Regenerative cooling system for small rocket engine having restart capability and using noncryogenic hypergolic propellant

Latto, W. T., Jr.

English

NASA Technical Reports Server

W. T. LATTO, JR 3,32 1,922 May 30, 1967 
SMALL ROCKET ENGINE 
Filed Got .  29, 1, 
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INVENTOR 
WILLIAM T LATTO JR. 
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ATTORNEYS 
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3,321,922 
William T. Latto, Jr., kakewood, Obh, assignor to the 
United States of America 59s represented by the Ad- 
ministrator o€ &he National Aeronaxtics and Space 
Administration 
ct. 29, 1954, Ses. NQ. 407,595 
3 Claims. (a. 60-260) 
The invention described herein may be manufactured 
and used by or for the Government of the United States 
of America for governmental purposes without the pay- 
ment of any royalties thereon or therefor. 
This invention is concerned with a regeneratively- 
cooled rocket engine that is small in size and has a 
relatively low thrust. More particularly, the invention re- 
lates to cooling the combustion chamber of a small rocket 
engine having the capability of restarting using non- 
cryogenic hypergolic propellants. 
Small rocket engines, such as those having low thrusts 
of 1000 pounds or less, are difficult to cool because the 
ratio of the heated area to the volume of coolant is high 
compared to larger rocket engines having the same com- 
bustion characteristics. Regenerative cooling has been 
used with large rocket engines, but satisfactory cooling by 
this method becomes difficult to achieve as the engine 
size decreases. This cooling problem is further complicated 
by the fact that most small rocket engines in the 10 to  
1000 pound thrust range are required to have restart 
capabilities so that they can perform attitude and tra- 
jectory control functions. 
A typical large regeneratively-cooled rocket engine has 
a combustion chamber and nozzle formed by an assembly 
of tubes that are shaped to the proper configuration. One 
end of the tube assembly is in communication with an 
injector while the other end is connected to a supply 
of cryogenic propellant which acts as a coolant as it flows 
through the tubes, and provisions are made both for 
interrupting the propellant flow to stop the engine as 
well as for again initiating this flow for restarting. When 
the propellant flow through the injector stops, the cooling 
of the engine is interrupted because the propellant ceases 
to flow through the tube assembly. 
While the use of cryogenic propellants for regenerative 
cooling is preferred because of their large heat-sink ca- 
pacity, low-temperature operating problems and long term 
storage difficulties arise when these propellants are used. 
The use of certain non-cryogenic or storable propellants 
for regenerative cooling to  alleviate the difficulties en- 
countered with cryogenic propellants gives rise to the 
problem of dissipating the residual heat in the chamber 
materials after shutdown without causing violent de- 
composition of the propellant trapped in the cooling tubes. 
Flush systems to remove the entrapped propellant and 
intercoolers to lower the propellant temperature are com- 
plicated and expensive. 
These problems and difficulties have been solved in the 
small thrust rocket engine of the present invention which 
has regenerative cooling passages formed in a high 
strength shell which encloses a refractory metal liner 
with an insulating liner between the outer shell and the 
inner liner. Adequate cooling is realized by providing an 
accumulator for receiving propellant from the regenera- 
tive cooling system after the propellant flow has been 
interrupted so that the propellant continues to  flow 
through the regenerative cooling system. The propellant 
received by the accumulator is used when the engine is 
restarted. The improved regeneratively-cooled rocket en- 
gine of the present invention utilizes storable hypergolic 
propellants to achieve the necessary restart capabilities 
without being dependent upon a separate starting system or 
requiring low temperature operation. 
It is, therefore, an object of the present invention to 
provide a small rocket engine which utilizes storable 
propellants and has an improved wall construction. 
Another object of the invention is to provide an im- 
proved regenerative cooling apparatus for a small rocket 
engine having means for continuing the flow of propellant 
through the cooling apparatus after flow to the com- 
bustion chamber has been stopped. 
A further object of the invention is to provide a small 
1o rocket engine having an improved wall construction for 
regenerative cooling using a storable propellant with pro- 
vision for enabling the propellant to flow into an accumu- 
lator for a period of time after propellant flow to the 
engine combustion chamber has been interrupted so that 
15 a sufficient amount of heat from the chamber wall is ab- 
sorbed to prevent decomposition of propellant in contact 
with the chamber wall. 
