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DOWNGRADED AT 3 YEAR VIZMALS;

DECLASSIFIED WIER 12 YEARS.

DOD DIR 5200.10

DECLASF.

3 0 OCT 1989

AIR FORCE SYSTEMS COMMAND

UNITED STATES AIR FORCE

PRELIMINARY DEVELOPMENT PLAN

FOR

HYBALINE-FUELED ACENA SPACE VEHICLE

System No. SS-01-B

26 July 1963
DECLASSSED

DQO

c' ' .1".2: 3 0 OCT 1989

3 0 OCT 1989

FOREWORD

This document presents a proposed development program for a Hybaline

fueled Agena Space Vehicle in response to Hq USAF message AFRSTD

76993, dated 9 July 1963.

The program is based on the results of research efforts by Union

Carbide Chemicals Campany, South Charleston, West Virginia, and

the Rocket Propulsion Laboratory, RTD, AFSC, Edwards AFB, California,

and results of a feasibility demonstration program conducted by the

Lockheed Missiles & Space Company, Sunnyvale, California, and the

Bell Aerosystems Company, Buffalo, New York.,

iii
TABLE OF CONTENTS

Page No.

Foreword iii

Section 1 Program Summary 1-1

Section 2 Schedules 2-1

Section 3 Program Management 3-1

Section 6 Acquisition 6-1

Section 11 Financial 11-1

Section 12 Requirements 12-1

Section 13 Authorization 13-1

Section 15 Security 15-1

iv
SECTION 1

PROGRAM SIMMARY

Hybaline (A-5) is a new high-energy rocket fuel developed by the Union

Carbide Chemicals Company which can provide si3nificant increases in

performance of the Agena space vehicle with only minor modifications

to the present design, with the exception of the engine.

Typical missions have been calculated to illustrate the performance

gains (See Figures 6.1, 6.2 and 6.3). Increments of 500 to 750 pounds

in burnout weight, which can be interpreted as increments in payload

and mission peculiar equipment, are available using the Improved Thor

launch vehicle. Using the Atlas launch vehicle, typical increments

in burnout weight are 500 to 900 pounds. The increased performance

can also be used to increase orbit altitude. For example, Program A

(Figure 6.3), using an Atlas launch vehicle, could increase orbit

altitude from 2000 to 3500 miles.

The modified engine will have 18,000 pounds thrust using IRFNA as the

oxidizer. It will have multi-start capability with a simplified start

system. and a maximum burn time of from 225 to 250 sec, depending on

which of the propellant tank configurations is selected. Engine specific

impiuse (Tsp) is estimated op be 325 sec under vacuum exit conditions

with an expansion ratio of 45. This exceeds the present UDMH/IRFNA

performance by more than 30 sec.

Configu ration No. 1 for vehicle structure is based on a minimum

modification concept.. The only changes are structural beef-up to

carry the higher payload and a minor relocation of the inner bulkhead

in Ih propellant tank to provide the proper ratio of fuel and oxidizer

wil i.rlrease in total volume.

Configuration No. 2 is based on a near minimum modification concept.

In w.HiLicn to:structural beef-up, the propellant tank volume is

aupoximately 15% in order to increase the propellant load

acleve increased performance gains within the constraint

,+:u14,,,f3 concept. This increase in volume is achieved by

iameter from 60" to 63" without changing the length.

e in diameter does not cause any change in handling

. qhange in AGE (for either alternative structure)

fuel system. This change would involve the

• rressurization system, which is not expected to

DOWNGRADED AT 3 YEAR INTERVALS;

DECLASSIFIED AFTER 12 YEARS.

DOD DIR 5200.10

1-1
The pacing item in this proposed program is the availability of

sufficient quantities of BYbaline fuel for engine development. The

lead time for a full scale production facility is approximately 12

months. Prior to that time, the only fuel available is from a small

pilot plant which has a capacity of only 1200 pounds per month.

SECTION 2

SCHEDULES
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SECTION 3

PROGRAM MANAGEMENT

No changes are anticipated to the existing Ageaa management structure.

Primary responsibility for program management rests with the Agena

Program Office (SSVA) within Hq Space Systems Division (AMC). The

prime contractor is Lockheed Missiles and Space Company, Sunnyvale,

California. LMSC subcontracts rocket engine development and production

for the Agena program to the Bell Aerosystems Company, Buffalo, New

York.

