FEASIBILITY STUDY THROTTLEABLE LIQUID BIPROPELLANT ROCKET ENGINE

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP89B00709R000400840007-6
• Release Decision: RIPPUB
• Original Classification: K
• Document Page Count: 37
• Document Creation Date: December 22, 2016
• Document Release Date: January 11, 2011
• Sequence Number: 7
• Publication Date: March 20, 1959
• Content Type: REPORT

STAT
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Report X-394 Copy No. 1
FEASIBILITY STUDY
THROTTLEABLE
LIQUID BIPROPELLANT ROCKET ENGINE
20 March 1959
HUGHES TOOL COMPANY--AIRCRAFT DIVISION
Culver City, California
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Frontispiece.
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TABLE OF CONTENTS
Page
SUMMARY 1
I. INTRODUCTION 2
II. PRELIMINARY DESIGN OF ENGINE 4
A. PROPELLANT SPECIFICATION 4
B. THRUST CONTROLLABILITY 5
C. BURNING DURATION 6
D. DEVELOPMENT PERIOD 6
E. EXHAUST EMISSIVITY 7
F. PERFORMANCE OF PRELIMINARY DESIGN ENGINE 8
III. INFRARED DETECTION OF EXHAUST FLAMES 12
IV. EXPERIMENTAL ROCKET ENGINE STATIC FIRINGS 25
BIBLIOGRAPHY 33
Figure
LIST OF ILLUSTRATIONS
Page
Frontispiece. Static Firing of Model 315 Test Engine
on RFNA and N2H4
11
1
Throttleable Liquid Rocket Engine, R19
9
2
Design Specifications, R19A and R19B
10
3
Vehicle Performance With R19 Rocket Engine
11
4
Model 315 Propellant Evaluation Engine
26
5
Model 315 Thrust Stand
27
6
Model 315 Feed System
28
7
Rocket Exhaust Emissivity Test Setup
29
8
Summary of RFNA-N2H4 Static Test Firings in Model 315
Rocket Engine
30
9
Effect of RA-N2H4 Exhaust Flame on Standard Source
Adsorption Profile
32
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SUMMARY
The feasibility of a throttleable liquid bipropellant rocket engine suitable
for a long range manned vehicle was investigated. Results of the stuoily indicate
that such an engine is indeed within the present state of the art and could be
developed to qualification status in a two-year period. The detection of such an
engine by present ground based IR systems would be extremely difficult, if not
impossible. The experimental phase of this program, in which several of the
capabilities of this type of engine were to have been demonstrated was terminated
after an accumulated five minutes of firing time on the test engine. This termina-
tion WAS necessitated by the limitations on program time and funds.
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I. INTRODUCTION
From a vehicle standpoint, the choice of a propulsion system depends
upon its performance, reliability, availability, and cost. It is the purpose of
this study to establish the status of these factors for a liquid bipropellant
rocket engine, so that a vehicle designed around such an engine can be compared
with vehicles designed around other types of engines, specifically, airbreathers.
The vehicle comparison is beyond the scope of this study, but in general
the basis for vehicle performance comparisons is range. In this regard, rocket-.
powered vehicles are at a disadvantage in travel through that part of the
atmosphere that can support combustion. In the vehicle comparison at hand,
however, the'performance criterion isnot solely range -- there is also an implied
requirement for escape from detection. Here, the rocket-powered vehicle is at
an advantage in comparison with an airbreather. This is so because the very
parameter that a vehicle seeks to maximize to escape detection, namely, altitude,
is that which imposes detection compromises in the case of an airbreathing
engine.
Thus, the Mach 3 or faster airbreathers capable of operation at altitudes
over 100,000 feet must resort to use of pyrophoric fuels to sustain combustion.
The exhaust products of such fuels, bearing as they do solid constituents in
relatively large amounts, can serve as rather efficient energy radiators in the
IN spectrum. This condition is not necessarily so for a liquid rocket engine,
inasmuch as its fuel constituent is not compromised by the necessity of combustion
in thin air. Thus, rocket propellants can be tailored for minimum exhaust
emissivity, and as a result a rocket-powered vehicle could enjoy a performance
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advantage over a vehicle powered by an airbreathing engine. This situation
depends, of course, on the importance of the detection requirement, an evaluation
that is, as previously stated, beyond the scope of this study.
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II. PRELIMINARY DESIGN OF ENGINE
The following criteria were set up as basic requirements for the liquid
rocket engine in this feasibility study:
A. Storable, high performance propellants
B. Thrust controllability throughout a large turn-down range
C. Long burning duration
D. Capability of development in two-year time period from present
r'
E. Nonemissivity of exhaust products.
These requirements in toto define an engine that is, at present,
nonexistent. However, as will be discussed, this nonexistence is at present
the result of lack of stated need rather than of fundamental developmental
L- problems. Individually, each of these requirements has been satisfied in various
operational or developmental engines. Since none of these requirements are
mutnAlly exclusive, it is definitely within the state of the art to design a
liquid rocket engine satisfying each and all of these requirements.
