PROJECT HAZEL PROPULSION, STRUCTURAL HEATING AND PRESSURIZATION REPORT NO. ZJ-026 OCTOBER 1958

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP89B00709R000400810001-5
• Release Decision: RIFPUB
• Original Classification: S
• Document Page Count: 49
• Document Creation Date: December 27, 2016
• Document Release Date: December 13, 2010
• Sequence Number: 1
• Publication Date: October 1, 1958
• Content Type: REPORT

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P~UJEC
H A Z E L
PROPULSION, STRUCTURAL H,EATING
AND PRESSURIZATION
OCTOBER 1958
A DIVISION OF GENERAL DYNAMICS CORPORATION
SAN WEGO, CALIF,
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C 0 N V
A I R
P I_ 4 zJ-026-----
4 I I1 F SF A',
,. ,n 4
1 C, r58
SAN DIEGO
HAZEL
TITLE 2,.,.
Gtfv~Ff N4Y.
PROPULSION a f'
STRUCTURAL HEATING r,=_ttT Ff '' W ' a E
AU1M' N 1 2
BiEN(Efi
AND DATE, AWS,
PRESSURIZATION STUDI$S
x ITI PE
SUMMARY REPC&O PA 1 E L' ? spa `3 A _ ,
PREPARED BY
R.
Ilk
K. Jo son
Al
GROUP Thermodynamics
R.
K.Livett
~/
(.,v+
REFERENCE
CHECKED BY
W. Broshar
APPROVED BYaa"""
R.
Nau
"
C. E. Chap6an
,,
v ~(
Chief of Thermodyri
~
,
G.
Nicoloff
Es
NO. OF DIAGRAMS
3 F. Brady
Dev pment Pro e t
REVISIONS Engineer
NO.
DATE
BY
CHANGE
PAGE AFFECTED
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ANALYSIS
PREPARED BY
CHECKED BY
REVISED BY
CONVAIR
A DIVISION OF GENERAL DYNAMICS CORPORATION
I SAN DIEGO)
PAGE i
REPORT NO ZJ-026
MODEL HAZEL
DATE 10-31-58
SECRET
This document contains information affecting
the national defense of the United States within the
meaning of the espionage laws, Title 18, U.S.C. Sections
793 and 794. The transmittal or the revelation of its
contents in any manner to an unauthorized person is
prohibited by law.
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ANALYSIS CO N VA I R PAGE ii
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REVISED BY DATE 10-31-58
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FOR=
This report is presented as one of a set describing the
Project "Hazel" study performed by the Convair San Diego Division
of the General Dynamics Corporation. The entire set of reports,
listed below,, represents Convair's fulfillment of the publica-
tions obligation specified in Contract NOas-58-812 (55-100) and
Amendment #1., issued lk August 1958 by the Bureau of Aeronautics.
ZP 252 Summary (Brochure of Charts with Text)
ZP 253 Aircraft Design
ZA 282 Aerodynamics
ZT 026 Propulsion,, Structure Heating,, and Pressurization
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FOAM I6I2411'1
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ANALYSIS
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A DIVISION OF GENERAL DYNAMICS CORPORA tION
I SAN DIEGO)
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TABLE 011 C LNTENT
Security Notice
Foreword
Table of Contents
List of Figures
List of Tables
Introduction
Summary and General Conclusions
Propulsion System
Page
Inlets
Fuels
Engines
Performance
Engine Teat & Facility Requirements
Conclusions and Recommendations
PAGE iii
REPORT NO ZJ-026
MODEL HAZEL
DATE 10-31-58
6
6
8
9
1.1
11
Structural Heating Analysis
13
Introduction
13
Summary
13
Recommendations
13
Discussion of Results
13
Fuel System Heating
13
Structural Heating
15
Structural Pressurization System
16
Introduction
16
Conclusions
16
System Requirements
16
Systems Considered
16
Proposed System
18
44
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ANALYSIS
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A DIVISION OF GENERAL DYNAMICS COPPORATION
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LIST OF FIGURES
PAGE 1
REPORT NO ZJ-026
MODEL HAZEL
DATE 10-31-58
Figure
PaRe
1
Combustor Inlet Total Pressures and Total Temperatures
23
2
Pratt & Whitney Engine Configuration
24
3
Marquardt Engine Configuration
25
4
Marquardt Engine Net Thrust Coefficient vs. Specific
Fuel Consumption - Pentaborane Fuel
5
Marquardt Engine Not Thrust Coefficient vs. Specific
Fuel Consumption - SF-1 Fuel
6
Assumed Combustion Efficinecy vs. Altitude - Marquardt
Optimization Study
28
7
Marquardt Engine Weight va. D3
29
8
Marquardt Off Design Performance - Pentaborane Fuel -
120,000 ft.
30
9
Marquardt Off Design Performance - Pentaborane Fuel -
135,000 ft.
31
10
Pratt & Whitney - % Diameter, Length and Weight vs. % Thrust
32
11
Validation of Design Point Data - Pentaborane - (A) =
0.015
33
12
Validation of Design Point Data - Pentaborane - (F) =
0.040
34
13
Validation of Design Point Data - Pentaborane - (F) =
0.019
35
14
P & W Testing Limits - Wilgoos Lab
36
15
Marquardt Engine Facility Capability
37
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ANALYSIS
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A DIVISION OF GENERAL DYNAMICS CORPORATION
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PAGE 2
REPORT NO ZJ-026
MODEL HAZEL
DATE 10-31-58
Figure
FaRe
16
Vapor Feed Fuel System Schematic
38
17
Proposed Pressurization System Schematic
39
18
Alternate Pressurization System Schematic
39
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ANALYSIS
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A DIVISION OF GENERAL D~NAMIL3 COPYORArION
I SAN DIEGO,
PAGE
REPORT NO
MODEL
DATE
3
ZJ-026
HAZEL
10-31-58
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LIST OF TABLES
Table
Page
1
Pentaborane vs. SF-1 Fuels - Ground Handling & Logistics
6
2
Pratt & Whitney Engine Data - SRJ-43D & SRJ-43E - SF-1
Fuel - Mach 3.0
3
Pratt & Whitney Engine Data - SRJ-43D & SRJ-43E - SF-l
Fuel - Mach 2.5
Pratt & Whitney Engine Data - SRJ-43D - Pentaborane Fuel
?1
Mach 2.5 and 3.0
42
5
Pratt & Whitney Engine Geometry
43
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ANALYSIS C O N VA I R PAGE 4
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INTRODUCTION
This report presents the results of studies of engine and
inlet performance, structural heating problems, and structural
pressurization systems, carried out by the Thermodynamics Group.
