A supersonic combustion rocket is provided in which a small rocket motor is substituted for heavy turbo pumps in a conventional rocket engine. The substitution results in a substantial reduction in rocket engine weight. The flame emanating from the small rocket motor can act to ignite non-hypergolic fuels

Franciscus, L. C.

Weber, R. J.

English

NASA Technical Reports Server

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON. D.C. 20546
G? . ' . DEC 2 ? 1973
TO: KSI/Scientif ic & Technical Information Division
„ Attentions Hiss Winnie M. Morgan
FROM: GP/Office of Assistant General Counsel for
Patent Matters
SUBJECT: Announcement of NASA-Owned U.S. Patents in STAR
In accordance with the procedures agreed upon by Code GP
and Code KSI, the attached NASA-owned U«S« patent is being
forwarded for abstracting and announcement in NASA STAR.
The following information is provided
U.S. patent No* ; _j. H I I i
Government or
Corporate Employee
Supplementary Corporate
Source (if applicable)
NASA Patent Case No. ; iCj/J - • || . Q 5 S* - /
NOTE - If this patent covers an invention made -by a corporate
employee of a NASA Contractor, the following is applicable:
Yes / / No / X/
Pursuant to Section 305(a) of tho National Aeronautics and
Space Act, the name of the Administrator of NASA appears on
the first page of the patent? however, the name of the actual
inventor (author) appears at the heading of column No. 1 of
the Specification, following the words "« »^ .^'«*yj^ > srsspect to
an invention of «, „ <,'
Elisabeth A« Carter
Enclosure *
Copy of patent cited above
https://ntrs.nasa.gov/search.jsp?R=19740005389 2020-03-17T14:10:53+00:00Z
Un&d States Patent
Weber et al.
[in 3,777,490
[45] Dec. 11, 1973
[54] SUPERSONIC-COMBUSTION ROCKET
[75] Inventors: Richard J. Weber, Cleveland; Leo
C. Franciscus, Lakewood, both of
Ohio
[73] Assignee: The United States of America as
represented by the Secretary of the
National Aeronautics and Space
Administration, Washington, D.C.
[22] Filed: Mar. 10, 1972
[21] Appl. No.:'233,519
[52] U.S. Cl 60/258, 60/259
[51] Int. CI F02k 9/02
[58] Field of Search 60/258, 259, 39.48,
60/39.49,39.12
[56] References Cited
UNITED STATES PATENTS
3,200,583 8/1965 Sobey et al 60/259
3,570,249 3/1971 Baum et al 60/259
2,936,577 5/1960 Amneus 60/258
2,408,111 9/1946 Truax et al ....60/259
3,178,885 4/1965 Loughran 60/39.48 X
2,558,483 6/1951 Goddard 60/39.49 UX
Primary Examiner— Carlton R. Croyle
, Assistant Examiner—Robert E. Garrett
Attorney—N. T. Musial
[57] ABSTRACT
A supersonic combustion rocket is provided in which
a small rocket motor is substituted for heavy turbo
pumps in a conventional rocket engine. The substitu-
tion results in a substantial reduction in rocket engine
weight. The flame emanating from the small rocket
motor can act to ignite non-hypergolic fuels.
5 Claims, 3 Drawing Figures
TNASA-Cas^LEB-11058-1)
SUPERSONIC-CGMBUSTION BOCKET
(NASA) 5 P
patent
CSCL 21H
Dnclas
00/28 24243
PATENTED DEC 11 1973 3,777,490
FIG. 2
FIG. 3
3,777,490
1 2
SUPERSONIC-COMBUSTION ROCKET The supersonic combustion rocket of this invention
noii-MM r>c TUC IHTWCMTIOM comprises a nozzle means, a mixing chamber means,ORIGIN OF THE INVENTION ,iquid propellant storage means means operatively
The invention described herein was made by employ- connecting storage and chamber means and a means
ees of the United States Government and may be man- 5 for introducing a high velocity gas into said chamber
ufactured and used by or for the Government for gov- means in a manner that a pressure differential is cre-
ernmental purposes without the payment of any royal- ated between the storage means and the combustion
ties thereon or therefor. chamber. This pressure differential causes the liquid
_, . _ propellant to flow from the storage tank means through
BACKGROUND OF THE INVENTION ,„ The Pconnecting means and into the chamber, with the
1 . Field of the Invention liquid propellant becoming entrained in the gas stream.