A. PECTILLANT SPECIFICATION
L_
The storability requirement rules out cryogenic propellants and restricts
the choice of liquid propellant oxidizers to two oxidizers: nitric acid and
r nitrogen tetroxide. In the performance study of the subject system that follows,
nitric acid was selected over nitrogen tetroxide on the basis of greater
development background and superior characteristics as a coolant. However,
nitrogen tetroxide is becoming increasingly acceptable as a storable oxidizer,
L-
and actually is somewhat superior on a performance basis. The performance study
thus reflects a certain degree of conservatism as regards range as a result of
the use of nitric acid performance rather than nitrogen tetroxide (about 5 percent).
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The choice of storable fuels is more extensive, but the restrictive
requirements in this regard relate to hypergolicity and emissivity of exhaust
products. Since it is considered that the advantage of bypergolicity in safe
and reliable engine operation outweighs the hazards associated with tankage
integrity, bypergolicity of propellants has been set out as a further requirement.
This decision rules out hydrocarbon fuels and focuses attention on the bydrazine
family. Considerable development background exists on both RFNA-N2H4 and
RFNA-UDMH systems, and there is little to choose between them on a performance
basis. However, considerations of exhaust emissivity may favor the choice of
N2H4 over the more easily handled UDMH.
B. THRUST CONTROLLABILITY
This requirement is probably the one creating the most problems from an
engineering standpoint. The developmental effort required for a controllable
thrust rocket engine has in the past motivated system designers to accept compro-
mise solutions involving multiple chambers, or preprogrammed boost-sustain
thrust patterns.
For the subject application, however, in which thrust must be capable of
matching drag throughout a relatively long period during which vehicle mass is
constantly decreasing, the aforementioned design compromises are not acceptable.
A truly modulating thrust control is now required. Fortunately such an engine
presents no fundamental problems incapable of solutionl. The main design com-
promise occurs in optimizing the feed system pressure drop-cooling flow rela-
tionship for the extremes of the thrust range. That is, the combustion chamber
1 The Naval Ordnance Test Station, Inyokern reports such an engine in development.
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coolant passages must be small enough to create coolant velocities sufficient
to maintain safe wall temperatures at the low thrust flow rates, and yet be large
enough to allow acceptable pressure drops at the high thrust flow rates.
C. BURNING DURATION
The subject vehicle trajectory requires burning durations of approximately
25 minutes. This requirement, as well as that of (B) above, represents an exten-
sion to the present state of the art. It is an extension, however, only in that
no requirement of this duration has heretofore existed. An increase in burning
duration in itself represents no fundamental difficulty, inasmuch as present day
regeneratively cooled thrust chambers operate in thermal equilibrium. Their
burning duration is limited solely by their available propellant supply.
D. DEVELOPMENT PERIOD
Restricting the development period to two years defines the propellant
choice more than it affects the hardware design. The hardware design itself
appears to be within the state of the art; at least, no hardware compromises
are envisaged because of the development period restriction. On the other hand,
performance increases, which are theoretically available if laboratory type
propellants were considered, are excluded by this requirement. Specifically,
storable high density halogen oxidizers appear to show promise for this applica-
tion -- promise that could be realized operationally in a few years if investiga-
tive and evaluative work were begun now. However, since propellant development
normally requires long lead times, the performance characteristics of the engine
have been limited to those available today.
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E. EXHAUST EMISSIVITY
Operational systems in the past have used carbon bearing fuels, which,
when used with either acid or LOX, give rise to long tail-plumes containing
incandescent carbon particles. The emissivity of the incandescent carbon
particles, being close to that of black boolly radiation, is 10,000 times greater
at 6,0000 K than the emissivity of the gaseous products of combustion. The
inference in such a comparative figure is that there should be no carbon atoms
(or at least no atoms producing solid particles in the exhaust) in the pro-
pellant system, in order to minize IR radiant energy. Thus, from an emissivity
standpoint, propellants could be qualitatively rated in the following order
(high intensity to low intensity):
(1) Propellants containing light metals
(2) Propellants containing chlorides
(3) Propellants containing carbon
(4) Propellants yielding exhaust molecules containing only
N2and H20 (RFNA-N2H4).
From the point of view of the possibility of detecting the exhaust emission,
much depends on the location of the detector. Projects to achieve very sensitive
IR detectors are now current for space environments. That is, the detectors
must be in space as well as the IR source, and, if so, detection as far as
1,000 miles out is hoped for. This subject is dealt with in more detail in
Section III. However, it can be stated here that for ground-located IR
detectors the IR radiation from gaseous molecules in space is attenuated by the
earth's atmosphere to such an extent that it would seem impossible to locate a
high-altitude source of the subject magnitude.