Management procedures, including configuration control, have been

fully implemented during the acquisition of SS-01-A and SS-01-B and

no charges are anticipated.

3-1
SECTION 6

ACQUISITION

6.1 System Description and Performance 

The Agena space vehicle is a versatile standardized vehicle which

can carry a wide variety of payloads into space, either as a second

stage booster or an orbital vehicle. Approximately thirty optional

kits are available for tailoring the Agena to specific missions.

First stage bor'st is provided by either Atlas or Improved Thor

standard launch vehicles. Launch facilities are available for both

the Atlantic Missile Range and the Pacific Missile Range.

Two configurations are presented for consideration. Both are

based on an 18,000 pound thrust Hybaline/IRFNA rocket engine which

is a minimum-change adaptation of the present Agena engine. The

minimum change concept has been applied even more stringently to the

other subsystems of the vehicle in order to avoid changes, insofar as

possible, in factory tooling, AGE, spares, launch equipment, and test

equipment.

Configuration No. 1 is essentially the same as the present Agena

except that the vehicle structure must be strengthened slightly to

carry the higher payloads, and the inner bulkhead of the -orrmellant

tank is relocated to optimize the oxidizer/fuel ratio.

Configuration No. 2 is a near-minimum change configuration, - but

one i17;ho-tant conession is made in order to increase performance --

the H;t.n 1.1-1k is enlarged by approximately 15*.

,Viapta!ion of the Agena vehicle for use with,Hybaline fuel will

:e ( , erational concept. It will, however, provide a

rya:7e in performance which is illustrated in Figures 6.1,

, Increase in performance can be used to increase

or payload, or, in some cases, it will allow

a er and less expensive Thor launch vehicle

:gena vehicle for use with Hybaline rocket fuel

ructural redesign in order to accommodate the

and increased thrust. The degree of redesign

11-formance increase desired.

6- 1
Configuration No. 1, being the same size and shape as the present

Agena, requires only a modest increase in strength, generally by

increasing thickness ofmaterial, with very little impact on tooling

and test equipment. In many cases structural components can be requalified by similarity. It will be necessary, however, to investigate

the requirement for additional thermal shielding in the vicinity of

the engine in order to cope with the higher combustion temperature.

Configuration No. 2 requires the same structural redesign as No. 1

plus redesign of the propellant tank. In order to obtain the additional

volume without impact on vehicle length or handling equipment, the

tank diameter is increased from 60 inches to 63 inches. While the

tank redesign and qualification is simple and straightforward, it does

require additional engineering, testing, and tooling as reflected in

the increased vehicle development costs. In addition, the increased

propellant weight will require some increase in skin guage on the

booster adapter.

A preliminary investigation of the impact of increased payload

and/or increased propellant weight on the launch vehicle indicates

that the Improved Thor will have to be strengthened in either case.

This effort is estimated at $1,900,000 and will probably require

eight months. Atlas, on the other hand, can carry the increased load

provided a low-g trajectory is used for Configuration No. 2. (See

Figure 6.4).

6.2.2 Propulsion

Based on the favorable results of the feasibility program at BAC,

we stronly recommend the immediate initiation of an engiL- :vvelopment

program. A prime design objective in this program will be the use of

YLE81.-BA-11. ani KR81-BA-13 design requirements and hardware configuration

i.o inu-e LwerchanGeability with these engines. This development proCruzn 1;Thll w' the following:

Th-11;•. Chamber

• e and full scale evaluation of thrust chamber

to optimize parameters of mixture ratio, combuss hrust coefficient (CF ), and chamber film coolant.

ptimum thrust chamber and nozzle extension

, erformance and durability at sea level (at

l'nold Engineering and Development Center).

6-2
b. Gas Generator

Develop a Hybaline/IRFNA injector suitable for use in

the gas generator (GG).

Determine injector performance, operating temperature,

solids buildup, duration, and restart capability.

c. Turbine Pump

r (1) Determine the effect of mixture ratio on the propellant

pumps •

Determine the effect of solids in the GG exhaust on the

turbine blades.

Evaluate a new pump-inducer combination capable of

operating at a lower net positive suction head.

(4) Evaluate a power take-off pad to drive the hydraulic

actuators.

d. Valves and Controls

Develop simplified

propellant valve.

and more reliable gas generator

Evaluate faster opening and closing main propellant valves

to reduce propellant pre- and post-flow losses and associa•ea start-up

and tail-ofT total tmpulses.