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ANALYSIS
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A DIVISION OF GENERAL DYNAMICS CORPORATION
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REPORT NO ZJ-026
MODEL HAZEL
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SUMMARY AND GENERAL CONCLUSION
Several inlets were examined by Convair, Pratt & Whitney and
Marquardt for this application. It was concluded that the fixed
isentropio spike diffuser with slight internal contraction if necessary
would be best,,
Engine performance as presented by Pratt & Whitney, and Marquardt
is exhibited in the report. This performance was checked by Convair and
found correct with the reservation that the combustion efficiencies
assumed will have to be verified by testing.
Hydrogen appears to be better than pentaborane from a propulsion
and handling standpoint. Both fuels are adequate for the altitude of
this mission.
Both engine companies have facilities that will be available by
1960 that pan handle the engines they propose. Government facilities
are also available at NACA and A.R.D.C.
Structural temperatures were found to be within operating limits
of the materials proposed, with the exception of some sections of the
engine, where additional materials study is indicated. Fuel heating
will not be a major problem for the fuels proposed. Wing surface temp-
eratures will vary from 6300 F at the leading edge to 400 and 3000 F one
and ten feet, respectively, from the leading edge.
Minimum structural pressurization system weight is obtained by
utilizing helium, stored in the liquid state and heated after evaporation
by mixing with hydrazine exhaust products from the auxiliary power unit.
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ANALYSIS
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A DIVISION OF GENERAL DYNAMICS CORPORATION
(SAN DIEGO)
PAGE 6
REPORT NO ZJ-026
MODEL HAZE,
DATE 10/31/58
SECRET
PROPULSION SYSTEM
Early proposals by Pratt and Whitney and Marquardt were somewhat conserva-
tive on pressure recovery, both using values of about .70. Boost and ranee
considerations indicated that the best system would most probably dictate ram
jet take over at or near the design Mach number. This allowed better diffuser
design point selection. Current peak pressure recoveries used were about .77 -
?79 at Mach 3.0. Under these conditions the best type of dif:'user appeared to
be the fixed isentropic spike. The nearest contender was the Internal Compres-
sion Inlet which may well have been selected on a total thrust minus drag basis
but was not because of higher weight. This resulted from its longer design and
moveable spike. The fixed isentropic spike inlets selected gave a total external
drag coefficient of .11 based on engine area. Of this, .06 was wave drag and
.041 was skin friction of the engine external surface.
The inlets had to be placed with respect to the wing in a way that satis-
fied radar visibility restrictions. In over wing location resulted and two
arrangements were found as satisfactory compromises. Two engines located out-
board about mid half span can be situated over the drooped leading edge so that
the upper wing surface with a minimum of flattening can give zero angle of attack
with respect to the inlet. One engine centrally located can be placed behind the
apex of the delta planform with a portion of the surface made plane at zero angle
of attack to the inlet.
It was found that the recovery penalty suffered from expansion over the
resulting flat surface when the vehicle was operated at higher than design and
of attack was leas than that suffered from the inlet in free stream at the same
off design angle. This is because the flow expands over the flat surface paral-
lel to the axis of the inlet. For a 2' positive angle of attack the loss in
pressure recovery is about 4% behind the flat surface and 6.6% in free stream.
Both Pratt and Whitney and Marquardt claim to have adequately tested the
inlets selected at the required design Mach number. Neither has matched the
Reynolds number of the flight condition, however.
A table of physical and handling characteristics is given below
TABLE i
P ABORANE VS SF-1 FUELS
GROUND HANDLING & LOGISTICS
Property Pentaborane SP-1
Price $/lb. lbday 20 1-10
Large Scale Production 3 .2
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ANALYSIS
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A DIVISION OF GENERAL DYNAMICS CORPORATION
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Prt Fentaborane
Explosive Hazard Pyrophoric
Toxicity
Storage
Hazards
Extremely toxic either by
inhalation or contact.
OK with anodized alum., cop-
per, steel. NACA RM E54 E12
has data on materials
Inert atmosphere. No leaks can
be tolerated
Toxicity and prophoric proper-
ties require inert atmosphere
transfer system and protective
clothing with respiratory
protection.
AIRPLANE PERFORMANCE CONSIDERATIONS
PAGE 7
REPORT NO. ZJ-026
MODEL HAZEL
DATE 10/31/58
Air mixtures can be igni-
ted by a spark
None. Can suffocate if it
displaces all oxygen. Cold
"burns" because of low
temperature.
Non corrosive. Can cause
low tem!)erature embrittle-
ment. it-8 steels; law C,
high He steel; Monels are
OK. Plastics will have to
be checked out.
Dewar tanks. Boil off
must be ventilated.
Protective clothing to
protect against cold
"burns
t
Pentaborane
v
preer
Heating value BTU/ib
Approx. 29,300 (sower)
51,500 (lower)
30,300 (higher)
Boiling Point OR
Approx. 600
37
Density at B.P. lb/ft3
Approx. 37
4.4
Use as Cooling Fluid
Decoatposes at temp. 260' F
Excellent coolant
Tanks and Lines
Inert transfer system
insulation needed
Insulation racy be needed on
tanks to prevent thermal ds-
composition.
It can be seen that hydrogen appears to have w%vantages in the handling area
since it is neither toxic nor pyrophoric. Also considerable experience with hand-
ling and pumping cold liquids has been gained with liquid rockets in recent years.
on performance it has higher heat content and clean exhaust. It does not have
a solid-liquid vapor phase as does pentaborane.
1antaborane has advantages in its ability to produce a strong stable flame
at high altitudes. It has better volume characteristics for tankage.
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Both fuels appear. adequate for the combueb.on conditions anticipated for the
Hazel vehicle. Hydrogen would seem to be somewhat marginal at altitudes above
150,000 feet at the chosen design Mach number. This is based on preliminary
results from Marquardt and depends partially on combustion efficiency assumptions.