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F. PERFORMANCE OF PRELIMINARY DESIGN ENGINE
The foregoing has summarized the effect of the vehicle requirements on
the powerplant. These requirements have been taken into consideration in the
determination of design specifications for an applicable engine. The general
geometry and essential features of such an engine are shown in Figure 1. Design
specifications for two such engines whose turn-down ratios (4:1 and 2 1/2:1) span
the area of interest are shown in Figure 2. It can be noted that their performance
is essentially the same, although certainly the attainment of the lesser turn-down
ratio would constitute a simpler development problem.
For evaluation of this engine's performance in a vehicle, Figure 3 is
presented. Here, duration of powered flight is shown as a function of vehicle
mass ratio and vehicle L/D for propellant Isp's of 275 and 300 seconds. For the
sake of illustration, a range scale is superimposed on the time scale, in which
the equivalence is based on a vehicle velocity of Mach 4. (In any particular
vehicle there is, of course, a functional relationship between L/D and vehicle
velocity, and therefore to preserve the generality of the graph a family of
range scales should be envisaged.) It should be appreciated also that the
vehicle range denoted by this range scale is solely that of the powered portion
of the flight. The full flight profile would include the glide phase, which
extends range considerably at the altitudes and flight speeds of interest.
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Throttleable Liquid Rocket
\O
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THRUST CHAMBER
THRUST AT DESIGN ALTITUDE
CHAMBER PRESSURE
THRUST CHAMBER COOLING
PROPELLANTS
MIXTURE RATIO
INJECTOR HEAD
NOZZLE THROAT AREA
NOZZLE AREA RATIO
CHARACTERISTIC LENGTH L*
TYPE OF IGNITION
CONTROLLABILITY
TURBOPUMP
GAS GENERATOR CONTROL
GAS FLOW
TURBINE INLET PRESSURE
TURBINE RPM
PUMP FLOW
RFNA
UDMH
PUMP DISCHARGE PRESSURE
PUMP INLET PRESSURE
POWER REQUIRED
PERFORMANCE AT ALTITUDE
Isp (ROCKET CHAMBER)
Isp (GAS GENERATOR)
Isp (EFFECTIVE)
WEIGHTS
THRUST CHAMBER
TURBOPUMP
TOTAL ENGINE
SPACE ENVELOPE
LENGTH
DIAMETER
DESIGN SPECIFICATIONS
R1 9A
(4:1 THRUST TURNDOWN)
1250 - 5000 LB
300 - 1200 PSIA
REGENERATIVE (RFNA)
RFNA - UDMH
2.6:1
VARIABLE AREA
3.0 SQ IN.
25:1
75 IN.
HYPERGOLIC PROPELLANTS
THROTTLEABLE, 1250 - 5000 LB
'VARIABLE AREA INJECTOR
WARIABLE AREA NOZZLE
0.06 - 0.87 LB/SEC
500 PSI
12,500 - 25,000 RPM
3.25 - 13.0 LB/SEC
1.25 - 5.Q LB/SEC
400 - 1450 PSI
20 - 50 PSI
10 - 145 HP
278 SEC
150 SEC
265 - 274 SEC
100 LB
100 LB
200 LB
48 IN.
18 IN.
R1 9B
(2.5:1 THRUST TURNDOWN)
1000 - 2500 LB
210 - 525 PSIA
REGENERATIVE (RFNA)
RFNA - UDMH
2.6:1
VARIABLE AREA
2.85 SQ IN.
25:1
75 IN.
HYPERGOLIC PROPELLANTS
THROTTLEABLE, 1000 - 2500 LB
_rVARIABLE AREA INJECTOR
\VARIABLE AREA NOZZLE
0.04 - 0.22 LB/SEC
500 PSI
12,500 - 18,700 RPM
2.6 - 6.5 LB/SEC
1.0 - 2.5 LB/SEC
300 - 675 PSI
20 - 40 PSI
6-34 HP
278, SEC
150 SEC
271 - 275 SEC
60 LB
60 LB
120 LB
48 IN.
18 IN.
Figure 2. Desiga Specifications, R19A and R19B
9-L000178001700n160L00868dCl-V10
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III,. INFRARED DETECTION OF EXHAUST FLAMES
EXhaust flames containing incandescent carbon radiate with nearly black
booty emissivity. The hot exhaust gases containing associated radicals such as
OH radiate considerably upon reassociation of these gases into their parent
molecules. Some polar molecules such as CO2 and H20 radiate discreetly in the
IR region. (The strong 002 and H20 bands are at 4.25 microns and 3.0 microns
respectively.) The polar molecules have lower emissivities than do carbon par-
ticles and, since they do not radiate in a continuum, their total intensity is
less than that of a black body radiating at the same temperature.
