V A

cv V A A A

r7V A

rnV A

SV A

*0 -Nl «3d- N0UV0N013

CV irv

-p

00 cv o

cv o cv IT»

• cv o •

cv o • «0 u ff

% 3 a cv t-i

to rH

to rH CV

r- «0 u

— ISd — cv o cv

C

i-H

8 O

8 rH

8 t-i O

8 I- X

0H3IA ©

6

•* ITS rH 5 » 3

1- o 7 — l9d —

cn 3 O

CV

cn O ^ 8

u er

<0 V t>- o o iH <r\ SS3ä±S cn o.

A O MD

0) 3±vwinn M3

CO CV vO vO cr\ vO

SH

OT

NO

. A

ND

S

PE

CIM

EN

LO

CA

TIO

N

ON

T

ES

T A

/ c

co ß o

•H •P „ ET

rH

to = = =

© 43

in rH

co

rH =

-* E-< o o E^ EH © EH

-|Vlti3XVlN CO 3 3 3 CO co

CO

CV cv CV CV cv CV CV

7f w ■A •A •A •A 1 1 -A SS C <D c a> c © c © SS ß CD

< UJ ff cd 0 cd 3 cö 3 cd V cd 3

NO

MIN

T

HIC

KN

A

ND

C

OLO

t

CO w £0 ää £ä co m S5ä ■p cJ co W

O "Ö cv r.

O T3 CV U

IT* T3 CV U

CV TS CV "O

O cd • o

o cd • o

O cd • o

O rt • o

• o

O cd • o

O cd • o

128

©

■H

1 O

3

«0

b c •H o o D -p •H C •H

«0 «0 H +3 o +> u © Ü 10 CJ

cr 2£ u a © TJ © © Ü o QC •H •P © T3 -p •H •H

tO < -P M

o a o

•H ££ +5 »H H © £ß° S2 «»

5 © -H +3 •H © •H 5^ © TH u ■p -p •H 43 ^ S Ä +J 4.1 r- £^ 'S) © -P

*o O •& © ^g la ä H ©

o S ß •H g CO ,0 •H a s ^ S o © H rH rH o o I n

m o CO O CO is a

< oc UJ o _J

3 m to 0 < > g o

u a

6 o H O © o

PQ 8 rH H

O o o rH

o ft rH

s§ m to

*

T°F

. E

MP

. P

E

&

AG

E

AT

A

o » to H o H rH H rH iTvCV c>-oj t>cv H H

<r-£OQ 3v av 8V 5v £v V V A A A A A

*

R -Nl Ü3d- cv o

CM cv o

cv o

cv o

«0 N0I±V0N013

%

• rH CM

• OA cv i—i

• t>- rH

• o H 8

• rH

IT

h- <0 U r-

rH H C^ CV H H H — iSd — S-S3b±S «t

8 o

vO H

O cr\

8 in

O

* 8 -4

C-- (V IT\ ITN l>- r>- CV

I ai3iA -tf -tf U> rH -* -d- c^

z — iSd— C"\ o o

ir\ vO cv

NO 1 C^V o -4

U

<0

SS-3ti±S cv -4" 0 •* o

3-LVWlJLm ir\ «0

cv to vO

to to

ö zz >o Z UO ^

ß O

•ri

ß 5

SH

OT

A

ND

S

PE

C IM

L

OC

AT

I O

N

TE

ST '

n a, H

■P © ü -P © CO

© ©

CO rH

= = 4J © © +3 V «5

fx< 3 P-. ß fo F>-, © fli

ß

H EH EH o E-< EH E-<

""1VIÜ3±VW 3 3 3 CO CO 3 CO ir>

CM CM CM cv CV CV CV

,<0

3UJ o: ?ZüO

1 1 1 1 1 1, T) X •H • •H • •H • *H * •n • ß © O B Ü s © s © S © § © CO 3 CO

rH Sa P cd rH HJ

3 os

>r\ ©

P CO 3 «5

d ^ 5^ZJ co PQ

rH T3

PQ

rH sO © o © -a:

•o o © o ©

IT\ U u> rH N Oi N CV N -4" N -* N z x v O cd O O -H O -H O -H O -H O -H

r- • • . ß • ß • ß • C • ß o O O O o o o

129

rTrTr^rSAH

X) ©

•H

% O O

•o ©

<0 P u

b D «0 <0

 4J •H H

CD ©

Vt ©

C

0*8 u X £E

CO M 1 © a o

p-, •Ö3

o 1

bJ

«-■2 Sa 8 o 8

O IT» 8

t H O < c

© ON PQ

en H

CV H

^1 o p H

m H

hi Q. OH .fi

* 4 cv o

C-rH

H CV O

<0 o CV H

HV A A

V A A A

*- <0 -Ni y3d- cv 1-1 cv cv CV t-\ H» NOIIVONOTJ ir\ o • • • o • 0 •

to % IT» r-i

vO to H

u or en en r-\ en H

•4- I- <o u

-ISd- f- cv

sO «0

8 1- o NO rH NO ITv vO

0131A IA H ir\ H u^ rH

o z -ISd- <3> <3>

CV <2> en

on en o

en

3 en 0 en

SS3U1S- cv

en

rH *

CV to

vO

o rH *

cv

<0 3±vwinn c- en NO en o en

9 zz ><J

SH

OT

A

NC

S

PE

C IM

LO

CA

TI

ON

T

ES

T 03

& H Cr,

r c s c —

E-i O E-« O EH 0

-IVIU31VW 9 cv

CO

cv

CO -* CV

9 CV

U4 CV

CO

cv

3 0 -d

cv t-< vO

P H o

+3 CV © i © 3

1 o O cö • o

cö O © • X O

© • K O A*35

f-i H

« a, g

© P +> bO

>l Ö

•H ©

fn 0 © Ä P

© p © C tw © p © ro

+J cd ctf o ä

•H T3 bO c C

•H •H

>J & Ü ■O O iH • T3 XI T3 ©

© JS -P Ö 0 cd CO

© 2 S ^ a

*H «J O -P ©

cd tt © © 2

O ► 0 0

© © ■O bO C cd & cd *H X! © ß

1 > •H p cd Ä X © to c

•H -H ß ^ cd •H •

p © S © O © © a, •H ft a, 0 2 P "Ä ß T3 9* 2? ♦H © i.s P *H Ü p p © 3 •H

fS p 0 si S O s »

=te

130

CM

l—"O "¥— •p +5 rH H o

tu Wi Q) CD •rl

«0 c ß a s -p •rl tH Tf -o Ü

H p •p V (D CD 0) <D bG W> _l to r-i H O .H O -rl -P bp ß ß a> d) cd *-i cd tu 3 C •H •H D s a «M «H CM -H M -H -P +5 to DC k -d f-l »Ö -P r-i rH

Id < Ü •H

Ü •rl 65 rH

3 -H CO rH

a> DC 2 -p -p O O

Ü Ü bb 0) bO « 0) O Ü Id e 4) ß i ß 1 o 0) •H ^H

h DC p -p •H 0) •H 0) ■P •P to s s T» U <D cd

T3 «^ Ü CD

O CO

cd TJ Cd "Ö T!Ö •P -p h- o o

o 3 H ß s 3

25 ss o cd M to &q U I SS Q. « 3 < Id DC _1 d> m Ü i*: co

o o

K cy

Q o <

18 o o O ^

t-, o o _J .-co

•" o rH O o o 0) o o 8 _J 4> O 0) o en m

cd "E

^ -p

O

Id

< DC a

en CP rH r-i rH

5 CO M

* — Ni U3d- to N0IJ.V0N013 o

c CO 6 ß

0>

H

O

to %

• to

• ON 5 * * 5 t»

H rH o o O o rH

0)

CD 0> CO Id a

CO CO a, CO DC

iH rH rH

to — ISd — 8 O

O 8 Id SS3ÜXS o CV © a> 0) a> -4- 1-

• •* rH rH rH r-i •* ai3iA NO

-4 3 •H 0}

•rl CO •d •H

CO

NO

T § ß 9)

c CÜ

ß CO 1-

O z o~\ <TN +3 -P -P -p cv

— ISd — o O Ö u oc SSBti-LS

CM cv

• O CO

o

to 3±vwi±~in N3 NO O O O o

*—4

o NO

Ö ZZ vO z ug ^

ß o u

c o u

ß o ß a o

u .ÎbZH •" û** oto O ° u

<D CO CO UOH UOrH CO CO H rH rH ß © ß 0> rH r-\ •rl -a:

•rl •H •a! £9 ^3 5!