This does not seem to be a problem, however, as the present mission does not
attain this altitude. Approximate combustion pressures and inlet temperatures
are shown on Figure 1, together with Marquardt and Pratt and Whitney test data
available on the two fuels. The test data is at or near Mach 2.0 but combustion
conditions may be expected to improve at the same pressures and higher tempera-
tures encountered at Mach 3.0. No combustion efficiency data was derived from
these tests, a fact th4t has led to a marked difference in design of combustion
chamber lengths as will be brought out later.
As was requested'by the Navy, Marquardt and Pratt and Whitney were the
only engine companies approached for performance and design data. The results
received from them are presented at the design points chosen'for each engine.
These are also substantiated by calculations made by Convair in the region of
the selected design points.
The engine size range was established by an interchange of estimated L/D'a,
gross weights and flight conditions between Convair and the engine companies.
Latitude on either side of the estimated design sizes was given to allow for
changes produced by more detailed calculations. Contacts with both companies
were made regularly by visit and mail to resolve design problems and interchange
data.
Somewhat optimistic engine performance and weight data was given earlier by
Marquardt, while the reverse was essentially true of Pratt and Whitney. Sub-.
sequent results received are in much better agreement between the two companies.
The mission performed starts at 125,000 feet and ends at approximately
140,000 feet. It is assumed that. the vehicle will be boosted to the design Mach
number of 3.0 and follow a Breguet range path at constant Mach number. Mach 3.0
was necessary to keep within the'structural limits of the Marquardt engine, as
well as the plastic airframe.
The Pratt and Whitney engine is shown on Figure 2. It is constructed of
high temperature steels throughout. The fuel system is designed to vaporize the
fuel within the cowl surfaces and center body. This general fuel system approach
is proposed for both hydrogen and pentaborane. The same basic geometry was held
using pentaborane as SF-1 except that the exit nozzle throat diameter was ad-
justed. This was done to match the higher combustion temperature considered op-
tinaun by Pratt and Whitney for pentaborane. The inlet shown is not the final
Pratt and Whitney design. In place of the two step cone, an isentropic spike
was used and the cowl lip geometry altered to match.
The use of the engine was limited earlier to 54.8" exit diameter by Pratt
and Whitney facility capability. This was relaxed to 104" diameter as later
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ANALYSIS C O N VA I R PAGE 9
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facility availability data revealed possible. A graph showing Pratt and Whitney
facility capability by 1959 and 1960 is ehawia in Figure 14. The subject of
facilities is discussed later.
The Marquardt engine is shown on Figure 3. The proposed construction is
plastic honneyccl>sb except for the fuel system., flaw holder, cooling shroud, and
engine mounts. The plastic Is a ceramic fiber impregnated with a high tempera-
ture phenolic. Plastics of this type are marketed under the trade name of
"Refraail". The maxim= skin temperature allowable is 800 - 904' . Very little
data is available for strength at these temperatures for time periods typical of
the Hazel mission. The critical point is at the exhaust nozzle throat where the
double skin area surrounding the exit nozzle has to be perforated to allow cooling
by radiation leakage.
Marquardt is facility limited to 8' diameter as shown on Figure 15. They
do not look at the scaling problem for this application as being a great risk,
however. This is backed with considerable experience in the ram jet field.
The Marquardt engine, as was Pratt and Whitney's, is designed for vaporized
fuel. In this case, too, the nose cone and recirculation zone walls are utilized
but additional heat exchanger surface supplied by Convair is required. This is
described in more detail under final tankage study results elsewhere in the report.
PF. 4RXANC H
Mar~wardt
Data presented by Marquardt for design point selection is shown on Figure 4
for the pentaborane fuel and. Figure 5 for the SF-i fuel. The "net jet" thrust
coefficients are based on "A3" as sham an the inset aketch Figure 5. These data
are based on the combustion efficiency variation assumed for a 16' combustion
chamber length and shown on Figure 6. Shorter lengths were exodned but there was
no real requirement. Uie weight of the additional length of combustion chamber
was negligible compared to the lose in range caused by a reduction in length.
The engine selected has a geometry peculiar to the design points shown on
Figures 4 and 5. The exit nozzle throat and exit areas "A5" and "A6", are given
as ratios to a reference area A A. These ratios are held for the entire graph while
the inlet area "Ac" is allowed ~o vary to place the diffuser always at design
pressure ratio. '.Thus, each point on the graphs represent a single engine geocetry.
The basis for choice of the particular engine in each case was a compromise between
engine size and beet specific fuel consumption. The choice of the particular. set
of 55/A3 and A6/A3 ratios resulted from the exchange of vehicle LID, and gross
weight data with Marquardt, which led to a narrower field of engine geometries
giving best range of the totalvehicle. The curves supplied by Marquardt and used
for engine weights are shown on Figure 7. lire-diameter and combustion chamber
length is given along with the effect of altitude on the design weight of the
engine at a combustion chamber lengtft of 16 feet. Off design per-
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ANALYSIS CO N VA I R PAGE 10
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S E C with ntaborane fuel
formance variation with Mach number is ven or two altitudes pe
on Figures 8 and 9 . Effect of angle of attack is also shown on these figures.
Pratt and Whitney
Data presented by Pratt and Whitney for design point selection Is presented
in Tables 2 through 5 . The size of this engine was fixed at what Pratt and Whitney
considered reasonable for full scale testing. Given the same general input on
mission requirements, a, basic engine geometry was established by Pratt and Whitney.
A passible alternate was provided for the SF-1 engine only. These are designated
SRJ-b3D for both the SF-1 and pentaborane engines, while the SF-1 alternate is the
SRJ-43E. The engines were scaled down in size where necessary but not up, as this
would emceed facility Halite. The scaling curves provl d by Pratt and Whitney
are given in Figure 10.
Performance and basic physical data of the engine is given at altitudes from 1
80,000 feet to 150,000 feet an a standard day at the design Mach number of 3.0 and
for 80,000 to 135s000 feet at the off design condition of Mach 2.5. The off design
data was rusted of Pratt and Whitney for turns and/or climb. Angle of attack
effect on performance was also provided by Pratt and Whitney to determine the effect!
of a 20 trim error on angle of attack. This is also shown on the tables.
Coarative Data
The engines selected are arranged in tabular form below with pertinent physi-
cal dimensions and performance. All engines are for a design point of Mach, 3.0 on
a standard day. The range for all cases is 3240 nautical miles.
Dimension In.