The present state of the art shows the development of infrared detection
devices that will seek out exhaust jets from a distance of several miles in the
rarified atmosphere above the earth. At the present time, the most sensitive
of these instruments can detect total intensities in the order of 10-11 to 10-12
watts/cm2. However, infrared detection instrumentation research toward improving
this range is now under wgy in several places. These research centers are working
on outer atmosphere infrared-sensitive devices (selenium sulfide cells, and so
forth) that will detect discreet radiation from 002 and H20 at distances in
excess of 100 miles.
The detection of discreet radiation through atmosphere containing water
vapor and carbon dioxide is quite difficult, since these molecules in the air
absorb this radiation and allow very little to pass through to the detector.
However, the effect of pressure broadening of the discreet bands has helped in the
problem of infrared detection in the atmosphere. An example might be given to
show the radiation intensity of a flame at a temperature of 2000? F, one foot
in diameter, radiating in a vacuum at a distance of 100 miles. The intensity
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would be approximately 10-12 watts. This approaches the maximum limits of
present day sensitivity and would be considerably less in transmission through
the atmosphere.
The infrared spectra may be generally considered to arise from the vibra-
tional and rotational energies of a molecule. Vibrations in a molecule arise
from a motion of the atoms along a line common to both in harmonic motion, or, in
some cases, anharmonic oscillation. The vibrational energy, including its inter-
action with electronic energy, is found in the spectral region from 1 to 20 P
Rotational spectra arise from the rotation of the whole molecule about the center
of gravity of the molecular structure. The energy levels of the rotational
aspects of infrared spectra are weak and are found in regions beyond 20 P and may
extend and be detected as far as 600 P, according to most recent work. For the
purposes of infrared detection of exhaust flames it is only the vibrational
energy levels that are important. For molecules possessing dipole moments
absorption of energy mgy occur if the incident energy corresponds to the frequendy
of mechanical vibration. On the other hand, emission of radiant energy may occur
if there is a transition in the quantum level from a higher to a lower plane.
Absorption of energy will occur when there is excitation from a lower to a higher
quantum level.
The following equations define the interrelationship of wave length (X),
wave number MI frequency (10, and speed of light (c):
1
- =
The problem of employing linear units of frequency or linear units of
wave langth has not been settled as yet and the above relationship provides a
means of interconversion between units. As may be seen, the frequency divided
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by the speed of light will give the reciprocal centimeter or wave number. The
range may cover 10,000 wave numbers (1 micron) to 10 wave numbers (1000 microns).
As it was previously stated, absorption of energy raises the molecule to
an excited state and when the molecule returns to the ground state it re-emits
the absorbed energy. The absorption or radiation of energy is characteristic
of the atomic moieties undergoing vibration and will occur at definite wave
lengths. However, there may be broadening of the spectrum of a characteristic
vibration for gases by the so-called pressure broadening or collision broadening
effect. In these cases there are statistical frequency perturbations caused
by the presence of extraneous atoms, through the medium of van der Waal forces
and phase perturbations produced by collisions. Schematically the problem is
described by assuming a time during which there is a constant frequency of
radiation with no perturbation and then assuming a time constant for perturba-
tions creating phase changes. The effect of pressure broadening is of consider-
able importance in infrared detection of the exhaust of missiles. For CO2 the
4.25 P band is shifted to 4.1 P. to 4.5 P. The effect of broadening apparently
is sufficient for the radiation to get through atmospheric water vapor, which
possesses various absorption bands at 2.5 4, 3.0 P5 and 5.0 to 7.0 P.
There is also a Doppler displacement of wave length that is character-
istic of a source of radiation that moves with speeds of significance in com-
parison to the
speed of light. This is a quadratic Doppler effect and the change
in wave length with velocity of source may be described as follows:
u2
= 10( .! cos e+-z--
2:2)
where: 9= angle of observation,
V = velocity of source, c = speed of light..
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For ordinary speeds, however, the Doppler shift is linear with velocity. The
Doppler shift mgy be significant in detection of CO2 in exhaust gases of a rocket
engine.
Radiant heat may be considered as a stream of energy from a radiating body
traveling with the same velocity as the speed of light, or 3 x 1010 cm-sec-1
(186,000 mi-sec-1). The radiant energy is stopped only when it comes in contact
with a solid, liquid, or gaseous body, being either absorbed or reflected. The
radiant energy from a point source or a spherical body spreads out equally in all
directions, following the simple geometrical law that states that the areas of
similar solids increase as the square of their linear dimensions. Thus the
intensity of radiation or the amount of energy that passes each second through
a plane of unit area perpendicular to the plane of flow, varies inversely as the
square of the distance from the radiating source.
The ability to radiate heat is called emissive power and is a function of
the temperature of a radiating body and its nature, especially that of the sur-
face. It is mathematically equivalent to the total energy emitted per second
per square centimeter of radiation body. The ratio between the emissive power
of a given surface and that of a perfect radiator or black body is called the
emissivity of that surface. The emissive power of a black body is 100 percent,
or 1, so that the emissivity or the ratio of emissive power of any radiating
surface to that of a black body radiation must necessarily be less than 1.