•rl

(0 to-1 1

cv 1

cv 1 1 ft. 1 Pu 1 1

E-" H E-< H H E-< E-t

■~IVm3-LVW CO

cv CO

3 CO

cv

CO

CV

CO

c> CO

O

CO

C^

«0 1 1 1 1 , 1 1

_Jto 3 d) ffl (D

-a cd CD

X)

9 0> c3 CD

■o

aj <D

5=*z.j ô< o

-P 3 -P 3 P 3 p 3 -p 3 -P 3 P fl COH CO rH CO H CO r-i CO rH CO rH CO r-i

CO. CQ m PQ CQ CO m sO IT\ vO ITN CV CV o

zx u r-lT3 O fn

CV Tj o &

r-i T) CV X) o u O (H St c^-S

• cd • cd • CO • cd • rt • cd • cd o o O o o O O

131

XI

1 •rl -P Ö O o

CM

CO 1-

CO

s 43

cd cf a §ß° C ng a

nd

ltin

g

to ß

* a) •rl •H •H •H •H -P -P s +3

CO CC «J r-\ rH cd H rH rH r-i

 CC-P -P +> -P +> CO 3 ^ Ü Ü ^ o O Ü Xi O u W a> a> <D tt> CD CD CD tjfl CD

H -p -p +3 -p -p -P +> .rl -ä O ß 3 3 ß 3 3 ?> H 3

ss M CrJ w M W W W CO KJ Ü I Q. <

u 08 o _J 8d o L." CD H _1 0 < * o O o o 8 o

.^ O _l H'S

o o rH H

m rH

o rH

in CD O Ü -H 8

< «8 © rH H r-\ rH rH cd ^ rH h- m V( (D

2 a

CO H

* — Nl U3d— CM CM H CO NOIJ.VON013 o O O

ß CD !J %

• o

1 1 B • • s

D CO

-* .n CM rH a 5 u o u O

ff a> ft CD a. a 1- en r^ CO en rH 05 CO

co u r-

— ISd- en to 8 o o O o

ss-3dis o o a> rH

en en CD CD r-{

X ai3iA o rH

o CM

•H 03 ß a>

3 3 •H CO c CD

•H CO

g \~ o en -p -n rH -P 43

z — lS"d — O o o O u 2* o CM O QC s-s^d-LS- to vO o •n

CO 3±vninn rH o en

o S3

CM NO

CM O 3

O

Ö ZZ v^ z ug <" olbzi-

u o a ß ß

O 00 E-*

tf ■p cd 2 E s

ß 2 S»S CD O. rH CD rH

CD rH

^?^f1ocÖ x CLO £ co co-1 ^

e-« H .H M E-<

•H

1

•H ■=■:

i 33

1

o e-< EH

iviusivw CO

3 CO rH

CO in

o CO

E-i CO &5 CO

NO c- CM in CM CM CM

i10

<UJ a:

i 1

■s CO CD

l X»

3 <D 'S cd a>

1 •rl •

3 cd

1

ß CD Cd 43

1 'S ß CD Cd 43

NO

MIN

T

HIC

KN

I A

ND

C

OLO

I -P 3 CO H m

-P 3 CO

-p 3 CO rH

-p 3 OQ

43 iH co i rH in T3 O fc

• 0$

CM

o u • cd

o in CM -d O ^

• cd

in a) in

8"E • cd

NO

• cd O O O o o O O

132

T3 •a n (50 no CO s„ c

cfl § T3 c

•H G •H •S

h- t£ ttu M C b0 b£ -P +3 cd !J co •o c T3 G T) G cd G G H rH +3

D >_ CO -H H -P

cd -H rH -P « 5s*^

•H +3

0) (D o

V) cc O H O rH Ü rH cd rH H Ld <

2 a> a> CD r-i a> Q) O Ü o

DC ^s "ÖS ^-1 S'1

S O

s •H +3

•H -P

43

_ o _, o Ü o Ü Ü TS

tO MMH «l.H bO-H U -H •H Q> 0) a> DC G -P a -p G -P bÜ4J HJ •p ■s

43 •H O •H O •H Ü fn Ü Ü a Ü

u +i a) +5 CD -P a) a) a) a> M 3 1- H +3 1-1 +2 rH -P -P 4J -p -Ö

a) 3 (0 3 0) 3 C p 3 o O G

O s w S3 W g w w 3 ss O o

I (0 a 43

<

o Id

_ co 8§ 8§ 8§

(0 (1) 43

Q aH CM rH rH

CM rH rH O Q * s O _i < !H U (V. ir\ o 5 o 6 o

_i H~S <t> o a> O a> o CM o H O rH O f-,

f ü -H ü .H Ü -H r-\ H d) t» (D C0 <D

^ oc a

cfl J-, cd ^ cd f-. CQ ra 43 «H (D «H a) <n 0

UJ b +» U 43 »H -p p

5 3 c 3 Ö 3 Ö G CO H CO M CO M

CO

«0 1-

-Nl d3d- rH H r-i H rH rH 0)

N0I_VDN0~I3 O rH O

e O t>- O

cd

^ _J D %

• • c*-\

• o

• -d

• • •-ä T3 rH CM CM •H H CM G a> CO O •H 3 u 9> a a: CO

.* •H O -p h

<o iH H r-{ H rH rH o

G O — ISd — 8

to 8 rH

O O O O H JO

C3 u h-

SS"3dlS o vO

0) O CM

IA si

•V •\ •s ■H •> •V ^ o CM

x CH3IA CO ß ^

-o -d

cd

<3 f-- CD +3 CM w O CM H dA CM rH CM O

cq z — iSd — o O o in O O •

< ui r-t O H irs O IfN CD CD E-5 _

l- co

S^dlS O 3

O O E-S 3_LVkNi±nn NO ^ ir\ 8^ x>

ö zz vü 2 uo <^

O CO

0) rH

<

ID rH •H

CD rH H

•H

■Win

g Pa

ne

■Win

g Pa

ne

1 3

SI i a. q h- 1 1 1

c-- 1

O- i •rl

(-1 CD <0 co -1 M

f-t cd (D JH O, CD

~IVI_3i.VW CO CO E-i CO CO CO CO

EH CO

P. > 3 cd

-4- -d -4- ^f -4- ir\ ir\ CM CM CM CM CM t>- i>- G

•H a> 0)

.£ | •p -p • • IH X) •H •H ■"Ö «> •o o (1)

< _ DC ZZQo

G a) cd -P « cd § ^ 1 s z 3 r-t ^

-P -H co £

Q) • Ä -P

0) • P3 +3

cd 3 fS a £ cm

en o CO O 3

NO

MI

HIC

K

AN

C

OU

CM

SO V

(a ^> -r) rH 03

CO ir\ .H CM CO

MD 'd rH 3

•rl LA a CM 3

.032

i

band

.

O « -Ö -d +> G

* O ft

• cfl O CD o a)

• a! O rH °.3 O .H cd

• Ä 43 1- o o ° 1 o o OStO o S co _________

133

•**s.:

*^_ .'-.,. —•'•">•"/

.'• " y'"5»*^

•-**

(a) Magnification: 150X

Description: Etched microstructure of sample heated below 900#F.

'1 c ^f'':~~ ?■' .' /i''

.«•,_e - •>„ ^v--- ..'«

'*"•£ •'.

(b) Magnification: 150X

Description: Etched microstructure of submitted sample showing evidence of surface melting.

Fig. A.l - Photomicrographs Showing Unaffected and Slightly- Overheated Aluminum Alloy Sheets

134

v \

"ZTt ~^PF~ .,-■'•< j

o

>%■

• r f ' ' V +

(a) Magnification: 250

Description: Etched micrestructure showing the effect of melting of base metal and unmelted cladding.

■s » '- «^V

. *K'

* *» r-y r-*'.' ■"» v**?.: ^v>;^*

- ^,v > ->•? £ ■. •■■ ~'r'-\. ■> V. v-v-^-^'frf ..'■- ■

»*-v. ■■-v.- ■-«. i""*- '-■- * *""■-- /.It»' ">A»-c.

**.

t^*l H^> -\4^!* - f

(b) Magnification: 250X

Description: Etched raicrostructure showing the effect of melting .of both the cladding and base metal.

Fig. A.2 - Photomicrographs Showing the Effect of Overheating 24S Aluminum Alloy Sheet

135

APPENDIX B

EFFECTS OF THERMAL RADIATION ON THIN ALUMINUM ALLOY PANELS EXPOSED ON THE

GROUND TO AN ATOMIC EXPLOSION

by

M. BENNON

NAVAL AIR MATERIAL CENTEä

PHILADELPHIA 12, PENNSYLVANIA

136

APPENDIX B

EFFECTS OF THERMAL RADIATION ON THIN ALUMINUM ALLOY PANELS EXPOSED ON THE

GROUND TO AN ATOMIC EXPLOSION

B.l OBJECTIVE

The objectives of these tests were: a. to determine the values of the equivalent thermal absorption

coefficients for aluminum sheet metal panels of different thicknesses coated with various aircraft lacquers when exposed on the ground to thermal radiation from an atomic explosion.

b. to determine the magnitude of structural strength changes in sheet metal panels prepared to simulate aircraft skins after exposure to radiation from an atomic weapon.

B.2 BACKGROUND

The tests from which the reported data were obtained were made at the request and with the cooperation of the Bureau of Aeronautics to obtain data supplementary to that from the drone flight test program during UPSHOT-KNOTHOLE. The circumstances under which these supplementary data were to be obtained required that only passive types of thermal instrumentation be employed. All test material was exposed during Shot 9.