Total
Engine
Total
Fuel-
Specific
Diameter
Weight
Thrust
air
FM1
B.
Fuel
Fi
I.iet dt
h
Pounds
Pounds
Ratio
Consumption
(Vehicle: PC 22; Altitude: 135,000 ft. at start of cruise)
2
SF'-1
73-
"-.3
226
17100
2382
.01)48
.880
PBoW
(Vehicle: PC 20; Altitude: 125,000 ft. at start of cruise)
2
P9
68.3
82.7
221
1590
3520
[_??66
t 1.950
(Vehicle:.MC2?; Altitude: 139,000 ft. at start of cruise)
2
SF-1
84.7
lk.7
808
920
2260 .0150
.970 --~
I~AR-
QUARDT
(Vehicle : ).I 10; Altitude : 125,000 ft. at start of cruise)
1
125.2
153.5
786
1460
3640
.0175
1.460
(Vehicle : MC 20; Altitude : 125s000 ft. at start of cruise)
2
PB
[ 85.7
105.3
808
1340
1695
.0175
1.460
__
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ANALYSIS C O N VA I R PAGE 11
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As can be seen, Marquardt presented pentaborane engine data at lower
fuel - ratios favoring the plastic construction. The lower fuel - air ratios
dictated a larger engine. Marquardt engine weights are lover despite the
larger size by a considerable margin. This is due to two factors; (a) the
plastic construction with a more liberal use of honeycomb structure and (b) the
size restriction by Pratt and Whitney which required that two Miller engines
be used instead of one larger engine, rvit'? )lit-1,111T ..~e i r_L- ,~ ? S_1~a~Y. in Figure 10.
;4n ~1 vIous difference in the two engines is in the over-all len.-ths.
This is due to the difference in combustion chamber len;;ths. Marquardt used
16 feet while Pratt and Whitney used 4 feet. Marquardt may be quite conserva-
tive but the questi?n of which is correct can only be resolved throughi adequate
testing. As yet, neither company has measured combustion efficiency accurately
enough. This may not be an important issue, however, as very little engine
weight is involved and the space requirements of most vehicles studied will
permit both engine lengths.
A check was made c-y Convair on the performance Wtimates of the two
companies. This was clone at comparative performance points using Convair
methods and without knowledge of the complete cycle assumptions made by either
engine company. The results are shown in Figures 11 through 13. Theagreement
was very good in all cases on both engine geometry and performance, and is con-
sidered adequate to substantiate both companies estimates. The differences
that do exist may be caused by slightly differing diffuser efficiencies, assump-
tions of combustion total head loss and degree of dissociation and recombination
in the exhaust nozzle.
TESTING AND FACILI:'t REQUI 1P. 4T8
Both Marquardt and Pratt and Whitney have adequate home facilities for
rain jet testing. Pratt and Whitney will have capacity by 1960 for testing its
86 inch engine as shown on Figure 14. At Mach 3.0 there appears to be aauffi-
clent margin to operate with the exit nozzle throat sonic. Marquardt facility
capacity is shown on Figure 15. The 8 foot diameter engine can be operated
with the exit nozzle throat conic simulacing the Mach 3.0 case at 125,000 feet.
Certainly, the two engine versions of both Pratt and Whitney and Marquardt' s
engines can be tested with the facilities available in. the time period. It is
also very p'obable that Marquardt 's single engine is within scaling distance of
the 8 foot diameter engine which would be simulated with sonic exit at 125,000
feet altitude and with subsonic exit and combustor Mach number matching at
1110,000 feet.
Capacity is also available at the present time at RAGA to handle a 10.5
foot engine to 117,000 at Mach 3.0 with sonic exit. A.E.D.C. plans within one
year to handle an 8 - 9 foot diameter engine at 135,000 feet under the same
conditions.
CONCLUSIONS AND IMCOMMIDATIONS
The estimates of engine performance appear to be correct depending upon
the validity of the combustion efficiencies assumed. It is recommended that
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ANALYSIS
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A DIVISION OF GFNFRAL DY NAMI'S CORPORATIUN
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PAGE 12
REPORT NO ZJ-026
MODEL HAZEL
DATE 10/31/58
SECRET
testing be directed toward substantiating the values assumed as this affects
range directly.
There appears little question that Pratt and Whitney can build the engine
presented using the materials selected. Similarly, Marquardt could build their
plastic design but it is obvious that more development wort, would be required
to obtain the advantage that plastic offers in weight saving. It is also not
entirely clear that by use of honeycomb structure the metal engine could not
have been made lighter.
Both fuels.: have undesirable logistics characteristics but are considered
essential to do this high altitude mission. Hydrogen appears to present the
least over-all problems-frcm the propulsion standpoint. Its volume characteris-
tics appear to require a two engine vehicle.
Facilities apparently can be made available in the required time period
that will satisfy the basic engine needs either at home facilities pr Government
test laboratories.
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ANALYSIS CO N VA I R PAGE 13
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REVISED BY DATE 10-31-58
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STRUCTURAL HEATING AMIS
INTRODUCTION
The basic problems of fuel and structural heating have been evaluated.
Aerodynamic heating and heating effects from the engine, contribute to increase
the temperature of the basic structures. Both factors are considered in the
analysis.
SUI~II4AR_,,, x
Structural temperatures will. not exceed reasonable operating limits for the
materials proposed, with the exception of some sections of the engine case. Some
areas of the engine case may require additional materials study for an optimum
design.
Fuel heating will not be a major problem if a liquid fuel system is selected.
No insulation will be required for a liquid pentaborane system. A hydrogen fuel
system would only require insulation to avoid icing conditions.
The wing surface temperature will vary from 6300 F at the leading edge to
400 and 300? 7 at one-foot and ten-feet from the leading edge respectively.
YEC ATIONS
During early stages of development, run heat flow test across simulated engine
walls to ascertain thermal transmission. Radiation properties are of prime
importance.
Determine rates of decomposition and deposits within fuel controls and the
heat exchanger if vaporized pentaborane is used as fuel.
DISCUSSION OF RESULTS
Fuel System Heating
Two types of fuel systems were investigated, gaseous and liquid injection.
Pantaborane and SF-1 were considered for both systems. With the liquid injection
systems, the possibility of fuel losses by evaporation, and malfunction of the
fuel system due to vapor entrainment, are two major problems. The major problems
in a gaseous distribution system are the correct sizing of generators, or heat
exchangers, to vaporize the liquid, and considerably larger flow controls than
normally needed for liquid systems.