The present invention relates to a liquid fuel rocket Preferably the means for introducing the high velocity
motor and specifically to a supersonic rocket motor gas is a small supersonic rocket motor having its nozzle
having novel means for introducing the liquid propel- directed toward the chamber.
lant. . 15 In one embodiment the storage means includes a fuel
2. Description of the Prior Art and an oxidant storage tank. The connecting means in-
Conventional rocket motors utilize heavy turbo eludes first and second conduits connected to the oxi-
pumps to transport liquid propellant components from dant and fuel storage tanks respectively, and to a third
their respective storage tanks to the combustion cham- conduit. This third conduit is connected to the cham-
ber. These pumps are heavy and complicated and pres- 20 ber means and contains the small rocket motor. Prefer-
ent complex design problems. Additionally, extra pro- ably the first and second conduits are connected to the
pellant is required -because of the weight of these third conduit upstream of the nozzle of the small rocket
pumps. The elimination of these pumps has therefore motor.
been a goal of engine designers. Unfortunately, no sys- In a second embodiment the fuel and oxidant tanks
tern has been developed which produces a decrease in 25 are connected to the chamber by first and second con-
weight without a corresponding decrease in the specific duits, respectively. A small rocket motor is positioned
impulse of the rocket motor. For exmaple, when a pres- in each of the conduits.
sure feed system is used, the resulting rocket has a In a third embodiment the oxidant and fuel storage
lower specific impulse than the conventional rocket tanks are operatively connected to the combustion
using the turbo pump feed system. Furthermore, much 3® chamber through first and second conduits. The small
of the weight loss achieved by eliminating the pumps is . rocket motor is connected to the chamber in a position
counteracted by the increase in weight of the storage wherein the exhaust from the small rocket creates a
tanks necessary for pressurization. In U.S. Pat. No. pressure drop between each of the storage tanks and
3,073,1 19, a feed system is disclosed in which one of the combustion chamber.
the supply pumps is eliminated. This is accomplished 35 nPSPRfPTfON nF THF DRAWING
by feeding one of the fuel components into the throat BRIEF DESCRIPTION OF THE DRAWINGS
of an ejector. The inlet for the other component is con- FIG. 1 is a schematic view of this invention using a
nected to the suction side of the ejector. This second single small rocket motor to inject a mixture of oxidizer
component is drawn into the combustion chamber as a and liquid fuel into the combustion chamber of the
result of the pressure differential between the throat 40 larger rocket motor.
and the second component inlet created by pumping FIG. 2 is a schematic diagram of a second embodi-
the first fuel component into the ejector. ment of this invention in which separate small rocket
SUMMARY OF THE INVENTION *"
The object of the present invention is to reduce the FIG. 3 illustrates a third embodiment of this inven-
weight of the conventional rocket motor. - tion in which a single small rocket motor is used to in-
Another object of the invention is to reduce engine ject separate streams of fuel and oxidizer into the mix-
weight by eliminating the heavy turbo pumps required ing chamber.
in prior art conventional rocket motors without a cor- INSCRIPTION OF THF PRFFFRRFn
responding reduction in specific impulse. 50 FMROniliFNT?
A further object of this invention is to provide a UMBUUlMtN 1 5>
rocket motor in which the propellant is travelling at su- FIG. 1 illustrates an embodiment of this invention
personic speeds in the combustion chamber when it is which can be used for hypergolic or nonhypergolic pro-
ignited.