A significant variable for assessing the damage to aircraft skin structure resulting from exposure to radiation of an atomic explosion is the peak temperature reached by the exposed sheet metal. This peak temperature for a given level of exposure is a function of surface condition, material composition and skin thickness. For convenience in assessing the effect of surface condition on peak temperature attained a coefficient of equivalent thermal absorptivity has been defined as follows:

137

WCAI1 * QA

u. (equivalent thermal absorptivity coefficient) = fraction e of QA absorbed by metal to produce temperature rise A T

W = density of material

C = specific heat of metal

AT= maximum temperature rise of metal during exposure

t = skin thickness

Q = total radiant energy normal to unit area of skin A surface

Bö T5ST DESCRIPTION

Three frames to which were attached 35 24 ST aluminum alloy panels were placed at each of three stations at various slant ranges from and with the faces of the panels normal to the intended burst point of Shot 9. Each panel was hinged along its upper edge to permit free rotation except for the frictional restraint offered by a sheet metal clip bearing on the stiffener attached to the lower edge of the panel. This frictional restraint was intended to initially^ maintain the orientation of the panel with respect to the burst point but not offer appreciable restraint against the blast pressure wave. The panels were of various thicknesses and four types of surface preparation. The surfaces of one group of panels were in the bare "as received" condition. The other three groups received a conventional wash primer coating followed by a 2 mil coat of Military Specification MIL-L-7178 cellulose nitrate lacquer on the surface to be exposed. One of these groups received a glossy sea blue (color #623) coat. A second group received a white (color #511 with omis- sion of the blue tint) coat. A third group received an aluminized coat. On the back close to the center of each panel was attached a Temp Tape temperature indicator. Each of the three frames carried along its upper edge a set of passive type thermal energy indicators. Figures B.l and B.2 show the details of the panel and frame construction. Figures B.5, B.7, 3.8, B.9 and B.10 show the thermal energy indicators on their mounting strip, a frame with mounted panels and thermal energy indicators, an illustration of a Temp Tape installation, a frame with the temporary protective covering (removed before the test exposure), and a frame set up at the test site with its supporting braces.

In addition to the three frames, four AD airplane elevator sections were exposed at each of two of the three frame stations. The orientations of the elevator sections are shown in Figure B.50. Each elevator section carried either two or three Temp Tapes on the inside skin surface of the exposed side.

138

Each panel carried an identifying number. The arrangement of panels in the frames is given in Figure B.3.

The actual burst point differed somewhat from the intended; however, the cosine of the angle between the lines from intended burst point and actual burst point to a frame in all cases exceeded 0.97. Slant ranges from actual burst point to the frames were as follows:

Station 1 (Frame 1) - 7150 ft Station 2 (Frame 2) - 5330 ft Station 3 (Frame 3) - 3800 ft

The AD airplane elevator sections were located at Stations 1 and 2. After exposure tensile test specimens (Figure B.4) were cut from

the panels and elevator sections and yield and ultimate strength data measured.

Locations and numbering of the elevator tensile specimens and Temp Tapes are given on Figure B.51. Specimens were taken in pairs from directly opposite sides of the section, one from the exposed and one from the unexposed side. An "E" is added to the identifying number for the specimen taken from the exposed side.

B.4 INSTRUMMTATION

B.4.1 Temperature Indicators

Peak temperatures reached by the sheet metal surfaces of the test specimens were recorded by Temp Tapes. These devices were developed by the University of Dayton under contract with the Wright Air Development Center. They consist of a fiberglas fabric coated on one side with a pressure sensitive adhesive to which are attached an array of 24 temperature sensitive discs each approximately l/ö in. in diam- eter. The discs are of two types, the four to indicate the highest temperatures are thin low-melting-point metals, the melting of which indicates that certain temperatures have been exceeded. The other sensitive discs are of the material developed by the Quartermaster Research Laboratories (see ref(19)). They consist of a black absor- bent paper base coated with white organic compounds of suitable melting points. When the coated surface of the paper reaches the melting point of the coating, the coating melts and is absorbed by the paper making the surface appear darker. ■ Figure B.ll shows the adhesive coated side of a Temp Tape and the temperature indicating discs. The Temp Tapes are fastened by means of the pressure sensitive adhesive to the desired surface. Calibrations made by the Naval Materials Laboratory were employed in interpreting the Temp Tape data. The highest temperature indicated on the Temp Tape is 640°F. Peak temperature indicated was generally taken as the average between the highest positive melting temperature indicated and the lowest positive non-melting temperature indicated. Several of the indicator discs which were found to be unreliable were neglected in reading the data.

B.4.2 Thermal Energy Indicators

The thermal energy indicators consisted of mounted cotton and

139

wool fabrics and maple-wood blocks. The degree of destruction or charring of the material when compared to the degree of charring under known conditions of irradiation provided an indication of the amount of radiant energy to which the indicators were exposed. The materials were provided by the Naval Materials Laboratory which also evaluated the resulting data.

B.5 RESULTS

Test data on the panels are presented in Tables B.l, B.2 and B.3» Ultimate and yield strength of unexposed test specimens taken from the same 24ST aluminum alloy sheets from which the exposed panels were fabricated are presented in Table B.4. Photographs of the panels in the three frames after the test exposure are shown in Figures B.12, B.13 and B.14. Individual photographs of the 35 panels after the test exposure are shown in Figures B.15 to B.49, inclusive. As shown in the photographs, some of the specimens were distorted or damaged as a result of their striking the back of the mounting cage at time of shock arrival.

Test results on the type AD airplane elevator sections are presented in Tables B.5 and B.6. Photographs of both the exposed and unexposed sides of the individual elevator sections are shown in Figures B.52 to 3.67, inclusive. Data for the equivalent absorptivity coefficient against peak temperature attained for each type of surface are presented in Figures B.68 to B.71. Data for the equivalent thermal absorptivity coefficient vs panel thickness for sea blue lacquered surfaces are plotted in Figure B.72.

Radiant energy at the three stations are presented in Table B.7. Values are given as indicated by the passive indicators mounted on the frames. In addition values extrapolated from measurements made by the Naval Materials Laboratory along the line of test frames utilizing metal foils are presented. The Naval Materials Laboratory data were used for the computations in this report as they were considered to represent more accurately the conditions at the various panel stations. In the case of Station 3 no value was available from the frame indicators.

The scatter in the thermal data obtained (Figures B.68, B.69, B.70, B.71, and B.72) obscures many of the trends which would probably otherwise be more distinct. Two such relationships are the expected increase in equivalent absorptivity coefficient with increasing sheet metal thickness and with declining total thermal energy. For the sea blue lacquered panels sufficient data exist to permit plotting the data by stations to demonstrate the trend (Figure B.72). A larger number of panels and temperature indicators having a more closely spaced set of temperature indications would have afforded more satisfactory data. Some uncertainty exists as to the rapidity of response of the temperature indicators, any lags in response would tend to make the peak temperature readings somewhat lower than the true peak for short temperature pulses.

HO

In frame 1, panels Ö and 10, although of the same gage metal and having similar coatings, experienced markedly different peak temperatures, the plate closer to ground reaching the higher temperature. An explanation for this behavior is not presently available.

B.6 CONCLUSIONS

Based on test results, the following conclusions are made: a. As indicated in Figures B.68 through B.71, the limited data

obtained show considerable scatter and accurate determination of the values of equivalent thermal absorptivity coefficients cannot be made. Since deterioration of the lacquer surface has occurred in some of the panels due to charring or flaming or the presence of a dust film prior to thermal exposure, the coefficients obtained in these cases may not be representative of the original type surface.

b. The expected increase of equivalent thermal absorptivity coefficient with increasing panel thickness and decreasing thermal irradiation is indicated by the set of data plotted on Figure B.72.

c. As indicated in Tables B.l through B.4, tests conducted on exposed and unexposed specimens of the 24ST panel material show, in general, only a slight reduction in strength properties due to the heating cycle. In a few cases melting of the panels has occurred as shown in Figures B.39 and B.42.

B.7 RECCttfrMflJATIONS

It is recommended that: a. Further measurements and associated analytic studies be made

to obtain a detailed quantitative picture of the factors influencing the peak temperature rise in coated thin metal panels.

b. Where practicable in future field test programs, attempts be madeto obtain temperature-time histories on similar test panels.

c. Attempts be made to improve the existing Temp Tape type of device to provide more reliable indicators in place of those which have proven unsatisfactory, with a possible increase in both temperature range and number of temperature indications, such instrumentation to be used where a large number of points are to be instrumented with minimum effort or where tine-history recording equipment cannot be employed.

d. That data be obtained to permit the dynamic error in peak temperature data obtained from Temp Tapes to be evaluated.

241

142

X (0 UJ

u a. n-T-

ILJ. -II n-

1 I -i'—■ ^ ■

* X h- o

z c* o -—*"• _l X

O H z m a:

Ö o •H

-p W 0 o o

aq

bo •H

?9tr-

143

PANEL- 10 SEA BLUE LACQUER 0.016 IN.

PANEL-3 WHITE LACQUER 0.012 IN.

PANEL-2 ALUMINIZED LACQUER 0.012 IN.

PANEL-I BARE METAL 0.012 IN.

PANEL-17 SEA BLUE 0.020 IN.

PANEL-15 WHITE 0.020 IN.

PANEL-13 ALUMINIZED 0.020 IN.

PANEL-11 BARE 0.020 IN.

PANEL-23 SEA BLUE 0.02 5 IN.

PANEL-26 SEA BLUE 0.032 JN,


PANEL-& SEA BLUE 0.016 IN.

FRAME NO. I













FRAME NO. 2



PANEL- WHITE 0.020 IN.

6



PANEL-I 4 ALUMINIZED 0.020 IN.


PANEL-2 8 SEA BLUE 0.032 IN.