El = emissive power of body
e = --
E = emissive power of black body
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Stefan formulated a law which states that the total rate of radiation
emitted by a unit area of a black body is proportional to the fourth power of its
absolute temperature:
Q = 04144
a = constant equal to 5.77 x 10-5 erg-sec-1 cm2-degree
Q = ergs.sec-1.cm-2 radiated
T = absolute temperature.
For bodies that are not perfect radiators the emissivity enters into the
relationship that depends upon the nature of the radiating body and its surface,
but for gases the emissivity is not even constant. Thus, the formula now
becomes:
= eaT4
where 'e = emissivity.
For a hot gas that is in thermal equilibrium the emitted radiation at any
given wave length is less than the emissive power of a black body at the same
temperature. However, for exhaust flames of a rocket engine equilibrium condi-
tions are not in evidence. Discontinuous radiation of exhaust flames of a
rocket are expected to occur and the characteristic radiation spectrum of 002
will be evident. Although the emissivities of gases at laboratory conditions
are optically thin and deviate considerably from black body radiation, rocket
exhaust gases may approach the black body emission conditions.
The emissivity of a hot gas is a function of the wave length of the
emitted radiation and therefore will have a series of values according to the
wave length of emission. Another term is sometimes introduced, namely, the
engineering emissivity, which is the total of the partial emissivities at the
various wave lengths.
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It must be recalled that it was stated that the emissivity of a gas is
not constant. It depends upon the optical density of a gas in addition to the
temperature. The optical density may be defined as the product of the partial
pressure of the gas in atmospheres or centimeters of mercury and the geometric
length expressed in linear measure such as cm or feet. Some data concerning
emissivities of CO2 are shown below as a function of optical density at a tempera-
ture of 600? K and a wave number 2349 cm-1.
pl pl Emissivity
cm-atmos ft-atmos Total
0.1 0.0033 0.019
0.5 0.033 o.o4o
1.0 0.033 o.o4o
5.0 0.164 0.057
15 0.492 0.062
50 1.64 0.063
100 3.28 0.065
200 6.56 0.067
p = partial pressure
1 = depth
Broadly speaking it may be said that the emissivity of a gas is decreased
as the temperature or pressure is decreased. However, there is a limiting
emissivity that is characteristic of a gaseous radiation at any temperature and
optical density. For example, the limiting emissivity of 002 at 300? K is 0.4
and will increase very slowlo, even with high values of pl (optical density).
The shape of the curve may be seen, on the following page, for emissivity
values of H2 at 12,000? K and P of 20 atmospheres.
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1.00 -
O. 90 _
0.80 _
0.70
0.60 -
:5
cn
.d 0.4?-
0 0.30 -
0.20
0.10
10 20 30 40 5o 60 70 80 90 loo 110 120 130 140 150 160 170 180 190
pl (cm)
The limiting emissivity of CO as a function of temperature may be seen in the
following graph.
200 600 1000 1400 1800 2200 2600 3000 3200
T ?K
eo = emissivity of the first overtone
ef = emissivity of the first fundamental.
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Emissivity data are also given by Lester Lees for air in equilibrium as a
function of density and temperature. Rowever, the data are based on samples of
optically thin slabs.
?1
8.Ma
Temperature
?K
Density (Relative
to sea level air)
lo-4 to 10-3
14,000
0.10 to 1.0
10-4 to 10-2'5
5,000
0.04 to 1.0
10-4 to 10-2
6,000
0.008 to 1.0
10-4 to l0-1?5
7,000
0.001 to 1.0
10-4 to 10-1'2
8,000
0.001 to 1.0
10-4 to 10-1
12,000
0.001 to 0.11
In addition, the emissivity of atomic hydrogen is shown below as a function
of temperature and pressure at 1 = 50 cm.
P atmos
8400? K
9200? K
11,300? K12,000?
K
?..?.
?
10
0.014
0.037
0.26
0.48
20
0.030
0.071
0.35
0.66
40
0.063
0.114
0.57
0.83
70
0.11
0.24
0.74
0.92
loo
0.15
0.32
0.82
0.96
150
0.22
0.44
0.89
200
0.30
0.54
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To solve the problem of the total radiant energy emitted by the exhaust
plume of a rocket engine, we may employ the formula (considering radiation to
take place in a vacuum):
Q=eerTil"
e = emissivity
a = Stefan Boltzmann constant = 5.672 x 10-5 erg-cm-2.degil-sec-1
T = absolute temperature
Q = ergs.sec-1.cm-2
We shall make an approximation of the emissivity of 0021 which is higher
than that of GO, following the limiting emissivity of CO2 given by Penner at
300? K as 0.4. We shall next make an approximation of the size of the radiating
sphere of hot exhaust gas as two feet in diameter. The hot gas sphere is con-
sidered as a stationary point source. Therefore, the formula is as follows
considering the radiation from a sphere of one foot radius:
e = low2 ecrri4
However, the radiation from the point source is now distributed at 100 miles
r'
away on the surface of the earth at an area of 4nR2. Therefore, we have
4rr2eCT4
-
4rR2
where R = distance to the missile flame source from the surface of the earth
r = the radius of the emitting source.