A:flembly.

A: H,"%c engine mock-up.

E•a'.-,;ate complete engine performance and durability at

9AC) 1n,! at altitude (at AEDC).

'hcoretical engine and propellant performance.

Unit

an auxiliary power unit to drive the turbine

Hyballne/IRFNA gas generator proves to be

• :red that in the eighth to tenth month of the

mm, a decision will be made as to which type

yie/IRFNA or some other) would be carried to

6-3
The schedule for the above engine development program (Figure 6.5)

is predicated upon the present production capability of Hybaline

(Figure 6.7). The schedule calls for completion of PFRT in twenty-one

months from go-ahead. It should be noted that the engine development

progrem could be completed in fourteen months if adequate quantities

of fuel were available.

6.3 Background, Difficulties and Approach 

The family of storable liquid rocket fuels known as the Hybalines

were developed by the Union Carbide Chemicals Company during the period

of late 1960 to 1962. This new class of propellants consisting of

aluminum borohydride adducts of organic nitrogen compounds, such as

primary and secondary amines, has the desirable characteristics of

relatively high performance, along with the physical properties required

of an earth or space storable fuel. On the basis of small (100# thrust)

motor firings at the Air Force Rocket Propulsion Laboratory in late 1961

and early 1962, the Hybaline A5 material evolved as the best candidate

from the standpoint of high performance and optimum physical properties.

The characteristics and properties of Hybaline A5 are shown in Figure 6.6.

Because of the relatively new nature of Hybaline A5, an experimental

evaluation of its characteristics was essential to determine its actual

performance and operational behavior.. To accomplish this task, the Air

Force Rocket Propulsion Laboratory at Edwards AFB, in November 1962,

procured a quantity of this material sufficient to permit thrust chamber

testing at a thrust level of approximately 5000 pounds and to conduct

supporting laboratory work in the areas of stability, chemical analysis,

and material compatibility. The A5 that was procured was .-- intended

to represent a final fuel but was considered a prototype material which,

if the experimental data was promising, would be further developed into

a well caracLeri :7,ed . propellant. Therefore, it was anticipated that

s)me u. arise concerning the purity and physical properties.

Tn€ pr(;qTram consisted of a performance testing program, a

comb:: ; : Fability program, and laboratory tests. The testing was

- fl any operational problems such as safety, handling,

the Rybaline as well as providing specific

Hybaline A5-N204 combination.

program, which began in December 1962, was

phase consisting of mixture ratio and injector

,b ish performance trends, and a determination of

'ensure upon combustion efficiency. The second

, rmance based on the trends established during
the first part of the program and would provide data on heat flux

and nozzle efficiency. In the initial phase of this program, over

96 tests have been conducted at the 2000 pound thrust level using

Hybaline A5 and N204. These tests have served to evaluate twelve

different injector patterns. The results of these tests show that

the best efficiency obtained at the present time lias been approximately

89% Isp which is equivalent to a delivered Isp of 257 sec at 500 psia

chamber pressure and sea level exit pressure or 275 sec at 1000 psia

chamber pressure and sea level exit pressure. Mixture ratio traverses

have indicated lower performance efficiencies at the fuel rich mixture

ratios, with the drop off in efficiency occurring around a mixture

ratio of 1.0 to 1.2. Varying the chamber pressure over levels of 300,

500, and 1000 psia did not appreciably affect the performance efficiency

for a given injector. Comparative tests between N204 and IRFNA with

hybaline using the same injector pattern have shown essentially equal

performance efficiencies for both combinations. This indicates that

it should be possible to extrapolate the trends established with the

A5-N204 combination to the A5-IRFNA combination.

A measure of the relative stability for the A5-N204 combustion

process was accomplished by firing the propellants in a pulse motor.

This is a combustion chamber into which calibrated pulses can be

introduced. The size of the pulse required to initiate combustion

instalability provides a relative measurement of the stability of the

combination and the injector pattern involved. The hardware utilized

in these tests had previously been used to test the N204-50/50 combination so that a baseline for comparison was available. The -e clats of

the tests indicated the stability of A5-N204 is considerably ,etter

than N204-50/50 for the conditions tested. The testing covered a

chamber pressure range between 300-800 psia and a mixture ratio

between 2.2 to 4.1 using an injector pattern with a performance of

approx 'sp. Thus the injector performance was fairly

good compare'l to the best injector tested so far at 87% Isp.