FRAME NO.3




Fig. B.3 Arrangement of Panels in Frames

LU

I <D

<D

&

P. •H

•P CO

o 4

u ID

<D

w

•H to

U5

w*.. - î-., 'i-, ^*«äf

IJJ ##g#ÄwÄH'--

CO

u ü

+3

eö

CO u o t3

C< CO

vO * m

fr.

| '

BKisÄPS

U6

^

• ■'■ tu

>-:h i! 1

* WÜr* ^^ ■

Fig. B.7 Frame with Mounted Panels and Thermal Energy Indicators

U7

■■■■■■■■■■■■■a ■■■■■■■■■■■■■a ■■■■■■■■■■■■■a ■■■■■■■■■■■■■a ■■■■■■■■■■■■■a iiiiiiiiiiiii ■■■■■■■■■mi IIIIIIIIIIIII ■■■■■■■■■■■■a ■■■■■■■■■■■■■ ■■■■■■■■■■aaa ■■imiiiiiii ■■■■iiiiiiiii ■■■iiiiiiiiii ■■■■■■■■■■■■I ■IIIIIIIIIIIII »■■■■■■■■■■in

i m

ii \

!!

3!

Fig. B.8 Mounted Panel Shov/ing Temp Tape Installation

MB

Fig. B.9 Frame with Temporary Protective Covering

Fig. B.10 Frame at Test Site before Exposure

H9

WS

INCHES

Fig. B.ll Temp Tape

Fig. B.12 Frame No. 1 after Exposure - Slant Range 7150 ft,Thermal I Energy 18.3 c&l/ca?-

150

CJ

CÖ CO H O CO

O I Cî

CD >> U 60

CO CD

* I -P 0)

ctj E-i

CV •>

U £ xi o

<D <D S tip CD t-t U cö

PQ

•H

151

Fig. B.14 Frame Number 3 after Exposure - Slant _ Range 3800 Ft, Thermal Energy 43 cal/cra

152

Ö o • •H

ß p M trt

-P (V ro H O 1 • o U

(1) 1 "i

0< 

a $H (!) a> -p u «Hl S H i

H X Q) o

3 p- H

ir\ H • m

• b£

•H (^

153

jälm

• C O

M •H P

sü m H -p o CO • o i

1 ?-. a>

CD

m cn ►-J

N CD W 3

H U (JÜ Q) P rt tw (1) cö CO

to 1

H -M 0) u C •H cr< X

P-i H

to r-{ • cq

• feO •n fr-,

• Ö H M

Ö CN; ü H •H o P • ctl O -P

1 CO

CD 1

U U 2 CD 0) d, o a* CL, o (3 & Vi CD 0) -P -P •rl CM ,c tö

c^ 1

H X (1) a

s •H

£L. E-i

r- H » 01

« «! •r) fe

154

V

\

V , „ „, 1..

c H

O CM iH O • Ö o O

•H 1 -P

CO <D P S-. en 3 (D i

& H

^ crt P CD

^ a CD P 0) <H SH

CO &

H H 1

H ^ a> Ü

§ •H

P-, H

o Cxi • P3

• to

•H (x.

H • Ü ß H o

•H vO p H CO O P • CO o 1

1 u Q) CO ?* 3 3 a1

co o o 3 & 0)

2 u rH 0 PQ p

S cd a> en

o r-i I

H ^ 0) o Ö •H cd Ä

&< fH

CN H • pq

. t»0

•H Ü.

155

V

Ö M H

O C cv Ü o •H • -P ü a?

+3 i co

a) 1 ^ 2 u m (1)

-f EL, a1

W^ C) w crt

t-1 u 0) a) -P +3 <M •H it

ir. H 1

H .* 0) o C •H crt £ eu H

cv CM • CP

• bn

•H fc

c • o Ö •rl M -P

crt O +3 rv rn O • I o u

1 (i) 3

<D CT< ^ o 2 ,3 s, T) 5^ a> W N

•H k C a> •H -p «HI

e crt H

«C

a 1

H ^ a) o

SÜ •H

P* IH

H cv • m

• &c

■H fe

156

■A

M O •H

IT\ -p CV 03 O +J

I I

0) 0)

3 & CO Ü O cti

CD CO -P

erf a)

CV I

H .Jci a> ü ß -rl CO .fi

0L< H

3

157

Ö C M O

•H O P CM crt O • £ o

I 1

S-. <D a> Ü 3 en o B.S £ a)

■3 u H 0) CU -p «H m cd Q> u A 3 oH

w O

% CO

ä a) 3 u H

0> w -P <H crt erf (11

en <r\ 1

H ^ (1) o § a. H

vO Oi • CP

• bf

ti

rH

c Ö H O

•H C\J -P r^\ CTl O ■P • ro O

1 1

fc <l) (I) u "t 3 oH

m o o crt &hJ

£ a> 3

?H H <» W •P «H <rt crt (1)

rn sü cv l r-H ^

(1) O

g •H Xi

P-. EH

IT» cv • PQ

• fcf

•H Cn

158

cv Ö o •

•-P1

H «1 -p

v0 en H O I • O u a>

1 -i CT1

a> O u 2 3 m o T> 9< (1)

& Ö

u •r) Q> pl

■P 3

3 MD i

H ^ (1) Ü fi ä ctl a, n

<o CM

GQ

W) •H &4

vO CV H O ti * o o •H

-P 1 crt

-p Q) co

& I tn o ■d ä -p

u (1) « -p ^ 13(58 »A i

H .* a> O

s •H

fit H

r- cv • ra

txo •H Cn

159

cv

\

• o

M •H -P

vO «1 H ■P O </) • o 1

1 k (1)

0) - O H •H O •P • «id o -P

1 en

+ o £ CO

(H 3) 0) •P -p •rt

§ £> i

H ^J a> Ü ö •iH CO £1

PH E-i

o cv • en

hj) •H &4

160

ITN CV CV O • Ö o o

•H 1 P

at CO P U »•3 2 CO 1 o Q< H

is CO P CD

u s «J -p (0

^ 1H

w o CV 1

H ■M U> f)

ä p^ H

CV en • m .

K) •r4 ÜH

CV • Ö C H O

■H O P CV ert o P • tn o

i 1

fn CO CD k 3 3 o' to Ü o ai Ut t-l

iS4 d) pi ^ H

0) UJ -P «H tf m (U

(/) «0 H 1

H ^ (P o

s •H

P-I IH

H m • CQ

• W)

•H fn

161

Oi

IA C CV O Ü •H • P Ü CO

P 1 CO

CD 1 h d IH CO a> a & ä o

crt M u

CD cu P -P «H •H cd fS

CM CM 1

H .JU CD O

Sä PH EH

-* en • CQ

• bO

s

CV Ö • Ü

C iH M p

cd lf\ +5 (V CO O • i o u

1 CD 3

<D a1

U o 2 CO m H-l O (**>t Tl Uj 0)

PH N •H

P c a) •H p Ö

% H «J

H CM 1

H .« CD o C •H cö Xi (U EH

OA c^ • CQ

• W) •H fe

162

cv

/ •

f

3S •rl

CV -P tn CO O -P

• CO o 1 ,x\ 1

1 u

d> Q)

&& ra Ü

..*'■

Q CO

* $% a <D 2 ; *,.. •*> f* rH

' -«%', r v © cq v;v:, »' y -P «H CO

" /'' C6 <D CO

t> <M 1

d •* a> o

cd Jd a. e-<

vO en * m

60

CM

a c H o

•rl lf\ 4J CV cri O -P • ro o

i 1

?H a> (1) h 3 3 a1

m ü

& cö

ä CD

JH rH a> m

-P

S CO

m -* cv i

H AS d) o Ü •H CO Ä P* E-i

IA m • CQ

• 00

El

163

cv

A

Ö c M o

•H H 4-> ir\ * o -P

• f.'l o

1

u (I) (1) u 0 3 Ü m r) o erf O.f-1

ä <D p

fn H CD ixi P <H> CTt

crt a> CO

UA f^ i

H x 0> o

% OH E-i

to c> • CQ

• w £

Oi

M0

Jf**1

V «i-A

ft '■■

c C H o

•H Ü -P -tf at Ü P • ro O

1 1

h a> <1) u 3 2 C* V) o o m ft(-a £ <ü

3 u rH o c •H m FM

PH E-+

£> o^ • ff)

• to

•H fc

164

en

Ä.

*% ,

#

C • o fi •H M 4->

01 O -P (V m o • i o

fn 1 a>

0 (1) cr u Ü

$3 o PHTJ

£ CO N

•H

0) .3 p ö «M 3 01 H

«Ü -4- H i

H X u> o C •H * .C

IX. En

O -* • d)

• bO

•H 1^

• fi H

O (V m (.)

m c o o

•H i -P

oi <D P

S en (0 1 O P* H Q ni

fxj p <n

h r5* a> P Cl> «H U 01 0)

:u Pi i

H ,* (1) o

s •H

PH EH

cr^ en • ra

* bO

•H li,

165

OA

#? *

Ö Ö M o

H O P (V IT) O +3 • 7) o

1 1

U a> a> u PV

Ü O o trt ft^ a a)

P J- H (D W

+5 «M crt til <D

W 0-> i-i 1

H -M a» o rt ■H nJ JÜ

H< bH

cv -d- • CO

• to

•H fc,

• c M C\ O Ö fV o o •H • ■P o CO

-P 1 C/J

<1) 1 ?H -( ^ in a> o ■± R< a* £ o

cfl t-1

k 0) CD -p -P <M •H «5 jg

NO H 1

rH « a> U

S PL, E-i

H -* • cq

• b£

•H U,

166

c^ ß c M o

•H CM P C"\ cö o -P • ro L>

1 1

fc 0) a> %4 ~s 2 o* UJ O O af

* .-1

u rH 0) w p «H m

(13 <D rn

G) cv 1

H ^ a> o a •H cö ,c G* E-«

3 • CQ

• U)

£

c^v

M O ■H

ir\ +J Q* crt O -P • en O

1 1

U 0) a> u g. 2 a1

to o

& a!