 55 pellants. The supersonic rocket motor includes the
These and other objects are accomplished by substi- basic rocket motor elements, to wit, an exhaust nozzle
tuting a source of high velocity gas for the conventional 1, a combustion chamber 2, propellant storage tanks 3
turbo pumps utilized in rocket feed systems. This high and 4, and a fuel injection means 5. The injection
speed gas source is preferably one which supplies gas means comprises the heart of the invention and in-
at supersonic speed and in the preferred embodiment eludes a mixing conduit 6 and a source of high velocity
constitutes a small rocket motor. The small rocket gas, such as the small rocket motor 7 illustrated.
motor is located between the fuel storage tank and the In operation the exhaust from the small rocket motor
combustion chamber with its nozzle directed toward creates a pressure differential between the storage
the latter. The high speed exhaust from the small tanks and the exhaust stream in the mixing conduit 7
rocket motor creates a pressure differential between ,5 which causes the liquid fuel and liquid oxidant to flow
the storage tanks and the exhaust stream which draws from their respective storage tanks into the mixing con-
the fuel components from the storage tanks into the duit where they are entrained in the exhaust gases and
combustion chamber. then propelled into the combustion chamber at super-
3,777,490
3 4
sonic speeds. In the system illustrated in FIG. 1, the is also possible to use different fuel and oxidant combi-
small rocket motor exhaust ignites the fuel mixture as nations for the small and large motors. In the former
it enters the combustion chamber. Alternatively, a sep- case the same fuel storage tanks can be used to supply
arate pilot light, not shown, can be used to ignite the both engines. However, it is also possible to use sepa-
mixture if the exhaust gas temperature of the small 5 rate storage tanks for both rockets,
rocket motor is below the ignition temperature. The advantages of the supersonic combustion engine
FIG. 2 illustrates an embodiment of the invention of this invention are clearly illustrated by the analytical
which may be used for hypergolic fuel systems. Once results obtained using a liquid oxygen-liquid hydrogen
again the same basic rocket components are present. propellant mixture in the embodiment of FIG. 3. Using
These include the exhaust nozzle 1, the combustion 10 a mixture ratio of liquid oxygen to liquid hydrogen of
chamber 2, the fuel and oxidant storage tanks 3 and 4 8 and operating at a chamber pressure of 500 p.s.i. an
and the fuel injection system 5. Here the fuel injection exit mach number of 2.16 is obtained in the small
system comprises two mixing conduits 6' and 6" and motor with an area ratio of 2.45. Using a large rocket
two small rocket motors 7' and 7". Downstream of the with a propellant flow about 10 times that of the small
rocket motors two mixing zones 8' and 8" are pro- 15 rocket and having a combustor area ratio of 56 with a
vided. liquid oxygen/liquid hydrogen mixture ratio of 8 the
In operation the fuel contained in tank 3 is drawn following exit conditions are obtained for the super-
into the mixing conduit 6' by the pressure differential sonic combustion rocket: A combustor exit pressure of
created by the exhaust stream from the rocket motor 0.83 p.s.i.; a temperature of 3,500"R.; a velocity of
7'. This component becomes entrained in the exhaust 20 13,800 feet per second; and a mach number of 4.2. The
from the rocket in mixing zone 8'. In a like manner, the vacuum specific impulse of this supersonic combustion
oxidant contained in tank 4 is intermixed with the ex- rocket is 470 seconds if the exhaust gases are expanded
haust from the rocket motor 7" in mixing zone 8". The to 0.15 p.s.i. Thus a small rocket which delivers 15,000
two propellant components entrained in their respec- pounds of thrust results in a supersonic combustion
live small rocket motor exhausts are combined in com- 25 rocket having a total vacuum thrust computed to be
bustion chamber 2 where they spontaneously ignite at 230,000 pounds. This large thrust is achieved with a
supersonic speeds. substantial reduction in rocket weight since the only
The embodiment of FIG. 3 is also useful in hypergolic turbo pumps used are those necessary for the small
fuel systems. A single small rocket motor is used to feed rocket motor which are significantly lighter than the
the oxidant and fuel components into the combustion 30 pumps necessary for a conventional motor having
chamber as in the embodiment of FIG. 1. However, un- 230,000 pounds of thrust.