Other objects of the invention will be apparent from 
the specification which follows and from the drawings in 
20 which like numerals are used throughout to identify like 
parts. 
In  the drawing: 
FIG. 1 is an axial quarter-section view, with parts sche- 
matically illustrated, showing a small regeneratively-cooled 
25 rocket engine constructed in accordance with the inven- 
tion; and 
FIG. 2 is a sectional view taken along the line 2-2  
in FIG. 1. 
Referring now to the drawings, there is shown in FIG.1 
30 a small rocket engine 10 that is constructed in accordance 
with the present invention. The engine PO includes a com- 
bustion chamber 12 and a nozzle 14 which is oriented 
in the thrust direction along the longitudinal axis away 
from the ccmbistion chamber. The engine 10 further in- 
35 cludes an injector 16 at the opposite end of the combustion 
chamber 12 from the nozzle 14 for bringing storable hy- 
pergolic propellants, such as nitrogen tetroxide and hy- 
drazine, from a plurality of sources into contact with- 
in the combustion chamber 12. These sources include 
40 a liquid fuel supply tank 18 and a liquid oxidizer supply 
tank 20 connected to the engine 10 by a fuel line 22 and 
an oxidant line 24 respectively. A suitable valve 26 in the 
line 24 shown schematically in FIG. 1 controls the flow 
of oxidant from the tank 20 to the injector 16 while a 
4; valve 28 controls the flow of fuel. An actuator 29 sche- 
matically shown in FIG. 1 is operably connected to the 
valves 26 and 28 for initiating and interrupting the propel- 
lant flow. The injector 16, valves 26 and 28, and the ac- 
tuator 29 may be of the type described in copending ap- 
50 plication Ser. No. 182,692, filed Mar. 26, 1962 now 
Patent No. 3,170,286. 
According to the present invention the small rocket en- 
gine 10 has an improved wall construction comprising 
an outer shell 30 enclosing an inner liner 32 with an in- 
55 termediate liner 34 between them. The outer shell 30 is 
of a high strength material, such as stainless or high 
chrome steel. Cooling passages 36 are formed in the outer 
shdl 30 for regenerative cooling. These passages are of 
variable height to  provide adequate cooling. 
The outer shell 30 may be fabricated in two parts with 
the oooling passages 36 being formed on the outer periph- 
ery of the inner portion in the form of a screw thread. 
This inner portion then is inserted in the outer portion 
with a suitable insert being provided at the throat of the 
65 nozzle 14. If desired, the passages could be formed on 
the inner surface of the outer portion of the shell 30 in 
the form of screw threads with the outer surface of the 
inner portion being smooth. Here again a suitable insert 
would be provided in the area of the throat of the 
The inner liner 32 is of a refractory metal, such as a 
tantalum-tungsten alloy or molybdenum. By way of ex- 
5 
60 
70 nozzle 14. 
3,321,922 
ample, an alloy comprising 80% tantalum and 20% 42 so that the chamber volume 48 is reduced to a mini- 
tungsten may be used. The thickness of this liner depends mum. 
on the structural requirements of the rocket engine 1% The accumulator 40 has a displacement that is ade- 
although it is uontemplated that a thickness between 0.01 quate to permit enough propellant to flow through the 
and 0.10 inch would be satisfactory for most applications. 5 passages 36 after the engine 10 has been deactivated to 
The intermediate liner 34 is of a refractory insulat- absorb the required quantity of heat from the engine wall. 
ing material, such as zirconia, zinc oxide, or alumina. Eecause the propellant received by the accumulator 40 
The thickness of this liner depends upon the cooling de- is returned to the system upon restarting the engine 10, 
sign requirements with good insulation being of prime im- no propellant is lost. Also no separate pressurization 
portance. system is required for the accumulator 48 nor are addi- 
As shown in FIG. 1 the fuel line 22 is connected to  ticnal valves used. 