£ a) 3

u H a> tq -p <+-! <0 01 o c •H * X!

PH &H

o~\ -tf • CQ

• tjfl

•H f*.

167

c M

H ir\ O O • d O O

•H 1 -P

n) a> -P u CO 2 tn 1 o 4" ■a

-p (1)

u s OJ -p 1)

S 51 H en 1

H j« (I> ü

s •H

A« H

sO -4" • CO

• bO

•H fo

ftf.

en ö d M O -. *H O -P

O +3

o I

1 u

u &

& 0) pj ^ rH

0) m +> <i-< cB crt a)

01 O en 1

H X 4> o d •H

P-. EH

UN

168

1»;

/

ti H en rH ß lf\ O O •H • -P O oi

-P 1 C/J

2 1

S & o ri u cr S ü

01 t-i u

ID a> -P -P

S ^ en C\ l

<1> •s s •H

(X, fri

3 • m

• 00

ti

CA

Ö • o c •H H •P

01 H -p lA CO O • 1 o

fn 1 a)

•2 0 a1

& o

to f-J

& T)

ä •H

k Ö 0> -P P *H P a 3 C\i en 1

■H ^! 0) O c •H cö Ä

&4 EH

r- —a • CQ .

bO •H fc.

169

\.

Fig. B.49 Panel 36 after Exposure - 0.051 In. Thick - Sea Blue Lacquer - Station 3

170

171

o

<0 co

Ul < 1 2 CO o Ul i.i u. CL < (0 1-

K y CO oc Ul D h- H

< u rr _l Ul CO Q. Z 2 ui Ul h- h-

8 n u

r~: t=xn —\\

8 : ca j

< |

i

i

1 i

i 1 " 1 1

< —

8

H

OQ

EDO !■ '

. 1 i

JZDZJ

i—i j 1 1 1 t

<

R r H I- i—i

..J ..-i _

EDO '

tzxn |

i—i

CO

CO z o

co

(0 CO z o I-

CO

(0

CM CO

CO

in

cO z o

a co

a E-<

P. s 0)

■s CO

03 c e

•rt o 8.

CO

-p CO CD

EH

«H

O

CO C o

•r-f

o o

(-1

bC ß

•H

o CO

CO ß o

•H -P o Q>

CO

o

I H

PQ

&0 •H fa

172

s

% +> O (0

CO

§ •H

IS o

1 ■ ÖS

•H

ISP

ÜfP'l^-

in Ö CNJ o

•H Ö -p o o •H <D ■P

CO d -P

(4 CO O -P 1 crt > Q) a). Tl H •H W CO

C\f IA • ra

• W)

•rH tx.

173

r • o

•H ■P O

CO cti •P

^ CO

3' a>

W CO

Ö O

•H

-p CO

pq

M •H

T3

O

c^v e H o

•H G P o Ü •H <1> -P

CO CO P

(4 CO o p 1 crt > 0) <n Tj

d •H CO

vO ir\ • fp

• W)

£

175

"8 CO

a X

ä

O -H m to

3 . cd > <D

«A

PQ

b0

5!

<D (0

•H C -P O ü -H (D -P

en eg -p

fn en o 3 ' >• © (D T3 H -H P4 CO

■CO

PQ

176

'-1 •g 0) o «• (I) Ö

J=>

tf\ C CM O

•H C +5 O Ü •H <D -p CO eri

-P fn en O

+3 i cd > a) r) H •H W CO

if?

AT , ITft,

d> to o

iA

c r\? o

•H ö -P o Ü •H V P

CO cd -P

SH CO O

-P 1 cti

* •rt d •H

CO

o vO • m

• W)

£

177

L

<D CO O

9* S

vO

o •H c -P Ü O •H U) -P

CO «J +5

u CO o -p 1 cd > d) to 'd

ä •rl CO

r^ vO • CQ

• b£

•H U<

vO

1

Ö cv o

■H c -P o c> •H 0) -P m CO

P u CM o -p 1 crt t> a> T)

iH •H w co

o; vO • pa

• tu':

•H fc

178

T3 CO CO O

& 0 Ü

179

a H o •H C •P O o •H Q) -P

co Ct) •P

fn to O

-P i a) > co (l) T)

ä •H CO

ir\ vO • CO

• M

•H fr.

.'"=?•;

CO

I fi rH O

■H C -P ü O •H a) -P co cd

•P u CO

-2 1 crt

£ CO T)

d ■H co

3 • PQ

• W)

•H fl4

CQ

«C

Ö r-i o

•H c •P ü O •H (!) P ro cd

P h C'i o -p 1 ert J> 0 a> n H •H W CO

sO NÜ • cq ,

w> •H fn

180

t.

<£ en (0 (0 UJ UJ o Ul T m m 11 u z U 1 U

o >«: - 1 ^ <J o 2 n rviQ —

— <Mojfn"<ttn H PQ

UJ OOÖOOO (d ö c! ö ö 6 6 <D

z i t i i i i CO

£ 0<X€>BD I

C

El X ® O CO P

p O -4-> •<

k CD u 2 p CO

Id U Z 0)

p. ® 2

b

S 0)

En

< •Ü X

Ld CO Q>

PH

co >

B 1 *l °2 0> *<J •H *a. Ü

eP 2 Ul r-

_l Ul z

•H «H <H 0) o o

11

< > ct •H P

*: < o

[ 3 co

«3<

<° H

f* 0>

£ P c Q) H cö >

•H 3 O« W

O i i

tO iD Tt (f) CVl — C > o o 6 ° c )

• bo

>Y 1N3DL U30D AHA! idyosov l N3" ivAino: \ t*4

181

d ö

*

ö

(O u I— o z

<n Lü

o z rvj«ooin— CVl(\JiO oqqqq öciöoo

• iiii 9Ho<KD

o o in

o o

o o (0

u. 0

Q U z < h- I- <

ui Q:

DC Ul D_

_J Ul Z

< Ul

o o CM

3V 1N3IDUJ300 A±lAlldy090V lN3nVAinD3

182

u

& O CO

CD -P •H

73 CD C

•H CO -P -P

© u P

a I CD

j« CO CD a,

CD •H O

•H «H

o >J

•p

s p p« u o CO

CO e

x: H

■s CD H cd

.5 W

PQ

bO •H

4—h4- o ö

CO Ö

-h-\-

$ (0 # WJ UJ UJ UJ Z

INC

I-

INC

I

X (\J<DOif) — h- (\lc\|tn _J oqooo UJ ooooo z <

■ i i i i #o<ixa Q.

o o <0

o o m

o o

u. o

Q UJ Z

< ui a:

? a: UJ Q.

UJ

_J UJ z

< UJ Q.

O O fO

SV 1N3DUJ3CO AllAlldäOS9V lN31VAinr>3

183

a)

<D X EH

1 H a)

o

PQ

en <n ÜJ z U X

_J Ul z

(0 Ul X« U z

(/) Ul

-X o z

(MoJ o ooo dodo

H «1 •P <D S a> fc A

o o 1 (0

0) u. c

•H 0 cö

-p n +3

ui <c

z a)

< Ö h- -p

1- CO

< 0)

p* Ul 0 t£ EH

o D .M o 5 cfl

CD in (T PH

Ul CO

OL > 2 Ul a) H •H

Ü

_l •H

Ul «H

Z 0)

< o OL

>» £ -p

•H o < > o Ul -P

* u. a, o CO

<

cfl e 0) x: E-<

■s a)

> o •d o a*

d fO d

CM

d d

CQ

3V-1N3DIJJ30D AllAIidaOSSV lN3TVAin03 «JO

•H

184

* V 1N3IDUJ30D AllAlldäOSQV iN3nVAinÖ3

135

TABLE B.I FRAME NO. I — PANEL TEST DATA THERMAL ENERGY-18.3 CALORIES PER CM'

PA

NE

L

NU

MB

ER

ME

TA

L

TH

ICK

NE

SS

IN

CH

ES

u CC D <0

PE

AK

TE

MP

. R

ISE

°F

A

BO

VE

6

2°

F A

MB

IEN

T

EQ

UIV

AL

EN

T

AB

SO

RP

TIV

ITY

C

OE

FF

ICIE

NT

^ e

UL

TIM

AT

E

ST

RE

NG

TH

-P

Si-

YIE

LD

S

TR

EN

GT

H

— P

SI

—

1 0.012 BARE METAL

406 0235 64,900 47,300

2 0012 ALUMINIZED LACOUER

530 0.306 65,700 48,700

3 0012 WHITE LACQUER

217 0.125 68,200 50,400

8 0.016 SEA BLUE LACQUER

>578 65,100 52,900

10 0-016 SEA BLUE LACQUER

456 0352 66,200 48,700

II 0.020 BARE METAL

317 0.306 67,500 45,200

13 0O20 ALUMINIZED

LACQUER 288 0278 66,900 43,900

15 0020 WHITE

LACQUER 154 0.148 67,500 45,100

17 0O20 SEA BLUE LACQUER

337 0 325 67,100 46,700

23 0025 SEA BLUE LACQUER 381 0.459 66,800 45,500


315 0485 66,400 43,400


244 0600 70,200 47400

186

TABLE B. 2 FRAME NO. 2 PANEL TEST DATA THERMAL ENERGY —30 CALORIES PER CM

PA

NE

L

NU

MB

ER

ME

TA

L T

HIC

KN

ES

S

INC

HE

S u

u

a. D V)