like the embodiment of FIG. 1, the two propellant com- The throat area between the chamber 2 and the noz-
ponents do not intermix until they enter the combus- zle 1 may be proportionately larger than that shown in
tion chamber as in the embodiment of FIG. 2. In opera- FIGS. 1, 2 and 3 and is large enough so that sonic flow
tion of the rocket motor depicted in FIG. 3, the exhaust 35 from the chamber to the nozzle is avoided. As is known
from the small rocket motor 7 creates a pressure differ- to those skilled in the art, sonic flow in the throat of a
ential between the storage tanks and the combustion rocket motor tends to create excessive heating in that
chamber inlets which draw the propellant components area. Advantageously, since the exhaust flow from the
from tanks 3 and 4 through conduits 6' and 6" respec- rocket motor embodying the invention is already at su-
tively. The two streams emanating from these conduits 40 personic speed, no necked-down throat between the
simultaneously become entrained in the exhaust gases chamber and the nozzle is required to produce super-
and intermix with each other thereby spontaneously ig- sonic velocity of the exhaust as in the prior art.
niting as they enter the combustion chamber 2. What is claimed is:
The embodiments of FIGS. 2 and.3 can also be used 1. In a rocket motor comprising a nozzle means com-
for non-hypergolic fuels. In the FIG. 2 motor it may be municating with a mixing chamber means, liquid pro-
necessary to provide a pilot light to ignite the fuel mix- pellant storage means, and means operatively connect-
ture. However, in the embodiment of FIG. 3 the ex- ing said storage means and said chamber means, a small
haust gases from small rocket 7 may serve this function. supersonic rocket motor means for introducing a high
In designing the supersonic combustion rocket of velocity gas stream into said chamber means, said small
FIG. 2, the various cross sectional flow areas are se- rocket having a nozzle directed toward said chamber
lected to achieve supersonic flow of the fuel compo- means and communicating therewith whereby a pres-
nents entering the combustion chamber. For the rock- sure differential is created between said storage means
ets of FIGS. 1 and 3, the propellants may be injected • and said chamber means, whereby the liquid propellant
at subsonic or supersonic speeds. However, in all cases is caused to flow from said storage tank through said
the burned mixture is to leave the combustor at super- connecting means into said chamber means, throat
sonic speeds. The overall rocket should be designed in means disposed between said chamber and said nozzle
a manner that the various flow areas, wall contours and means, -said throat being sufficiently large to avoid
pressures yield maximum thrust and specific impulse. sonic flow in the throat area, said liquid propellant be-
All of these parameters can be determined by those
 6(J coming entrained in the high velocity gas stream then
skilled in the rocket engine art. burning and leaving said chamber at supersonic veloc-
The liquid propellant for the small rocket motor is ity.
provided through a conventional turbo pump feed sys- 2. The rocket motor of claim 1 wherein said liquid
tem. Alternatively, as a result of the low chamber pres- propellant storage means comprises an oxidant storage
sures required by the small rocket motor, a pressure
 65 tank and a fuel storage tank, said connecting means
feed system will provide satisfactory results. Generally, comprises first and second conduits connected to said
the propellant used for the small rocket motor will be oxidant and fuel storage tanks respectively, a third con-
the same as that used by the larger motor. However, it duit connected at one end to said chamber means, and
3,777,490
means connecting said first and second conduits to said
third conduit, said small rocket motor being positioned
in said third conduit.
3. The rocket motor of claim 2 wherein said means
connecting said first and second conduits to said third
conduit is located upstream of the nozzle of said small
rocket motor!
4. The rocket motor of claim 1 wherein said liquid
propellant storage means comprises an oxidant storage
tank and a fuel storage tank, said connection means
comprises a first and second conduit means connecting
said oxidant and fuel storage tanks to said chamber
means, respectively, said small motor means compris-
10
ing first and second small rocket motors located in said
first and second conduit means, respectively.
5. The rocket motor of claim 1 wherein said liquid
propellant storage means comprises an oxidant storage
tank and a fuel storage tank, said conduit means com-
prises first and second conduits connecting said oxidant
and fuel storage tanks respectively to said chamber,
and means connecting said small rocket means to said
chamber in a position wherein the exhaust from said
small rocket creates a pressure drop between each of
said storage tanks and said chamber.
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Supersonic-combustion rocket

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