the passages 36 at the downstream end of the nozzle 14. While a preferred embodiment of the rocket engine 
A fuel, such as hydrazine, is supplied Under Pressure to has been described, it will be appreciated that various 
the Passages 36 from the tank 18 for cooling the engine structural modifications may be made without departing 
10. The fuel leaves the passages 36 at the Upstream end 15 from the spirit of the inven)tion or the scope of the sub- 
of the rocket engine 10 though a line 38 to the valve 28 joined claims. For example, while the passages 36 are 
that controls the supply of fuel to the nozzle 16. shown in the form of screw threads it is contemplated 
In operation, the valves 26 and 28 are opened by the that parailel passages may be used in larger engines, such 
actuator 29 to start the engine 10. An oxidizer, such as as those havine thrusts between 500 and 1000 uounds. 
nitrogen tetroxide, flows through the line 24 to the injec- 20 
tor 16 and the fuel likewise flows through the line 38 
to the injector where these propellants are injected into 
the combustion chamber 12. The refractory material of 
the liners 32 and 34 of the multilayer chamber wall forms 
a heat sink which is cooled by the fuel as it flows through 25 
the passages 36. When it is desired to  stop the engine 10 
the actuator 29 closes the valves 26 and 28 thereby in- 
terrupting the flow of the propellants to the injector 16. 
An important feature of the invention is the provision 
of an accumulator 48 in the form of a chamber that is 30 
in communication with the line 38 through a by-pass con- 
duit 42. A piston 44 in the accumulator chamber forms a 
wall that is mounted for reciprocable movement between 
an empty position shown in FIG. 1 and a filled position 
wherein the piston is located at the opposite end of the 35 
acculmulator chamber. A spring 46 biases the piston 44 
toward the by-pass conduit 42 to provide a minimum 
chamber volume 48 on the propellant receiving side of 
the piston. It will be appreciated that other types of 
resilient means, such as bellows, may be used to move 40 
the piston 44. It is further contemplated that a com- 
pressible fluid may be utilized to perform this function. 
After the engine 10 has been stopped or deactivated 
by the closing of the valves 26 and 28 propellant fluid 
which is under pressure is transferred through the by-pass 45 
conduit 42 to the interior of the accumulator 40. This 
pressurized propellant moves the piston 44 against the 
biasing force exerted by the spring 46 toward the end 
of the accumulator 40 remote from the by-pass conduit 
42 which increases the chamber volume 48 to a maximum 50 
value. In this manner pressurized fluid propellant con- 
tinues to flow through the passages 36 for a period of 
time after the engine 10 has been deactivated, and this 
flow continues until the piston 44 has moved to the Mled 
position with a maximum chamber volume. 
vides adequate propellant flow through the passages 36 
to absorb a sufficient quantity of heat from the engine 
wall so that the propellant trapped in these passages after 
this flow ceases does not become overheated. 
When the valves 26 and 28 are reopened by the actua- 60 
tor 29 to restart the engine 110 the spring 46 moves the 
piston 44 toward the by-pass conduit 42 because of the 
system pressure loss as the propellants pass through the 
injector 16. This movement of the piston 46 expels the 
propellant from the accumulator 4@ into the combustion 65 
chamber 12 through the injector 16. 