The atorat.ory

Hyl, a!Ine - have

wi.1 h a: .

i.a

'

0( •

tests and operational experience gained from the

iemonstrated that the A5 is relatively easy to work

major problems as far as operating a test system

nc and transfer of the Hybaline A5 on the test

ijshed on a routine basis with no difficulty.

offers no hazard from toxicity or danger of

:iust produced with either N204 or IRFNA contains

1Aeoirable products. Hybaline A5 has been

i 4 11-most of the commonly used materials in

, ,4;h Hybaline A5 is sensitive to moisture

REN:144•
and reacts violently with water, there has been no operational

difficulty from this aspect providing proper precautions are observed.

Several incidents have occurred where impurities in several batches

of the A5 have resulted in problems in the test system, but it appears

that modifications in the production process can eliminate these

difficulties.

On the basis of the promising results from early testing at AFRPL,

an interest was generated in applying the Bybaline A5 fuel to the Agena

propulsion system. This interest was further stimulated by the apparent

suitability of Hybalirie A5 in Agena with IRFNA as an oxidizer. The •

chemical and physical properties of A5 make it well suited for application to a system such as Agena. The theoretical performance with IRFNA

offers a considerable improvement over the existing IRFNA-UDMH propellants.

In addition, the density of A5 is similar to UDMR, the chemical nature

in terms of compatibility is reasonably similar to UII4H and the mixture

ratio for Hybaline A5-IRFNA is in the proper range to make a substitution

of A5-IRFNA into the Agena system appear quite attractive. Besides the

chemical and density similarities between A5 and DEM, the retention

of IRFNA as the oxidizer might permit the use of IRFNA as a coolant

in the same manner as it is used in the present , Agena engine.

In January 1963, discussions between AFSSD, AFRPL, Lockheed Missile

and Space Company, and Bell Aerosystems Company were held to explore

the possibility of test work with Hybaline A5-IRFNA directed towards

an Agena application. These preliminary studies and discussions led to

a program initiated on 20 March 1963 to demonstrate the feesibility of

utiliz4 ng Hybaline in the YLR81-BA-11 engine with minimum _ranges to

the present hardware. The program was organized into five parts to

investigate the areas of combustion evaluation, gas generator and

turbine rump evaluation, thermal transport properties, chemical analysis,

an,J. :‘nanges t.o the engine.

The task on Combustion Evaluation had several objectives: First

of all wa:; t del ermine the combustion efficiency of the A5-IRFNA

Ac,ena type hardware. The present results indicate a

can be obtained from Agena type injectors in

?-x1.- • chambers. The heat rejection rates were also

water cooled chambers. This was an important area

:‘ates that are too high would preclude the possibility

- with IRFNA using the present Agena type aluminium

f the work showed that heat rejection valves were

e i values and were approximately the same for

..filues obtained indicate that cooling with IRFNA

the same margin of safety as presently obtained.

6-6
The presence of solids in the exhaust posed a potential loss in

nozzle efficiency. To investigate this area tests were performed in a

diffuser with an Agena thrust chamber using a fully expanded nozzle to

an area ratio of 45 to 1. The average Of delivered from three tests

97.4% of theoretical. Cf for the UtiMH-IRMA engirt is 97.1%. This

is an encouraging result which apparently indicates no appreciable

losses are occurring during expansion through the nozzle. The overall

results from the thrust chamber performance tests have indicated it

is possible to operate an existing Agena type chamber using A5-IRFNA

with the IRFNA acting as a regenerative coolant and to deliver thrust

chamber impulse at attitude of over 327 sec.

The purpose of the gas generator and turbopump work was to

identify problem areas involved in operating a gas generator using

A5-IRFNA as propellants and in driving a turbopump with an exhaust

containing an appreciable amount of solids. The choices of either

oxidizer rich gas generator operation or fuel rich gas generator

operation were possible. Oxidizer rich operation produces less solids

than fuel rich operation, but results in a more severe environment for

the turbine materials because of the hot oxidizing gases. The results

of testing under both conditions indicated the turbine and pump could

operate satisfactorily for durations of over 30 seconds with no evidence

of harmful effects. This was originally considered to be a critical

area and it had been predicted that either the solids or the oxidizing

exhaust would have an adverse effect on the turbine. Although the

turbopump itself caused no problems, a problem was encountered with

burnouts of the gas generator under both oxidizer rich and rich

conditions. The problem appears to be one of injector design and it

is believed that the burnouts would be eliminated by utilizing a

different injector pattern.