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ANALYSIS
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PAGE 14
REPORT NO ZJ-026
MODEL HAZEL
DATE 10-31-58
Pentaborane fuel can be used as a liquid in the proposed tank arrangements
without requiring insulation to avoid overheat, or boiling, at a 15 psi fuel
system pressure. This requires a fuel temperature at take-off below 60? F, which
is not considered restrictive. This system will result in lesser maintenance
problems due to the absence of deposits of decomposed fuel elements. The analysis
was based upon cylindrical fuel cells. If the internal wing volume were used to
store fuel in bulk, a small amount of insulation may be required on the lower
surface to maintain fuel temperatures below boiling. A schematic diagram of a
liquid system is shown on Figurel6. Due to a possible fire hazard, the fuel tank
pressure relief line must be vented downstream of the vehicle. Fuel decomposition
is negligible.but the system should be flushed after each flight. Deposits in the
fuel system that occur, due to temperature, are absorbed by fuel at temperatures
below 100? F and sea level pressure. Therefore, fuel may be used to flush the
system after each flight.
Figure 16 describes a vapor feed system for pentaborane. With a vapor feed
system a large amount of energy is absorbed by the fuel during vaporization. It
is evident from the analysis that a minimum of 1500 aq.ft. of external surface
would be required to evaporate the fuel, at the required rate, by aerodynamic
heating. A heat exchanger may be made as an integral part of the engine wall
using only 75 sq.ft. of surface. Any deposits within the heat exchanger can be
removed by flushing after each flight. The fuel flow diagram shows the gaseous
fuel bubbling through the liquid fuel. This will minimize the deposits within
the flow controls and spray nozzles.
Sufficient vapor for starting must be stored within the tanks. to minimize
the storage volume the pressure at light-off must be at a system maximum, and the
temperature must be at the boiling point. This will allow vapor generation by
lowering the tank pressure during the time the heat exchanger is becoming
operative. The required vapor boil-off rate is maintained by controlling the
pumping rate through the heat exchanger.
SF-1 fuel has the inherent problem of boil-off at very low temperature. While
this is helpful in flight, in reducing heat exchanger size, it creates high fuel
losses and icing problems during and previous to launch.
Approximately two-inches of insulation will be required to avoid icing. A
weight. saving of the vehicle may result by developing rapid fuel handling techniques
and accepting the icing penalties encountered during last minute ground check out
of launch. A minimum weight exchanger would probably be one that is an intergral
part of the engine wall. This arrangement would require a heat exchanger of
approximately 50 sq.ft., based on a fuel consumption of 1+000 #/hr.
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ANALYSIS
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A DIVISION OF GENERAL DYNAMICS CORPORATION
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PAGE 15
REPORT NO. ZJ-026
MODEL HAZEL
DATE 10-31-58
SECRET
STRUCTURAL HEATING
The basic airframe heating is caused by the usual aerodynamic and solar heat-
ing. Some areas are also effected by radiation from the engine surfaces.
The selection of materials is a major factor in the thermal analysis of the
engine case. Due to the high temperature of the combustion gases, 3100? R, the
inner surface of the engine absorbs large quantities of heat, both by radiation
and convection. The amount of structural cooling done by inner passage air flow,
or external flow, is limited by the high energy level of the ambient air stream
(approximately a 1250? R boundary layer and a 1350? R stagnation temperature) and
the low convective heat transfer coefficients. This means that engine structures
must rely on thermal radiation to the atmosphere for cooling. Preliminary
investigation shows that the materials selected by the engine manufacturer can be
surfaced to control thermal emissivity and, therefore, the temperatures can be
maintained within the limits to which the materials can perform.
Limited information is available on the deposits of combustion products on the
engine walls. Additional data is also required on the gaseous radiation to the
engine walls. Both of these areas will have to be investigated for an optimum
design of the engine structure.
I Placing the heat exchanger, for vaporizing the fuel, on the inside engine
wall will result in lower structural temperatures in a local area. Some advantage
may be gained by this in the detailed design.
The maximum heating during the cruise portion of the trajectory of the pro-
posed Hazel vehicle occurs at its beginning (M.= 3 @ 125,000 ft.)
The temperatures of the wing were determined from steady state equilibrium
heat balances by equating the engine, solar, aerodynamic, and terrestrial heating,
to the radiation to space. The flow field at this condition would be laminar.
The aerodynamic heating for the flat portion of the wing was evaluated by the
Reference Temperature. Method (2) and the predicted temperatures are 400 and 300? F
at one-foot and ten-feet respectively, from the leading edge.
The nose stagnation temperature of the vehicle was determined by the method
of Sibulkin (1). The temperature determined by this method was 725? F, considering
two inch radius.
The temperature of the stagnation line of the leading edge was also determined
by the method of Sibulkin, but modified by the cosine of the effective sweep angle
in order to account for the sweep of the leading edge. The temperature determined
by this method was 630? F, considering a two inch radius.
None of the above predicted temperatures have been found to be prohibitive for
the proposed structure.
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ANALYSIS
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I SAN DIEGO)
PAGE 16
REPORT NO. ZJ-02b
MODEL HAZEL
DATE 10/31/58
SECRET
STRUCTURAL PRESSURIZATION SYSM
INTRODUCTION
The inflatable configuration of the vehicle consists of a rigid pilot's
capsule and engine structure supported by a pressurized airfoil. This report
evaluates various systems for supplying this pressurization and outlines those
found most promising. From these, selection is made on the basis of minimum
weight and operational suitability.
CONCLUSIONS
1. Minimum system weight is afforded by a system utilizing helium
stored in the liquid state and heated by direct mixing with hot gas from the
monopropellant APO' hot gas generator. Total weight of the proposed system is
142 pounds.
2. Pure helium free of the hydrazine decompositon products can be sup-
plied to the structure by an alternate system for a 19 pound weight penalty.
Alternate system weight is 161 pounds.
3. Tests should be conducted to determine compatability with structure
and explosive hazard of gas mixture containing hydrazine decomposition products
at operating conditions.
4. Data on leakage rates for materials and construction employed should
be obtained and all possible steps taken to reduce these quantities.