The position of the piston 44 depends upon the differ- 
ence between the pressure of the propellant in passages 
36 and the pressure exerted by the spring 46. It is pref- 
erable that the spring pressure is sufficiently limited SO 70 
that the spring 46 is fully compressed when the pressure 
in chamber volume 48 reaches that in the remainder of 
the system. It is also desirable for the spring 46 to exert 
a force great enough to move the piston 44 into contact 
with the accumulator wall adjacent the by-pass conduit 75 
This pro- 55 
What is clai&ed is: 
1. A rocket engine comprising, 
a combustion chamber and a nozzle in continuation 
thereof, 
said combustion chamber and nozzle having a wall 
including an outer layer of high strength mate- 
rial having cooling passages of variable heights 
therein, an intermediate liner of refractory in- 
sulating material selected from the group con- 
sisting of zirconia, zinc oxide, and alumina, and 
an inner layer of an alloy comprising 80% 
tantalum and 20% tungsten forming a heat sink, 
an injector closing the end of the combustion chamber 
opposite to said nozzle, 
said cooling passage forming regenerative cooling 
means surrounding said combustion chamber and 
nozzle to provide a flow path for regenerative cool- 
ing flow of a propellant fluid to said injector, 
a plurality of sources of supply of pressurized pro- 
pellant fluid, 
first conduit means connecting one of said sources to 
said cooling passages, 
second conduit means connecting another of said 
sources directly to said injzctor, 
valve means in the flaw path from each of said sources 
to interrupt propellant flow to said injector, one of 
said valve means being located between said cooling 
passages and said injector, 
an accumulator chamber for receiving a propellant, 
a bypass conduit having one end in communication 
with the interior of said accumulator chamber and 
the other end in communication with said cooling 
passages and said one valve means, 
a wall mounted for reciprocable movement in said ac- 
cumulator chamber between a first position remote 
from said one end of said bypass conduit thereby 
providing a maximum volume in said accumulator 
chamber to a second position adjacent said one end 
of said bypass conduit thereby providing a minimum 
volume in said accumulator chamber, said bypass 
conduit serving to transfer pressurized propellant 
Buid to said accumulator chalmber thereby exerting 
a pressure on said wall to move the same to said first 
position whereby continued flow of propellant fluid 
through said cooling passages is provided for a period 
of time after said engine has been deactivated to cool 
said heat sink, and 
resilient means within said accumulator chamber for 
biasing said movable wall from said first position to 
said second position, said resilient means serving to 
supply propellant to said injector by moving said 
wall to said second position when said one valve 
means is opened to reactivate said engine. 
2. A rocket engine comprising, 
a combustion chamber and a nozzle in continuation 
thereof, 
3,321,932 
6 
3. A composite wall for a rocket engine combustion 
an outer layer of high strength material having cooling 
an intermediate liner of refractory insulating material 
selected from the group consisting of zirconia, zinc 
oxide, and alumina, and an inner layer of an alloy 
comprising 80% tantalum and 20% tungsten. 
References Cited 
5 
said combustion chamber and nozzle comprising 
an outer layer of high strength material having 
cooling passages of variable heights therein, 
an intermediate liner of refractory insulating passages of variable heights therein, 
material selected from the group consisting of 5 
zirconia, zinc oxide, and alumina, and an inner 
layer of an alloy comprising 80% tantalum and 
20% tungsten forming a heat sink, 
an injector closing the end of the combustion chamber 
said cooling passages forming regenerative cooling UNITED STATES PATENTS 
chamber and nozzle comprising, 
10 opposite said nozzle, 
1,860,891 511932 Goddard. means surrounding said combustion chamber and 
2,476,185 711949 Goddard. nozzle in said outer layer and providing a flow path for regenerative cooling flow of a propellant fluid, 2,548,485 411951 Lubbock _ _ _ _ _ _ _ _ _ _ _  60-39.69 a plurality of sources of supply of propellant fluid, 15 2,555,080 511951 Goddard. 
first conduit means connecting one source of supply 2,814,929 1211957 Morley et al. _ _ _ _ _ _ _ _  6 G 3 5 . 6  to said cooling passages, 3,170,289 211965 Kramer _ _ _ _ _ _ _ _ _ _ _ _  60-35.6 second conduit means connecting another source of 3,220,180 1111965 Heckert _ _ _ _ _ _ - _ _ _ _ _  60-35.6 supply directly to said injector, 
OTHER REFERENCES valve means in the flow path from each of said sources 20 of supply to interrupt flow to said injector head, 
an accumulator chamber, and 
bypass conduit means extending from said accumulator 
chamber to said one valve means and said cooling 
passages, said bypass conduit $means serving to trans- 25 
fer propellant fluid from said one valve means to 
provide a continued flow of propellant fluid to said 
cooling passages for a period of time after said en- 
gine has been deactivated. 
Porter, H. B.: “Rocket Refraotories,” NavOrd Report 
Pages 14, 15, 22, 4893, NOTs 1191, PB 131503, 1955. 
23 relied on. 
MARK NEWMAN, Primary Examiner. 
D. HART, Examiner. 


Small rocket engine Patent

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