In :7-umm3xy, it appears possible to operate a turbopump with either

a fuel rich gas generator. or an oxidizer rich gas generator provided

a satisfactory gas generator can be developed.

Tr:.(- Irmal transparent properties was undertaken to

invest- .he:mal stability of A5 under conditions of heat input

sue!) &m . , ered in injector manifolds or coolant passages.

In 1,, :1.1. of using A5 as a regenerative coolant was

en.acArl :hawed A5 would decompose under both static

till.ions with the evolution of gases and production

,-- . 11.ved whether this thermal decomposition was

impurities present in the A5, but is believed

, jes would reduce the decomposition rate. The

n marginal regenerative coolant with a

valent to that of IRFNA.

6-7
The effort on chemical analysis was undertaken to determine the

" quality of the fuel", to establish quick analytical techniques for

determining quality, and to determine physical properties. The study

i nd i cated that existing analytical technique need to be improved to

accurately determine quality control of A5. The work effort in this

area is still in progress.

"Difficulties and Approach"

The experimental data and experience obtained from the BAC feasibility program and from the AFRPL studies with Hybaline A5 indicate that

the development program to incorporate the Hybaline A5-IRFNA propellants

into the Agena system is feasible. Certainly no problems have been

encountered to date which would rule out the possibility of such a

system. On the other hand, the evidence is strong that •ybaline can be

incorporated into an Agena system with a minimum of major design change.

In summary, it can be stated that the characteristics of Hybaline A5 in

teris of physical properties, handling data, operating experience,

thrust chamber performance, combustion stability, and heat transfer

have been experimentally investigated and the results are encouraging

for an Agena development program.

Some additional work is required in several areas to support a

development program. These areas consist of analytical procedures on

the Hybaline A5 propellant to more fully define "quality" of the

propellant and to check consistency from batch to batch and run to run.

Other areas concerning the fuel are to more accurately establish

theoretical performance values, to develop more complete physical

property data, and to optimize the manufacturing process. Furt'n- test

work is required in the engine area to optimize thrust chamber performance, achieve durability of the hardware, develop an A5-IRFNA GG

injector. and demonstrate integrated engine operation.

6.4 Test n'a

Category and II testing, as defined by AFR 80-14, will not be

required f(., pl-,:nt.11. However, development testing requirements

include auita Cruz Test Base.

of tt:e

be H..:

i • luired to demonstrate the altitude performance

Later, of the complete propulsion subsystem.

pper stage propulsion systems has shown that

performance problems which cannot otherwise

&ITC will provide the most reliable data

- :11 subsystem performance.
Testing at the Lockheed Santa Cruz Test Base is required to

demonstrate integration of the propulsion subsystem into the complete

Agena vehicle. It will also serve as a training program for LMSC

launch base crews and as a verification program for operating and

handling procedures. This is particularly important for handling and

storage of Hybaline fuel because of the necessary modification to the

ground fuel system.

The schedules for these tests are shown in Figure 2.4.

6.5 Hybaline Production Data

The only additional facilities required for this program are

production facilities for Hybaline fuel. A survey of existing facilities indicates that the USAF Pentaborane Plant (Installation No. 8699)

at Moskogee, Oklahoma, is the only existing plant which could be converted for Hybaline production at reasonable cost. The availability

of Hybaline fuel is the pacing item for the development schedule of a

Hybaline fueled Agena. The pacing item for conversion of the Muskogee

plant is physical access for Union Carbide facility design engineers

to the Muskogee plant and to the necessary detailed design data on the

Muskogee plant. The generalized data currently available to them is

sufficient to assure feasibility of the conversion, but is not sufficient

to allow them to start detailed design work.