All calculations are in the cgs system and conditions are stated as follows:
Flame plume: 2 feet in diameter or (2 x 30.48 cm)
Flame temperature: 2000? F or 820? K
Distance of flame: 100 miles or (1609.35 x 100) meters.
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Thus:
Report X-394
4n 24
_ reaT
4nR2
= r2eaT4
Q
R2
_ (30.48 x 1)2(0.50)(5.67 x 10-5)(820)4 ,
4.59 x 10-5 ergs.sec-1.cm-2
(100 x 1609.35 x 100)2
4.59 x 10-5
- 4.59 x 10-12 watts.cm-2
107
(Note: One watt is equivalent to 107 ergs.sec -1.)
This energy is detectable by present day devices.
The calculations have assumed a radiation in a vacuum and have ignored
the significant absorption properties of the layers of water vapor in the
atmosphere. In the infrared and red portions of the spectrum quantitatively
the absorptions due to B20 and 02 are most evident. There is also the absorp-
tion due to CO2. For example, the vibration spectrum of B20 completely obscures
the solar spectrum above 24 p. and almost completely above 16 P.
As it was previously stated, the pressure broadening effect on CO2 gas
apparently enables the CO2 radiation spectrum to get through the atmosphere.
At the present time there is disagreement in the field of high altitude detec-
tion. Reliability of detection of infrared radiation at 100 miles altitude is
theoretically possible, as seen by the calculated intensity. However, the
feasibility of detection would be greatly enhanced on an air-based platform at
a height of about 40,000 feet. From the ground the reliability of detection is
poor and depends upon weather conditions, including the degree of water vapor
saturation.
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From purely theoretical considerations then, it is apparent that the CO2
radiation spectra may get through to the surface of the earth, especially when
the broadening effect is considered. The order of transmitted power from our
theoretical radiation source was found to be 10-12 watts.cm-2. This then must
be correlated with the available detectors presently in use. To date tiny
crystals that change in conductivity when excited by infrared radiation are
known. The signal variables in a fixed bridge circuit are fed to servo movements
that guide the missile to its target.
At present, the photoconductor materials available may be tabulated as
follows:
Crystal
Wave Length (I1)
Temperature
0 K
Minimum Detection Power
watts.cm-2
Lead sulfide
2 to 5
193
10-11 to 10-12
Lead sulfide
2 to 4.8
Uncooled
10-10 to 10-11
Lead telluride
2 to 7
90
10-10
Lead selenide
2 to 8.5
90
10-9 to 10-1?
Indium antimonide
2 to 9. 5
90
10-9 to 10-9'5
Germanium
(specially treated)
2 to 11
Cooled
10-9
Thermal detectors
2
10-8
It may be observed that the crystal detectors do not have the infrared
total wave length response characteristic of the thermal detectors; however,
their sensitivity in narrow wave lengths is somewhat greater. It is also seen
that greater sensitivity of the photoconducting crystals is increased by cooling.
For cooling purposes, air-borne cryostats of low weight and size are available
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employing pressurized nitrogen, which cools upon expansion and is continually
recirculated to compression from expansion.
Optical techniques are also available for collecting and focusing the
radiation and are similar to those employed in the visible portion of the spec-
trum. However, the optical materials cannot be opaque to infrared radiation.
Characteristic materials are tabulated below.
TRANSMISSION PROPERTIES OF MATERIALS TO BE USED AS PRISMS OR FILTERS
r
Material
Approximate
Transmission
In Percent
Wave Length
(P)
Thickness
Arsenious sulfide (A523.3)
80
0.5 to 10
2 mm
'
Silver Chloride (AgC1)
80
2.0 to 20
10 mm
L-
rl
Potassium Bromide (KBr)
Sodium Chloride (NaC1)
90
90
0.5 to 27
0.4 to 15
10 mm
10 mm
Linde Sapphire (A1203)
90
0.3 to 5.5
2.5 mm
Lithium fluoride (LiF)
90
0.3 to 5
10 mm
F
Calcium Fluoride (CaF2)
90
0.3 to 9
10 mm
Thallium Bromide Iodide (KRS-5)
70
0.6 to 38
2 mm
Fused Silica (3102)
90
0.2 to 2.0
10 mm
Optical Silicon (Si)
55
1.5 to 6.5
0.1 inch
I
Germanium (Ge)
50
2.0 to 16
4 mm
F
Some problems of detection arise in background radiation, especially that
of the sun. Appropriate use of filters can block out undesirable radiation,
allowing the CO2 spectra to pass. The materials used for filters are similar
to those employed for optical radiation concentration. The feasibility of
employment of filters is immediately evident from the consideration of trans-
mitting and absorption wave lengths of the materials tabulated.