PE

AK

T

EM

P

RIS

E

°F

AB

OV

E 6

2°

F A

MB

IEN

T

EQ

UIV

ALE

NT

A

BS

OR

PT

IVIT

Y

CO

EF

FIC

IEN

T

UL

TIM

AT

E

ST

RE

NG

TH

—

PS

I-

YIE

LD

S

TR

EN

GT

H

— P

SI—

5 0-016 BARE METAL 530 0.249 64,500 51,300

6 0016 ALUMINIZED LACQUER

337 0158 67,600 51,800

7 0.016 WHITE LACQUER

406 0.191 63,400 45,500

9 0.016 SEA BLUE LACQUER 469 0220 59700 37,000

18 0020 SEA BLUE LACQUER

530 0.312 66,500 49,600

20 0025 BARE METAL 530 0389 66,600 50,400

21 0025 ALUMINIZED

LACQUER 406 0298 64,600 44,900

22 0.025 WHITE LACQUER 211 0-155 67,900 47,000

24 0025 SEA BLUE LACQUER 469 0344 69,300 47,500

27 0.032 SEA BLUE LACQUER 406 0-312 65,300 44,300

29 0.040 SEA BLUE LACQUER 340 0400 65,500 42,400

35 0.051 SEA BLUE LACQUER 317 0-475

68,800 45,400

187

TABLE B. 3 FRAME N03-PANEL TEST DATA THERMAL ENERGY-43 CALORIES PER CM v

PA

NE

L

NU

MB

ER

ME

TA

L

TH

ICK

NE

SS

IN

CH

ES

id

D <0

PE

AK

TE

MP

R

ISE

°F

AB

OV

E

62

°F

AM

BIE

NT

EQ

UIV

AL

EN

T

AB

SO

RP

TIV

ITY

C

OE

FF

ICIE

NT

UL

TIM

AT

E

ST

RE

NG

TH

—

PS

I —

YIE

LD

S

TR

EN

GT

H

— P

SI —

12 0.020 BARE

METAL TEMP TAPE

MISSING

14 0.020 ALUMINIZED LACQUER 481 0.197 64,400 44,800

16 0.020 WHITE

LACQUER TEMP TAPE

ILLEGIBLE 65,500 47,600


TEMP TAPE

MISSING 57,800 36,400

25 0.025 SEA BLUE LACQUER >578 59,700 39,700


469 0308 57,500 38,600


469 0.384 62,700 48,600

31 0051 BARE METAL

>578 66,300 53,400

32 0051 ALUMINIZED

LACQUER 340 0.356 68,400 45,200

33 0.051 WHITE LACQUER

340 0-356 69,700 46,700

36 O05I SEA BLUE LACQUER

456 0478 70,500 50,000

188

TABLE B.4

ULTIMATE STRENGTH AND YIELD STRENGTH OF UNEXPOSED 24 ST ALUMINUM ALLOY PANEL MATERIAL

METAL THICKNESS

INCHES

ULTIMATE STRENGTH

P.S.I.

YIELD STRENGTH

RS.I.

0.012 67,500 49,800

0.016 66,400 49,100

0020 65,200 44,200

0025 69,300 47,200

0032 69,200 4^800

0.040 68,800 49,000

0.051 70,400 49,700

189

TABLE B5

ELEVATOR SECTION DATA - STATION I 24STALCLAD- SLANT RANGE-7I50FT AMBIENT TEMP 62° F

ELEVATOR SECTION #3 "A" TEMP TAPE-403°F "B"TEMP TAPE-592° F

TEST SPECIMEN NO

SECTION SIDE

METAL THK IN

ULTIMATE STRENGTH PSI

YIELD STRENGTH PSI

1 UNEXPOSED 0.016 61,500 44,700

IE EXPOSED 0.016 61,000 46,500

2 UNEXPOSED 0.016 61,300 45,600

2E EXPOSED 0.016 58,800 44,000

3 UNEXPOSED 0.032 67,900 52,500

3E EXPOSED 0.032 64,000 46,000

4 UNEXPOSED 0.032 64,600 NO DATA

4E EXPOSED 0.032 66,700 NO DATA

ELEVATOR SECTION** ., „ "A'TEMP TAPE-I80°F "B"TEMP TAPE-351 F C TEMP TAPE-235F

TEST SPECIMEN NO

SECTION SIDE

METAL THK IN


YIELD STRENGTH PSI

1 UNEXPOSED 0.016 62,300 47,000

IE EXPOSED 0.016 62,700 46,200

2 UNEXPOSED 0.016 61,700 45,900

2E EXPOSED 0.016 63,100 47,300

3 * UNEXPOSED 0.040 28,000 15,900

3E* EXPOSED 0.040 33, 500 16,700

4 UNEXPOSED 0.032 66,200 48,300

4E EXPOSED 0-032 67,500 49,700

* 52-SO ALUMINUM ALLOY

190

TABLE B5 (CONTINUED)

ELEVATOR SECTION4*?

VTEMP TAPE-226°F *B"TEMP TAPE-35I°F

TEST SPECIMEN NO

SECTION SIDE

METAL THK IN


YIELD STRENGTH PSI

1 UNEXPOSED 0.016 62,600 44,700 1 E EXPOSED 0.016 58,300 41,800 2 UNEXPOSED 0.016 62,400 4 5,300 2E EXPOSED 0.016 58,400 43,100 3 UNEXPOSED 0.032 70,200 56,300 3 E EXPOSED 0.033 67,600 52,300 4 UNEXPOSED 0.032 63,300 37 , 700 4E EXPOSED 0.032 64,200 49, 800

ELEVATOR SECTION*^

VTEMP TAPE-306*F V TEMP TAPE-NO DATA V TEMP TAPE ~2 79° F

TEST SPECIMEN NO

SECTION SIDE

METAL THK IN


YIELD STRENGTH PSI

1 UNEXPOSED 0.016 62,500 44,200 IE EXPOSED 0.016 NO DATA 45,600 2 UNEXPOSED 0.016 62,800 47,000 2E EXPOSED 0.016 62,300 45,600 3 * UNEXPOSED 0.040 28,800 NO DATA 3E* EXPOSED 0.040 28,900 NO DATA 4 UNEXPOSED 0.032 65,300 46,500 4 E EXPOSED 0.032 66,000 47,000

* 52-SO ALUMINUM ALLOY

191

TABLE B6

ELEVATOR SECTION DATA STATION 2 24ST ALCLAD - SLANT RANGE-5330FT AMBIENT TEMP 62° F

V

ELEVATOR SECTION #l "A" TEMP TAPE- NO DATA "B"TEMP TAPE->640°F

TEST SPECIMEN NO

SECTION SIDE

METAL THK IN


YIELD STRENGTH PSI

1 UNEXPOSED 0.016 60,500 49,900

IE EXPOSED 0.018 51,600 32,300

2 UNEXPOSED 0.016 61,300 45,600

2E EXPOSED 0.017 51,700 34,800

3 UNEXPO SED 0.032 66,300 52,300

3E EXPOSED 0.033 65,000 48,000

4 UNEXPOSED 0.032 67,000 55,000

4E EXPOSED 0.032 66,700 50,600

ELEVATOR SECT ION #2 VTEMP TAPE-NO DATA "B"TEMP TAPE- NO DATA

TEST SPECIMEN NO

SECTION SIDE

METAL THK IN


YIELD STRENGTH PSI

1 UNEXPOSED 0.016 62,300 47, 100

1 E EXPOSED 0.016 57,300 41,000

2 UNEXPOSED 0.016 62 , 500 47,700

2E EXPOSED 0.016 59,300 37,700

3 UNEXPOSED 0.032 65,900 51,100

3E EXPOSED 0.032 66,800 48,800

4 UNEXPOSED 0.032 66 , 700 54,200

4E EXPOSED 0.032 65 , 7 00 50,500

192

TABLE B6 (CONTINUED)

ELEVATOR SECTION* 5

"A"TEMP TAPE-468°F ttB"TEMP TAPE-)640°F

TEST SPECIMEN NO

SECTION SIDE

METAL THK IN


YIELD STRENGTH PSI

1 UNEXPOSED 0.016 63,900 46,300 IE EXPOSED 0.016 63,300 50,000 2 UNEXPOSED 0.016 64,600 44,800 2E EXPOSED 0.016 62,500 47,000 3 UNEXPOSED 0.032 68,700 54,900 3E EXPOSED 0.032 68,000 53, 800 4 UNEXPOSED 0.032 70,300 56,600 4E EXPOSED 0.032 68,200 55,300

ELEVATOR SECTION*6

V TEMP TAPE-408°F "ß"TEMP TAPE-NO DATA

TEST SPECIMEN NO

SECTION SIDE

METAL THK IN


YIELD STRENGTH PSI

1 UNEXPOSED 0.016 67,300 41,800 IE EXPOSED 0.016 62,000 48,100 2 UNEXPOSED 0.016 64,300 45,700

2E EXPOSED 0.016 59,900 45,600

3 UNEXPOSED 0.032 64,900 52,300 3E EXPOSED 0.032 64, 000 49,200

4 UNEXPOSED 0.032 60, 800 50, 300

4E EXPOSED 0.032 57, 600 45,700

193

TABLE B.7 RADIANT ENERGY MEASUREMENTS

STATION SLANT RANGE FT.