SYSTEM RE
Initial pressurization is supplied on the ground prior to take-off. Means
must be provided for the controlled escape of a portion of this gas with (i)
decreasing ambient pressure as the vehicle is lifted to 45,000 feet and boosted
to 125,000 feet and (2) increased internal temperature due to aerodynamic heat-
ing during cruise. Following this loss and stabilization at cruise conditions,
gas must be added to offset leakage and maintain the given 15 psig pressure
differentail as increasing ambient pressures are encountered during let-down
from altitude. The inlet gas must be injected at such temperatures as to pre-
clude thermal damage to the structure and to minimize total system weight. The
pressurizing medium chosen must remain a gas over the temperature and pressure
range encountered within the structure.
SYSTE(B CONSIDERED
The requirements on the pressurizing medium of (1) remaining a gas over the
operating temperature range of the structure, and (2) having low weight, suggest
the use of the low molecular weight gaseous elements. Table I below lists some
properties of these gases.
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ANALYSIS
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C O N VA I R PAGE 17
A DIVISION OF GENERAL DYNAMICS CORPORATION REPORT NO. ZJ-026
(SAN DIEGO) MODEL HAZEL
DATE 10/31/58
SECRET
TARLZ Z
Properties of Pressurizing Gases
Critical
Density Critical Pressure
Gas Referenced to H2 Temperature, ?F atm Remarks
Hydrogen (H2) 1 -400 12.2 Highly Inflamable
Helium (He) 2 -450 2.3 Inert
Nitrogen (N2) 14 -233 33.5 Inert
Oxygen (02) 16 -182 49.7 Reactive with
Structural Material
Neon 10 -380 26.9 Inert
Argon 20 -88 48.0 Inert
Hydrogen has the lowest density and thus affords the minimum weight penalty for
the pressurizing gas itself but presents an explosive hazard. Oxygen is heavy
and could react chemically with structural members at elevated temperatures.
Neon offers no advantages over Helium, is heavier and less readily available.
Similarly, Argon affords no advantage over Nitrogen. Thus, the choice from this
group for the pressurizing gas is between Nitrogen and Helium. Table II below
lists the advantages and disadvantages of these two gases as the pressurizing
medium.
TABLE II
Comparison of Helium and Nitrogen as Pressurizing Medium
Gas Advantages Disadvanta;;es
Helium Density 1/7 as great Must be transported to point of use;
higher leakage rate; liquifies only
at extremely low temperature.
Nitrogen Can be produced at site of use For equal volume leakage, 7 times
in either liquid or gaseous weight of He required.
form; liquifies at higher
temperature.
Also to be considered are the low molecular weight compounds existing as
gases at the temperatures and pressures considered. Those include such compounds
as ammonia (NH , M * 17), methane (CHt,, M o 16), and others. Some of these, such
as ammonia, ofrer the advantage of remaining a.liquid at ambient temperatures and
only moderate pressures an4 would thus require a simpler and lighter container
system. However, almost all of these compounds are toxic and/or inflammable and
afford no over-all weight advantage as seen below.
Based on the vehicle requirements as outlined in the following section of
this report, the required weights of various gases for the let-down re-pressuri-
zation with zero leakage were calculated and listed in Table III below. This
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ANALYSIS
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PAGE 18
REPORT NO. ZJ-026
MODEL R AZM
DATE 10/31/58
weight is for gas alone and includes no allowance for container weight.
TABLE III
Required Weights of Gases for Let-Down Re-Pressurization
Helium 34 lbs
Nitrogen 238
Ammonia 144
Methane 136 "
extreme corrosiveness while hydrogen and neon were discussed and rejected previously
PROPOSED SYSM
The following calculation demonstrates the necessity of using a low molecular
weight gas for the pressurization of vehicles of this size. Assuming the required
weight of helium to be 34 lbs., data from Reference 3 gives the weight of a
suitable storage container for the gas in liquid form as
container weight 16 + 1.53 x (weight of He)
container weight = 68 pounds
or a combined gas and container weight of 102 pounds. Considering the gas weight
alone, the density of a second gas must be less than = 3 times greater than -144 34
that of helium to show a weight saving. This second gas must t3ierefore have a
molecular weight less than 12. This condition is met by only three substances
besides helium which are gases under the operating conditions; hydrogen fluoride,
hydrogen, and neon. Hydrogen fluoride is dropped from consideration due to its
The configuration and conditions assumed are listed below in Table IV.
TABLE IV
Configuration and Assumed Conditions
Configuration
M-10
Cruise duration
98-minutes
Descent duration
10-minutes
Assumed gas temperature at cruise
3)0? F
Assumed gas temperature at landing
-50? F
Assumed gas temperature at take-off
600 F
Wing area
1985 ft2
Total area of pressurized sections
3967 ft2
Inflatable volume
1720 ft3
Structural pressure differential
15 psig
Launch altitude (boost)
45,000 feet
Cruise altitude
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125,000 - 137,800 feet
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The proposed system is shown schematically in Figure 1'T and an alternative
system in Figure 18. Both systems utilize initial pressurization before take-off
with helium gas, with make-up gas for leakage and let-down re-pressurization
supplied from a storage bottle of liquid helium. Both systcrns utilize hot gas
from the monorpopellant (hydrazine) auxi,Ltary power supply hot gas generator
for beating the very cold helium prior to its use in the structure. They differ
only in the method by which this heating is accomplished. The former utilizes
direct mixing of the hot and cold gaacU while the latter passes the gases through
a heat exchanger allowing only pure helium to enter the structure. The proposed
direct mixing system requires less total gas and does not require the added
weight of the heat exchanger. However, the compatibility with structural materj-
als and the safety of the resulting gas mixture containing the hydrazine decom-
position products must be proven by tests. As shown in Table V, the maximum
concentration of hydrogen within the structure is 5% by volume which is within
the explosive limits of hydrogen in air (4.1 to 47.2% by volume). Oxygen within
the inflated volume will however, be limited, due to the positive pressure
differential above the ambient, to leakage from the pilot's capsule. In addi-
tion, it should be noted that due to adding the helium gas cold for leakage
make-up as noted below, the structure contains only pure helium during all phases
up to and including cruise with hydrazine gas added only during the let-down
phase.
Both systems make use of electrical heaters within the liquid helium
storage tank for maintaining internal pressure as gas is withdrawn. Operation
of both is based on the assumption that make-up for leakage during cruise would
require very low flow and could be made with unheated gas direct from the liquid
tank with heating supplied from the hot structure.