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Report X-39/4
The fact that the missile plume radiation may be considered as a point
source against background radiation, which is diffuse and widespread, lends
itself to design characteristic that uses this phenomenon to advantage. Essen-
tially, alternate transmitting and opaque bars may be used in a scanning device.
It is apparent that the point source will be modulated, whereas the diffuseness
of the background radiation over considerable steradians will not be modulated.
The problem of detection of the theoretically possible radiation available
from an altitude of 100 miles can next be considered on the basis of improvement
of amplification devices and processing of the signal subsequent to crystal
photoconduction. There are many infrared frequency power measurement devices
based on theoretical equivalence of work and heat.
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Report X-394
IV. EXPERIMENTAL ROCKET EgGINE STATIC FIRINGS
It was the intent in this phase of the program to demonstrate certain
capabilities of liquid bipropellant rocket engines pertaining to the subject
feasibility
(1)
(2)
(3)
(4)
questions. The demonstrations were
Long burning
duration,
accumulation
to include:
of 50 minutes
firing
time
Start-stop-restart engine operation
Performance of RFNA-N2H4 and RFNA-UDMH
Comparative emissivity of exhaust flames of RFNA -N2134 (carbon-free)
and RFNA-UDMH (carbonaceous).
The rocket engine planned for use in these demonstrations was the Hughes
Tool Company--Aircraft Division Model 315 water-cooled propellant evaluation
engine, Figure 4. This engine design had originally been developed for evalua-
tion of high energy monofuels in an Air Force program [Contract AF33(60D)-3671,
but since its ignition system uses a 4ypergolic hydrazine slug the hardware was
readily adaptable to bipropellant operation. Figures 5 and 6 are schematics of
the engine thrust stand and feed system used in the static firing test setup.
The emissivity measurements were obtained with a recording IR spec-
trometer (Beckman Model IR-2S). This instrument measures the distortion effect
produced by the radiant energy source of interest on the energy reception profile
of a standard light source. In the static firing test setup the standard light
source was mounted near the radiometer detector and its beam directed by means
of reflecting mirrors to traverse the exhaust flame laterally before entering the
detector. Figure 7 is a photograph of this test setup.
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- - -
\AA AA.
ZS
If
z
0
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kkl
ti ?
11
II
II
II
II
ii
I I
ii
II
II
I
II
II
II
II
kk, at?4,
o
Is<
vIkk
kl. kk, o
NI Y\
4 )
Nt a I,) fk A'kk!kIk`ki,
1 I ! :
q a-X EX- I I I :
? 28
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Emissivity Test Setup
29
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The static test firings, so far as they went, were quite successful in
demonstrating smooth reproducible engine operation using RFNA-N2H . The longest
runs were of two minutes duration, sufficient to obtain a complete scan of the
IR spectrum. A total of approximately five minutes operating time was accumu-
lated on the engine using this propellant combination. The results of these runs
are summarized in Figure 8.
Unfortunately, however, program limitations on time and funds necessitated
a termination of the static firing tests prior to operation with the RFNA-UDMH
combination, so that no comparison of the relative emissivities of the two exhausts
is reportable at this time.
Item
. Units
Run Number
315-5
315-6
315-7
315-8
315-9
Injector Manifold Pressure, Pi
PSIA
343
363
328
329
331
Combustion Chamber Pressure, Pc
PSIA
285
286
260
273
272
Mass Flow Rate Oxidizer, co
Lb/Sec
1.35
1.36
1.32
1.29
1.29
Mass Flow Rate, Fuel, (*et
Lb/Sec
1.13
1.16
1.12
1.12
1.11
Mass Flow Rate, Bulk Propellant, ot
Lb/Sec
2.48
2.52
2.44
2.41
2.40
Mixture Ratio, 0/F
--
1.19
1.17
1.18
1.15
1.16
Mass Flow Rate, Coolant mc
Lb/Sec
5.63
5.33
5.57
6.23
6.25
Coolant Temperature Rise, AT
0 F
22.5
24.5
25.2
30.0
32.4
Total Heat Transfer, k
BTU/Sec
127
130
141
187
202
Duration, t
Sec
3.4
4.3
28.0
120.0
120.0
Thrust, F
Lb
507
507
493
522
515
Thrust Coefficient, CF
,
--
1.24
1.24
1.32
1.33
1.32
Specific Impulse, Isp
Sec
204
201
204
216
214
Characteristic Velocity, O*
Ft/Sec
5340
5270
5040
5260
5260
Figure 8. Summary of RFNA-N2H4 Static Test Firings in Model 315 Rocket Engine
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Figure 9 shows qualitatively the effect of the interjection of the
RFNA-N H exhaust flame into the view of the spectrometer. This figure is a
superimposition of two scannings of the instrument, the dotted line representing
the reception of the standard beam through the atmosphere, and the heavy line the
reception of the standard beam through the atmosphere and the exhaust flame.