ENERGY FROM PASSIVE INDICATORS ON FRAMES CALORIES PER CM2

ENERGY VALUES EXTRAPOLATED FROM NML DATA CALORIES PER CM2

1 7150 15 18.3

2 5330 32 30.0

3 3800 >44 43.0

194

BIBLIOGRAPHY

(1) Bureau of Aeronautics Structures Memorandum No. 33 of April 1952, Effects of Airburst Atomic Weapons on Naval Aircraft, Secret Restricted Data.

(2) Naval Air Material Center Report No. ASL NAM DE-260, Part I of April 1952, Instrumentation of Model AD-2 Airplane, BUNO 122363 for Blast Tests. Calibration of the AD-2 Airplane Wing for Flight Load Measurement.

(3) Naval Air Material Center Report No. ASL NAM DE-260, Part II of January 1953, Instrumentation of Model AD Airplanes for Blast Tests; Load Calibration of the Horizontal Stabilizer of the AD-2 Airplane; BUNO 122363. and Load Calibration of the Wing and Horizontal Stabilizer of the XBT2D-1 Airplane, BUNO 09103.

(4) Naval Air Material Center Report No. ASL NAM AD-276 of 17 August 1953, Modernization of/and Use of NAXSTA Radar Tracking Equipment

for Structural Flight Test Programs.

(5) Naval Air Material Center Project No. TED NAM AD-260, Reports to be issued, Instrumentation of Model AD Airplane for Atomic Blast Test.

(6) Naval Air Material Center Report No. ASL NAM DE-260.2, Part I of February 1953, The Positioning of Model AD Drone Aircraft for the Investigation of Critical Atomic V/eapons Effects, Secret Restricted Data.

(7) Armed Forces Special Weapons Project, Operation Tumbler-Snapper, Project 8.3 Report WT-543 of March 1953, Thermal Radiation from a Nuclear Detonation, Secret Restricted Data.

(8) Armed Forces Special Weapons Project Report, Department of the Army Technical Manual TM 23-200, Department of the Navy OPNAV Instruction 003400.1, Department of the Air Force AF0AT 385.2 of October 1952, Capabilities of Atomic Weapons, Secret.

195

(9) Bureau of Aeronautics Report DR-1434, Part I of June 1953 and Part II of November 1953, Reflection of Radiation from an Indefinite Plane.

(10) Bureau of Aeronautics Report DR-1485 of April 1953, Atomic Weapon affects from Standpoint of Air Delivery Vehicles, Secret Restricted Data.

(11

(12

(13

(14

(15

(16

(17

(18

(19

(20

Massachusetts Institute of Technology, Department of Aero- nautical Engineering Report of January 1950, Handbook of Normal Shock Relationships.

Douglas Aircraft Company Report No. ES 6613 Vol. 1 of December 1944, Wing Loads. Model XBT2D-1.

Douglas Aircraft Company Report No. ES 6705 of March 1945, Empennage Loads, Model XBT2D-1.

Bureau of Aeronautics Structures Project Report No. 42 of August 1953, Analytical Determination of the Effect of Aero- dynamic Cooling on the Temperature of a Metal Plate Heated by an Atomic Bomb Thermal Pulse, Confidential Restricted Data.

Douglas Aircraft Co. Report No. ES 17329 of May 1953, Skin Samples Tension Tests, Confidential

Douglas Aircraft Co. Laboratory Report No. ES-3 134 of May 1953, Short Time High Temperature Radiant Head Studies, Secret,

Armed Forces Special Weapons Project, Operation Tumbler Project 1.2 Report WT 512 of February 1953, Air Pressure vs Time, Secret.

Aluminum Co. of America Handbook, ALCOA Aluminum and Its Alloys, pgs. 113-114.

Quartermaster Research and Development Laboratories, Pioneer- ing Research Laboratories Research Service Test Report No. GC-30 of October 1951, Temperature Indicator Papers.

Armed Forces Special Weapons Project, Report No. 903 of November 1953

196

DISTRIBUTION Military Distribution Categories 5-21, 5-30, and 5-60

Copy

2

3

1*- 6

7

8- 10-

9 13

1k- 18

19

20- 22 23 21*

25 26

27

28

29-

31-

30

32

33- 31*

35- 36

37 38

39

i*o

1*1

1*2

U3

1*1*- 1*5

1*6- 1*7

1*8- 50

51

52

53

51*

55- 56

Washington 25, D.C.

57

ARMY ACTIVITIES

Asst. Chief of Staff, G-3, D/A, ATTN: Dep. CofS, G-3 (RR&SW)

Deputy Chief of Staff for Logistics, D/A Washington 25, D.C. ATTN: Director of Research & Development

Chief of Ordnance, D/A, Washington 25, D.C. ATTN: OEDTX-AE

Chief ßignal Officer, D/A, P&O Division, Washington 25, D.C. ATTN: SIGOP

The Surgeon General, D/A, Washington 25, D.C. ATTN: Chief, R&D Division

Chief Chemical Officer, D/A, Washington 25, D.C. The Quartermaster General, CBR, Liaison Officer, Re-

search and Development Div., D/A, Washington 25, D.C. Chief of Engineers, D/A, Washington 25, D.C. ATTN: ENGNB

Chief of Transportation, Military Planning and Intel- ligence Div., Washington 25, D.C.

Chief, Army Field Forces, Ft. Monroe, Va. President, Board #1, OCAFF, Ft. Bragg, N.C. President, Board #2, OCAFF, Ft. Knox, Ky. President, Board #3, OCAFF, Ft. Benning, Ga. President, Board #1*, OCAFF, Ft. Bliss, Tex.

Commanding General, U.S. Army Caribbean, Ft. Amador, C..Z. ATTN: Cml. Off.

Commander-in-Chief, European Command, APO 128, c/o PM, New York, N.T.

Commander-in-Chief, Far Ea3t Command, APO 500, c/o PM,

Commanding General, U.S. Army Europe, APO 1*03, c/o PM, New York, N.Y. ATTN: OPOT Div., Combat Dev. Br.

Commanding General, U.S. Army Pacific, APO 958, c/o PM, San Francisco, Calif. ATTN: Cml. Off.

Commandant, Command and General Staff College, Ft. Leavenworth, Kan. ATTN: ALLLS(AS)

Commandant, The Artillery School, Ft. Sill, Okla.

Secretary, The AA&GM Branch, The Artillery School, Ft. Bliss, Tex. ATTN: Lt. Col. Albert D. Epley, Dept. of Tactics and Combined Arms

Commanding General, Medical Field Service School, Brooke Army Medical Center, Ft. Sam Houston, Tex.

Director, Special Weapons Development Office, Ft. Bliss, Tex. ATTN: Lt. Arthur Jaskierny

Commandant, Army Medical Service Graduate School, Walter Reed Army Medical Center, Washington 25, D.C.

Superintendent, U.S. Military Academy, West Point, N.Y. ATTN: Prof, of Ordnance

Commandant, Chemical Corps School, Chemical Corps Training Command, Ft. McClellan, Ala.

Commanding General, Research and Engineering Command, Army Chemical Center, Md. ATTN: Deputy for RW and Non-Toxic Material

Commanding General, Aberdeen Proving Grounds, Md. (inner envelope) ATTN: RD Control Officer (for Director, Ballistics Research Laboratory)

Commanding General, The Engineer Center, Ft. Belvoir, Va. ATTN: Asst. Commandant, Engineer School

Commanding Officer, Engineer Research and Development Laboratory, Ft. Belvoir, Va. ATTN: Chief, Technical Intelligence Branch

Commanding Officer, Picatinny Arsenal, Dover, N.J. ATTN: 0RDBB-TK

Commanding Officer, Frankford Arsenal, Phila- delphia 37, Pa. ATTN: Mr. C. C. Fawcett

Commanding Officer, Army Medical Research Laboratory, Ft. Knox, Ky.

Commanding Officer, Chemical Corps Chemical and Radio- logical Laboratory, Army Chemical Center, Md. ATTN: Tech. Library

Commanding Officer, Transportation R&D Station, Ft. Eustis, Va.

69- 70

71

72

73

75- 71* 76

77

78

79- 80 81

82

197

Copy

58 Director, Technical Documents Center, Evans Signal Laboratory, Belmar, N.J.

59 Director, Waterways Experiment Station, P0 Box 631, Vicksburg, Miss. ATTN: Library

60 Director, Armed Forces Institute of Pathology, 7th and Independence Avenue, S.W., Washington 25, D.C.

61 Director, Operations Research Office, Johns Hopkins University, 7100 Connecticut Ave., Chevy Chase, Md. ATTN: Library

62- 68 Technical Information Service, Oak Ridge, Term. (Surplus)

NAVY ACTIVITIES

Chief of Naval Operations, D/N, Washington 25, D.C. ATTN: OP-36

Chief of Naval Operations, D/N, Washington 25, D.C. ATTN: OP-37MOEG)

Director of Naval Intelligence, D/N, Washington 25, D.C. ATTN: 0P-922V

Chief, Bureau of Medicine and Surgery, D/N, Washington 25, D.C. ATTN: Special Weapons Defense Div.