Table V shows the amount and composition of gas present in the structure
at various phases of the flight for both systems.
TABLE V
Structural Gas Content and Composition
Phwe Flitf
Item Di oct Mixing
Pore Hp in 3toct.
Take-off from
Gas in Structure (He) 37 lb.
37 lb.
Sea Level
Air Displaced 132 lb.
132 lb.
Net Lift 95 lb.
95 lb.
Start of Boost
Gas in Structure (He) 28 lb.
28 lb.
Stabilized cruise
Gas in Structure (He) 13 lb.
13 lb.
Landing
Gas in Structure 58 lb.
47 lb.
He 42 lb.
47 lb.
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ANALYSIS
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r
A DIVISION OF GENERAL DYNAMICS CORPORATION REPORT NO ZJ-026
(SAN DIEGO) MODEL HAM
DATE 10/31/58
Phase of
Flight Item
Landing H2
(continued) N2
NH3
SECRET
TABLE V
(continued)
Prop aed System Alternate System
Direct Mixing Pure He in Struct.
1 1b. 0 1b.
10 lb. 0 lb.
5 1b. O lb.
Air Displaced 132 lb. 132 lb.
Net Lift 74 lb. 85 lb.
Gas Composition % by Volume
He 72 % 100 %
192 N2 17 9
N3 9 %
A weight breakdown of the two systems is given in Table VI. This weight
includes a 20% safety factor on required amounts of gas.
System Weight Breakdown
Proposed Alternate System
System Direct Mixing Pure He in Structures
Liquid Selium 35 1'2 lbs. 40 1'2 lbs.
Helium Dewar 70 182 lbs. 77 1'2 lbs.
Mixing Chamber 2 lbs.
Heat Exchanger 5 lbs.
Valves 7 lbs. 7 lbs.
Ducting & Miscellaneous 5 lbs. 5 lbs.
Sub Total 119 lbs. 134 lbs.
Hydrazine (Hot Gas) 19 lbs. 22 lbs.
Sub Total 138 lbs. 156 lbs.
Batteries for Electric Heater 4 lbs. 5 lbs.
142 lbs. 161 lbs.
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ANALYSIS
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PAGE 21
REPORT NO. ZJ-026
MODEL AA +'L
DATE 10/31/58
1 - Weight includes a 20 percent safety factor on required gas weight.
2 - These weights will be increased due to leakage as outlined in
following paragraph.
At the time of writing of this report no data was available on leakage
rates for the materials and construction proposed for use. An expression for
the weight penalty was therefore d&=rived on the basis of known flight para-
meters and presented as a functlol of the leakage rate. For this purpose it was
assumed that all leakage consisted of, and was replaced by, pure helium gas.
For a given leakage rate the total gas weight lost is given by
(gas weight lost) _ J/?rP 7'A (DP) no.
where:
leakage rate - efm$tp/ft2 psi
tp ? gas density at t s o'C, p >4 1 atm - lb/ft3
flight duration- minutes
A - surface area for leakage - ft2
Ap R pressure differential - psis;
By data from Reference 1
(additional weight de?*ar) - 1.53 (gas 'weight lost)
(total added wei& t) - 2.53 O,, r 7'. (pp) lbs.
From the configuration data of Table IV
(weight gas lost) - 71.5 0 x 103 lb.
(added bottle weight) a 10g 0 x 103 lb.
(total added weight) a 1810 x 103 lb.
If the structure is assumed to consist of rubber 4-mils in thickness,
extrapolation from International Critical Tables data gives
0 - 33.4 x 10-6 cu/ft` psig
(weight gas lost) - 2.4 lb.
(added bottle weight) - 3.6 lb.
(total added weight) s 6.0 i~.
However, the leakage rate stated for a somewhat similar material yields
results 50-times the above. In addition, it should be noted that this addi-
tional weight calculated above compensates for leakage by diffusion through the
material only and not through any holes which may be present due to construction
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ANALYSIS
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PAGE 22
REPORT NO. LJ-.0 6
MODEL I1 ZEL
DATE 1431/58
or damage. Loss of Helium alone due to this latter cause could run to the
order of 6.5 pounds/minute for one-square inch r hole. This would give a
total weight penalty of 16.5 pounds for each square Inch of holies per minute
of leakage time.
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TEST RESULTS,
MARRUARDT.:
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Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
 Declassified and Approved For Release 2012/05/31 CIA-RDP89B00709R000400810001-5 e 15
MAR%UARDT
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TEMPERATURE
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Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
 Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
ANALYSIS
PREPARED BY
CHECKED BY
REVISED BY
CONVAIR
A DIVISION OF GENERAL DYNAMICS CORPORATION
(SAN DIEGO)
SECRET
no ,affft wEm
Puel Pressure
Control Valve
I
Vapor
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Figure 16
SECRET
PAGE 38
REPORT NO. ZJ-026
MODEL I ZEE
DATE 10/31/58
Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
 Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5 s 39
aD
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Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
 Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
ANALYSIS
PREPARED BY
CHECKED BY
REVISED BY
C ONVA I R
PAGE 40
ZJ-026
HAZISI
10/31/58
A DIVISION OF GCNERAL DYNAMICS CORPORATION
J SAN DIEGO)
SECRET
REPORT NO.
MODEL
DATE
TABLE
Pratt & Whitney
8RJ-43D
Altitude
80,00
100,000
135,000
150,000
Mach No
3.0
3.0
3.0
3.0
Pt2/~rto
0.784
0.784
0.784
0.759
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601.9
601.3
599.5
618.17
gt2
0.780
0.307
0.0709
0.0387
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Tt2
623
703
854
917
-
2440
2685
2895
2856
f/A
0.0113
0.0133
0.015
0.015
Fiq
19,050
7,725
1,732
876
TB7C
0.689
0.760
0.831
0.902
Drag
979
387
99
58
Fi
18,071
7,338
1,633
818
ITan
0.726
0.800
0.881
0.967
Weight
1075
1075
1075
1075
Length
256
256
256
256
Weight/F1
0.048
0.119
0.535
1.070
Engine + Fuel Wt./n. 1.108 1.239
1.695
2.309
Pratt & Whitney
SRJ-43E (Alternate Design)
Altitude
80,000
100,000
135,E
150,000
No
Mach
3-0
3.0-
3.0
r
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2
0.784
0.784
0.784
0.7yg
W
Yei/&t2
601.9
601.3
599.5
618.17
Ut2
0.780
0.307
0.0709
0.0387
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Tt2
623
703
854
917
_
2840
3075
3285
3200
f/A
0.0143
0.0168
0.020
0.020
FM
22,600
9,145
2,059
1,025
TSFC
0.731
0.811
?.933
1.029
Drag
979
387
99
58
Fi
21,821
4x778
1,960
967
ITSFC
0.763
0.846
0.98
1.09
Weight
1057
1057
1057
1057
Length
252
252
252
252
Weight/Ft
0.039
0.998
0.439
0.890
Engine + Fuel lit. /Fi 1.149 1.28
1.729
2.285
SECRET
Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
 Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
ANALYSIS
PREPARED BY
CHECKED BY
REVISED BY
PAGE 41
REPORT NO.