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Percent Transmission
10
90
BO
70
60
50
40
30
20
I0
1
I
I
1
I
\
'I%
1
\
6
Reception
Through
H4 Exhaust
Background
Flame?
/
It
/1
/
1
1
%
s
I
RFNA-N2
---- ? --Atmospheric
/
/
/
/
/
/
1
?
\
1
%
?
\
r
I
I
I
I
1
?...".?
%
%
%
%
%
I
/
/
/
/
/
%
1
1
%
i
Operating Conditions
Beckman Model
on
IR 25 Spectrometer
10 x gain
3 mm slit
speed
/
/
/
I
1
1
1
1
1
?
6 in./min chart
2 minute operation
/
/
/
/
/
/
? .
/
I
/
/
/s\J
/
........--,--%.
N.\
./
/
/
/
/
/
?
\
\
?
S.'
?./
......
......-"
,
....-*-

2
L12
-
L311
1.84
2.56
340
418
a..O 5An 5.88 6.1
A (Wave Length in Microns)
Figure 9. Effect of R1NA-N2H4 Exhaust Flame on Standard Source Adsorption Profile
[
t(6C-x q..zodau
r
I
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BIBLIOGRAPHY
1. Lester Lees, Space Technology, Lecture 6A (February 1958), UCLA
2. S. S. Penner, Approximate Emissivity Calculation for Polyatomic Molecules
I. 002, Jour. Applied Phys. 25, 660, May 1954
3. H. Aroeste and Wm. C. Benton, Emissivity of Hydrogen Atoms at High Tempera-
tures, Jour. Applied Phys. 27, 117, February 1956
4. Instrument News, Perkin Elmer Corporation, 9, 1 (Fall 1957).
5. E. A. Brande and F. C. Mhchod, Determination of Organic Structures by
Physical Methods, Academic Press, Inc., New York, N. Y. (1955)
6. W. Brode, Chemical Spectroscopy, John Wiley and Sons, New York, N. Y. (1949)
7. G. P. Harrison, R. C. Lord, J. R. Loofbourow, Practical Spectroscopy,
Prentice-Hall, Inc., Englewood Cliffs, N. J. (1955)
8. E. H. Kennard, Kinetic Theory of Gases, McGraw-Hill Book Co., New York,
N. Y. (1938)
9. G. Herberg, Molecular Spectra and Molecular Structure, I Spectra of
Diatomic Molecules, D. Van Nostrand Co., Inc., Princeton, N. J. (1957)
10. G. Herzberg, II Ingrared and Raman Spectra of Polyatomic Molecules (1956)
11. L. B. Loeb, Basic Processes of Gaseous Electronics, UCLA Press (1955)
12. P. Debye, The Dipole Moment and Chemical Structure, Blackie and Son, Ltd.,
London and Glasgow (1931)
13. W. Finkleburg, Conditions for Black Body Radiation of Gases, J. Optical
Soc. of Am., February 1949, p. 185-186
14. S. S. Penner, Emissivity Calculations for Diatomic Gases, Jour. of Applied
Mechanics, March 1951, p. 53-58
15. S. Wozowski, On the 2Til-4'2ng Bands of CO2, Part II, Phys. Reviews,
September 1942, p. 270-279
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16. H. J. Ramey, Project Squid, Heat Transfer Coefficients for Gasses - -Effect
of Temperature Level and Radiation, Purdue University, 1953, T.R.Pur-24-P
17. U. S. National Bureau of Standards, Energy Transfer in Hot Gases, NBS
Circular 523 (November 1957)
18. D. V. Cogate and D. S. Kothari, Flow of Energy in Thermal Transpiration for
a Bose-Einstein and a Fermi-Dirac Gas, Phys. Reviews, 349-358 (March 1942)
19. R. Gordon, Emissivity of Transparent Materials, J. Am. Cer. Soc., 39,
August 1956, p. 278-287
20. Infrared Emission from High Frequency Discharge in CO2, J. Ap. Physics, 28 ,
June 1957, 10. 737-741
21. R. E. Peck, Fagan, Plerlein, Heat Transfer Through Gases at Low Pressures,
Transactions of the ASNE, April 1951, p. 281-287
22. H. Latzko, Heat Transfer in a Turbulent Liquid or Gas Stream, TM 1068,
October 1944
23. Jackson, Statler, Jinkoff, Heat Transfer to Bodies Traveling at High ?Speed
In Upper Atmosphere
24. Robert NhcFee, Aerojet Corporation, personal communication (1958)
25. Robert S. Neiswander, HAC, personal communication (1958)
26. Robert S. Neiswander, H&C, TM 481 (Secret)
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