Chief, Bureau of Ordnance, D/N, Washington 25, D.C. Chief, Bureau of Ships, D/N, Washington 25, D.C. ATTN:

Code 31*8 Chief, Bureau of Yards and Docks, D/N, Washington 25,

D.C. ATTN: D-1*1*0 Chief, Bureau of Supplies and Accounts, D/N, Washing-

ton 25, D.C. Chief, Bureau of Aeronautics, D/N, Washington 25, D.C. Chief of Naval Research, Department of the Navy Washington 25, D.C. ATTN: LT(jg) F. McKee, USN

Commander-in-Chief, U.S. Pacific Fleet, Fleet Po3t Office, San Francisco, Calif.

83 Commander-in-Chief, U.S. Atlantic Fleet, U.S. Naval Base, Norfolk 11, Va.

81*- 87 Commandant, U.S. Marine Corps, Washington 25, D.C. ATTN: Code A03H

88 President, U.S. Naval War College, Newport, R.I. 89 Superintendent, U.S. Naval Postgraduate School,

Monterey, Calif. 90 Commanding Officer, U.S. Naval Schools Command, U.S.

Naval Station, Treasure Island, San Francisco, Calif.

91 Commanding Officer, U.S. Fleet Training Center, Naval Base, Norfolk 11, Va. ATTN: Special Weapons School

92- 93 Commanding Officer, U.S. Fleet Training Center, Naval Station, San Diego 36, Calif. ATTN: (SPWP School)

91* Commanding Officer, Air Development Squadron 5, VX-5, U.S. Naval Air Station, Moffett Field, Calif.

95 Commanding Officer, U.S. Naval Damage Control Training Center, Naval Base, Philadelphia 12, Pa. ATTN: ABC Defense Course

96 Commanding Officer, U.S. Naval Unit, Chemical Corps School, Army Chemical Training Center, Ft. McClellan, Ala.

97 Joint Landing Force Board, Marine Barracks, Camp Lejeune, N.C.

98 Commander, U.S. Naval Ordnance Laboratory, Silver Spring 19, Md. ATTN: EE

99 Commander, U.S. Naval Ordnance Laboratory, Silver Spring 19, Md. ATTN: EH

100 Commander, U.S. Naval Ordnance Laboratory, Silver Spring 19, Md. ATTN: R

101 Commander, U.S. Naval Ordnance Test Station, Inyokern, China Lake, Calif.

102 Officer-in-Charge, U.S. Naval Civil Engineering Res. and Evaluation Lab., U.S. Naval Construction Bat- talion Center, Port Huenema, Calif. ATTN: Code 753

103 Commanding Officer, U.S. Naval Medical Research Inst.., National Naval Medical Center, Bethesda ll*, Md.

jpnlff]

Copy

lOH Director, C.S. Naval Besearch Laboratory, Washington

25, D.C. ATTN: Code 2029 105 Director, The Material Laboratory, New York Naval Ship-

yard, Brooklyn, N.T. 106 Commanding Officer and Director, U.S. Navy Electronics

Laboratory, San Diego 52, Calif. ATTN: Code lt223 107-110 Commanding Officer, U.S. Naval Radiological Defense

Laboratory, San Francisco 2k, Calif. ATTN: Technical Information Division

111-112 Commanding Officer and Director, David W. Taylor Model Basin, Washington 7, D.C. ATTN: Library

113 Commander, U.S. Naval Air Development Center, Johns-

vllle, Pa. Ilk Director, Office of Naval Besearch Branch Office, 1000

Geary St., San Francisco, Calif. 115 Officer-in-Charge, U.S. Naval Clothing Factory, U.S.

Naval Supply Activities, New York, 3rd Avenue and 29th Street, Brooklyn, N.T. ATTN: BSD Division

116-122 Technical Information Service, Oak Ridge, Tenn.

(Surplus)

Copy

152 Commander, Headquarters, Technical Training Air Force,

Gulfport, Miss. ATTN: TA&D 153-1514 Commandant, Air Force School of Aviation Medicine,

Randolph AFB, Tex. 155-160 Commander, Wright Air Development Center, Wright-

Patterson AFB. Dayton, 0. ATTN: WC0ESP I6I-I62 Commander, Air Force Cambridge Research Center, 230

Albany Street, Cambridge 39. Mass. ATTN: CRQST-2 I63-165 Commander, Air Force Special Weapons Center, Kirtland

AFB, N. Max. ATTN: Library 166 Commandant, USAF Institute of Technology, Wright-

Patterson AFB, Dayton, 0. ATTN: Resident College 167 Commander, Lowry AFB, Denver, Colo. ATTN: Department

of Armament Training 168 Commander, 1009th Special Weapons Squadron, Head-

quarters, USAF, Washington 25, D.C. 169-170 The RAKD Corporation, 1700 Main Street, Santa Monica,

Calif. ATTN: Nuclear Energy Division 171-177 Technical Information Service, Oak Bidge, Tenn.

(Surplus)

AIB FORCE ACTIVITIES

123 Asst. for Atomic Energy, Headquarters, USAF, Washing-

ton 25, D.C. ATTN: DCS/0 12li Director of Operations, Headquarters, USAF, Washington

25, D.C. ATTN: Operations Analysis 125 Director of Plans, Headquarters, USAF, Washington 25,

D.C. ATTN: War Plans Div. 126 Director of Research and Development, Headquarters,

USAF, Washington 25, D.C. ATTN: Combat Components

Div. 127-128 Director of Intelligence, Headquarters, USAF, Washing-

ton 25, D.C. ATTN: AF0IN-1B2 129 The Surgeon General, Headquarters, USAF, Washington 25,

D.C. ATTN: Bio. Def. Br., Pre. Med. Div. 130 Deputy Chief of Staff, Intelligence, Headquarters, U.S.

Air Forces Europe, APO 633, c/o PM, New York, N.Y. ATTN: Directorate of Air Targets

131 Commander, U97th Reconnaissance Technical Squadron (Augmented), APO 633, c/o PM, New York, N.Y.

132 Commander, Far East Air Forces, APO 925, <7° ™> San Francisco, Calif.

133 Commander, Strategic Air Command, Offutt Air Force Base, Omaha, Nebraska. ATTN: Special Weapons Branch, Inspection Div., Inspector General

I3U Commander, Tactical Air Command, Langley AFB, Va. ATTN: Documents Security Branch

135 Commander, Air Defense Command, Ent AFB, Colo. 136-137 Commander, Air Materiel Command, Wright-Patterson AFB,

Dayton, 0. ATTN: MCAIDS 138 Commander, Air Training Command, Scott AFB, Belleville,

111. ATTN: DCS/0 GTP 139 Commander, Air Research and Development Command, PO

Box 1395, Baltimore, Md. ATTN: BDDN \kO Commander, Air Proving Ground Command, Eglin AFB,

Fla. ATTN: AG/THE 1IH-IH2 Commander, Air University, Maxwell AFB, Ala. 1^-3-150 Commander, Flying Training Air Force, Waco, Tex.

ATTN: Director of Observer Training 151 Commander, Crew Training Air Force, Randolph Field,

Tex. ATTN: 2GTS, DCS/0

178

179

180

OTHER DEPARTMENT OF DEFENSE ACTIVITIES

Asst. Secretary of Defense, Research and Development,

D/D, Washington 25, D.C. U.S. National Military Representative, Headquarters,

SHAPE, APO 55, c/o PM, New York, N.Y. ATTN: Col.

J. P. Healy Director, Weapons Systems Evaluation Group, OSD, Bm

2E1006, Pentagon, Washington 25, D.C. Asst. for Civil Defense, OSD, Washington 25, D.C. Armed Services Explosives Safety Board, D/D, Building

T-7, Gravelly Point, Washington 25, D.C. Commandant, Armed Forces Staff College, Norfolk 11,

Va. ATTN: Secretary Commanding General, Field Command, Armed Forces Spe-

cial Weapons Project, PO Box 5100, Albuquerque, 5. Max. Commanding General, Field Command, Armed Forces, Special Weapons Project, PO Box 5100, Albuquerque, N. Mex. ATTN: Technical Training Group

Chief, Armed Forces Special Weapons Project, Washington

25, D.C. Technical Information Service, Oak Hidge, Tenn.

(Surplus)

ATOMIC ENERGY COMMISSION ACTIVITIES

208-210 U.S. Atomic Energy Commission, Classified Technical Library, 1901 Constitution Ave., Washington 25, D-C ATTN: Mrs. J. M. 0'Leary (For DMA)

Los Alamoa Scientific Laboratory, Report Library, PO Box 1663, Los Alamos, N. Mex. ATTN: Helen Redman

Sandia Corporation, Classified Document Division, Sandia Base, Albuquerque, N. Mex. ATTN: Martin

Lucero University of California Radiation Laboratory, PO Box

808, Livermore, Calif. ATTN: Margaret Edlund 222 Weapon Data Section, Technical Information Service,

Oak Bidge, Tenn. 223-260 Technical Information Service, Oak Bidge, Tenn.

(Surplus)

181 182

183

181*-189

190-191

192-200

201-207

211-213

2ll*-219

220-221

198

AEC, Oak Ridge, Tei