The USAF plant at Muskogee, Oklahoma, is presently being maintained

in a stand-by status by Callery Chemical Corporation under ASD Contract

No. 33(600)-41757. The decision has been made, however, to deactivate

the plant. in October 1963. A portion of this plant, acquired in 1962

at a cost of approximately $32,500,000, could be converter ' an

initial production capacity of 2,000,000 pounds per year at a cost of

approximately $2,700,000. The relative cost of conversion of this

plant as copared to augmentation of one of the U.C.C. plants is mislealiur, That part of the Muskogee plant not used for production of

Hybaline the peculiar capability of being easily converted for

productin of sodium borohydride, which is the critical raw material

for pro:iuion of Hybaline. The current national capacity for producUon fY i,ydride is sufficient for no more than 2,000,000

plan' ter year. Construction of a new sodium borohydride

par; we expensive than conversion of the Muskogee plant.

requirement for Hybaline for the Agena program ,

Funds per year, it is reasonable to assume that

A' this new fuel will stimulate usage on other

capability for future expansion should not
Union Carbide Chemicals Company, developer for Hybaline fuel, has

made a study which indicates that they could augment facilities at

either of two existing plants in south Texas (Seadrift and Brownsville)

to produce up to 2,000,000 pounds of Hybaline per year. Both land and

major utilities are available at these sites. Cost of this augmentation

is estimated at $3,700,000 and would take about 14 months from go-ahead

to initial production delivery.

Consideration of the above factors leads to the obvious conclusion

that the Muskogee plant conversion is preferable to augmentation of one

of the U.C.C. plants. This choice, however, demands access as soon as

possible to the detailed design data on the Muskogee plant. In addition,

it is also concluded that a full go-ahead on conversion of the Muskogee

plant must be given at the very beginning of any further development

effort, however limited, unless we are prepared to accept a day for day

slippage of the over-all program.

Production capability in either the converted Muskogee plant or

an augmented U.C.C. plant builds up rapidly after completion of the

facility, as shown in Figure 6.7. Alter three months, production

capability exceeds requirements for testing. At that time it is anticipated that production would level off at about 506 of maximum for a

few months and then drop back to about 25% of maximum. Fortunately,

the production process is largely automated so that these large fluctuations in production rate can be economically accommodated.

The cost of Hybaline is strongly affected by production rate, as

shown in Figure 6.8. Manpower costs are essentially a function of time

rather than rate, due to the high degree of automation. In addition,

raw materials, particularly sodium borohydride, are subject to rispAntity

discounts. However, since the national capacity for production

sodium borohydride is only sufficient for 2,000,000 pbunds of Hybaline

per year, there q tr a disf:,ontinuity in Figure 6.8 at this point.

Above pnis per year, it would be necessary to obtain

additional n fuilities for sodium borohydride. This can be

accomplished etaity at the Muskogee plant with only minor modification

to existing. i.es. Construction of new facilities would be quite

expensive:. in Figure 6.8 are therefore based on production , .. ) 1 W Muskogee for Hybaline rates in excess of

2,000,- c.

!:tie Hybaline production facility, the only

sot-, r :mall pilot plant operated by Union Carbide

This pilot plant, originally designed

month, is currently operating at 1200

development plan is based on the 1200
pound rate in the pilot plant, it is estimated that this output could

be further expanded to 1800 pounds per month at a cost of approximately

$30,000. Lead time for this modification would be four months. Production at the 1200 pound rate could continue for 3 months, at which

time the plant would have to be shut down for two weeks for installation of the new reactors. The increased output would be more efficient,

thereby reducing the cost of Hybaline from $80 per pound to $70. The

net gain in production capability prior to completion of the full scale

production facility would be 4200 pounds, which would accelerate the

engine development effort by about one month at a net cost of $170,000.

In constrast to expansion of the pilot plant, acceleration of

the construction schedule for the full scale production facility would

result in a direct saving of $74,000 per month due to decreased cost of

Hybaline and at the same time allow earlier completion of the over-all

nrogram by an amount almost equal to the acceleration of the Hybaline

production facility. Management attention should therefore be directed

toward expediting completion of this facility. There is also the

possibility of a favorable trade-off between premium construction costs

(overtime) and the reduced cost of fuel, although this trade-off is

complicated by administrative problems because two different categories

of funds are involved.
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PHYSICAL PROPERTIES OF HYBALINE A5

Name Mixed methylamine-dimethylamine

monoaluminum borohydride

Structural Formula, CH3NH2•Al(BH4)5 53% by wt

(CH3 ) 2 N14• 411(Bil4 ) 5

477. by wt

Empirical Formula C1.47N1117.94A1B3

Molecular Weight 109.19

Density, ga/m1 20°C 0.734

Vapor Pressure, min Hg 25°C 3 (a)

Freezing Point, °C -50.0 (b)