MODEL
DATE
ZJ-026
HAM
10/31/58
CONVAIR
A DIVISION OF GENERAL DYNAMICS CORPORATION
ISAN DIEGO)
SECRET
Mm
Pratt & Whitney
W-
Altitude
60,000
100,000
135,E
Mach No.
2.5
2.5
2.5
0.89
0.89
0.89
z /&t2
Wag/V e
635.8
635.6
635.0
T
0.41
0.162
0.0374
6T2 'F
415
481
60
T *F
1600
1740
2160
f 7A
0.0071
0.00765
0.0104
FIT
7980
3110
790
Tsn
0.644
0.679
0.760
Dreg
34,40
1371
322
Fi
45
1739
468
ITSF'C
1.133
1.212
1.282
weight
1075
1075
1075
Length
256
2-,56-
256
Weight/11"i
0.193
0.503
1.87
Engine + Fuel wt. /Fi 2.173 2.541
3.895
by 2%
Pratt & whims
SRJ-43E (Alternate De
Altitude
80,000
100,000
135)E
M(a)ch No.
2.5
2.5
2.5
.
/
p
0.89
0.89
0.8g
V2
,
ig
V9
'/St2
635.8
635.6
635.0
r
*2
0.41
0.162
0.0374
'F
TT2
415
481
608
-
'F
-LYW
2094
2530
f7A
0.0093
0.010
0.0138
FN
10,100
3,965
950
TSFC
0.666
0.695
0.835
Drag
3440
137.
322
Fi
6660
2594
628
IT dFC
1.00
1.062
1.262
Weight
1057
1057
1057
Length
252
252
252
weight/Fi
0.129
0.332
1.370
Engine + Fuel Wt-/F1
1.877
2.117
3.360
For 2? angle of attack reduce thrust by 2% and increase &FC
SECRET
'?"" Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
 Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
ANALYSIS
PREPARED BY
CHECKED BY
REVISED BY
CONVAIR
A DIVISION OF GENERAL DYNAMICS CORPORATION
I SAN DIEGO)
PAGE
REPORT NO.
MODEL
DATE
42
ZJ-026
HAZEL
10/31/53
SECRET
TA 4
Pratt & Whitn
BRJ--
Fentaborane Fuel
2
3.0
.2
-
A1titu*
100
1
80
100
135
150
Pr2/!to
0.89
0.89
0.89
.784
.784
.784
.714
IFT2
.412
.163
.0375
.780
.308
.0709
.040
TT2 *F
415
479
6o6
623
703
85'4
917
N
635
8
6
635
635
601.9
601.3
599.5
613.9
wa
2
.
.
w 16/sec
2D1
76.5
16.6
324
123
26.7
14.7
P/A
.0158
.0176
.0217
.0245
.0278
.04
.04
.991
.985
.934.
.989
.982
.945
.918
TB OF
2420
2200
2400
2720
2870
3280
3173
9440
3821
877
20179
8467
1933.
1003
Tarc
1..213
1.272
1.482
1.354
1.453
1.992
2.112
3440
1371
322
979
387
99
58
IFN
6000
2450
555
20100
8080
1834
945
ITBFC
1.908
1.984
2.342
1.41^0
1.523
2.100
2.241
SPWT
.1733
.424
1.892
.052
.129
.567
1.101
Pr/Pi
499
694
3
582
5
2. u4
3.324
3.970
E +
3.
.
.
DC in
86
A Ft
L.
25.73
DD In
68.7
L/Vr
0.87
LN In
59.8
DE In
104
LIn
266
Wt ? Lbs.
1150
For 20 angle of attack reduce thrust by 2% and increase BF'C by 2%
SECRET
?"" ' Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
 Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
ANALYSIS
PREPARED BY
CHECKED BY
REVISED BY
CONVAIR
A DIVISION OF GENERAL DYNAMICS CORPORATION
(SAN DIEGO)
SECRET
TABLE
L5
Lz
L3 --
Pratt and Whitney
one- Gecxaetry
PAGE 43
REPORT NO. 'Z.J-026
MODEL HAZEL
DATE 10/31/58
r
SRJe.l3D
sRJ-43E
Alternate Design
Fuel
Pautaborae
Sr-1
SF?2
DZ In.
86
86
86
DT in.
68.7
67.3
70.1
DE In.
104
3.04
104
L1 In.
98.5
98.5
98.5
L2 In.
60
6o
60
L3 1n'
48
36
36
L4 In.
60
62
55
L5 In.
266.5
256.5
252.5
SECRET
Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
 Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5
ANALYSIS
PREPARED BY
CHECKED BY
REVISED BY
CONVA{R
A DIVISION OF GENERAL DYNAMICS CORPORATION
(SAN DIEGO)
PAGE 44
REPORT NO. ZJ-026
MODEL HAZL!I.
DATE 10i31/58
SECRET
RAE
1. Sioulkin, M. ?H., "mat der, Near the 8tagsation Point of t Body
of Revolution," Journal vf Av"nautiea1 Sciences, Vol. 10, P. 570.
2. Eckert, Ernest R. G. j, "Survey on That Transfer at High Speeds, "
WADC Technical Report 54-70.
3. Private communication with Glen E. McIntosh, Group Engineer, Boulder
Division, Beechcref t Research and Development, Inc.
SECRET
FORM lilt-A-I
Declassified and Approved For Release 2012/05/31 : CIA-RDP89B00709R000400810001-5