Boiling Paint, °C 263

Viscosity, cp 20°C 6,78

Thermal Stability, °C 90-100 (c)

Shock-Sensitivity, Kg/cm 120.0 (d)

Air-Sensitivity Oxidizes slowly in air without ignition

Beat of Formation, Kcal/gm-mole +11.0 (e)

Surface Tension, dynes cm 48

Auto-combustion Temperature, 298

Enthalpy, cal/gm, 10°C 6.0

Specific Heat, cal/gm, 20°C 0.621

Critical Temperature, °C 350 to 500

Heat of Vaporization, 25°C cal/mot 4990

Thermal Conductivity, cal/sec.

cm °C at 20'C 340 x 10-6

re compound.

1 -50°C.

5.7 per cent decomposed after 5.5 hours at 215°C.

Otin Mathieson drop weight tester.

+17.7 ANc ° for aluminum borohydride, and the

„amine lilands.

, rptrical method. Dovenammo AT 3 YFAR 1141112VAi

DECIMETRE!) AFTER 12 YEARS.

tna 5200.10
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SECTION 11

FINANCIAL
.

SECTION 12

REq1=MENTS

Directors of Program 162 and SAFSP programs have expressed great

interest in performance increases of the order of magnitude of the

increased proposed in this plan. Justification of requirements

will be left to the program directors concerned.
SECTION 13

AUTHORIZATIONS

FM HQ USAF WASH DC

TO RUE AFF/AFSC ANDREWS AFB MD

INFO RUEAFF/AFSC (SCSB) (COLONEL CRISTRADORO) ANDREWS AFB MD

RUWBBK/SSD (SSZA (LT COLONEL LEBECK) LOS ANGELES CALIF

RUWHBF/ROCEET PROPULSION LAB (DGGD) EDWARDS AFB CALIF

AF GRNC

NT

UNCLAS FROM AFRSTD 76993

SUBJECT: HYBALINE PROGRAM. REFERENCE OUR MESSAGE 75215, 1 JULY 1963,

SAME SUBJECT. UNDER SECRETARY MCMILLAN HAS CLARIFIED HIS REQUEST FOR

INFORMATION ON THE HYBALINE PROGRAM. HE WOULD LIKE A PRELIMINARY

DEVELOPMENT PLAN PUT TOGETHER FOR AlaBALINE FUELED AGENA WHICH INDICATES

THE QUANTITIES AND TIME PHASING OF HYBALINE REQUIRED. HE HAS STATED

THAT THIS PLANNING DOES NOT NECESSARILY INDICATE THAT WE WILL PROCEED

WITH AN ENGINE DEVELOPMENT PROGRAM EVEN IF THE PRESENT BELL EFFORT IS

SUCCESSFUL. DR. MCMILLAN HAS ASKED FOR A BRIEFING ON TEE PRELIMINARY

DEVELOPMENT PLAN AROUND THE END OF JULY. PLEASE CONFIRM DATE THIS

BRIEFING WILL BE AVAILABLE.

P 1520457

FM DCMSF ANDREWS AFB MD

TO RUVHBK/SSD WS ANGELES CALIF

RUEiBET BOLLING AFB WASH DC

11170 =1-1E,/ROCKET PROPULSION LABORATORY EDWARDS AFB CALIF

BI

UNC LAS MSFA 15-7-22

FOR SS'-; (COL BLUM). RTD FOR RTNP; RPL FOR DGGD. THIS CONFIRMS 12 JULY

171,0N BETVEEN COL BLUM (SSV) AND MSFAM PERSONNEL. REQUEST SSD

170NIU •ITfl USK' MESSAGE AFRSTD 76993. FURTHER REQUEST RTD ASSIST SSD

IN FP.'t, -YION J,ND PRESENTATION OF THE REQUIRED PLAN. THIS OFFICE .

11IEFING WILL BE AVAILABLE' WEEK OF 29 JULY. WE WILL

--.0DINGLY AND ADVISE OF PRESENTATION SCHEDULE AS SOON
SECTION 15

SECURITY

The chemical formula, detailed physical properties, and detailed

performance data on Hybaline (A5) rocket fuel are confidential.

The performance characteristics of the Agena Space Vehicle, in

general, are confidential. When the performance of the Agena with

specific payloads is presented, the security classification of their

combined performance is controlled by the security classification of

the using program except that it will never be less than confidential.