A rocket propellant feed system utilizing a bleed turbopump to supercharge a topping turbopump is presented. The bleed turbopump is of a low pressure type to meet the cavitation requirements imposed by the propellant storage tanks. The topping turbopump is of a high pressure type and develops 60 to 70 percent of the pressure rise in the propellant

Ginsburg, A.

Hartmann, M. J.

Stewart, W. L.

English

NASA Technical Reports Server

United States Patent [191 
Stewart et al. 
[111 4,171,615 
[45] Oct. 23, 1979 
[54] SUPERCHARGED TOPPING ROCKET [561 References Cited 
PROPELLANT FEED SYSTEM U.S. PATENT DOCUMENTS 
[75] Inventors: 
[73] Assignee: 
[21] Appl. No.: 
[22] Filed: 
Warner L. Stewart, Parma Heights; 
Ambrose Ginsburg, South Euclid; 
Melvin J. Hartmann, Westlake, all of 
Ohio 
The United States of America as 
represented by the Administrator of 
the National Aeronautics and Space 
Administration, Washington, D.C. 
545,793 
Apr. 21,1966 
2,625,007 1/1953 Truax ..................................... 60/260 
3,028,729 4/1962 Ledwith ................................. 60/259 
3,168,807 2/1965 Ledwith et al. ....................... 60/203 
2,816,417 12/1957 Bloomberg .......................... 60/39.32 
Primary Examiner-Stephen C. Bentley 
Attorney, Agent, or Firm-Gene E. Shook; John R. 
Manning; Norman T. Musial 
1571 ABSTRAm 
A rocket propellant feed system utilizing a bleed turbo- 
pump to supercharge a topping turbopump. The bleed 
turbopump is of a low pressure type to meet the cavita- 
tion requirements imposed by the propellant storage 
tanks. The topping turbopump is of a high pressure type 
I511 Int. c l . 2  ................................................ FO2K 9/02 
[58] Field of Search .......... 60/257, 259, 240, 39.46 R, 
60/203 
and develops-60 6 70 percent ofthe pressure rise in the 
6 Claims, 2 Drawing Figures 
1521 US. a. ......................................... 60/203; 60/259 propellant. 
A - 26 
https://ntrs.nasa.gov/search.jsp?R=19800005931 2020-03-21T02:01:26+00:00Z
Sheet 1 of 2 4 
i4 - 26 
FIG.  I 
US. Patent OC~. 23, 1979 Sheet 2 of 2 4,471, 
84- 
90- ii -98 
t 
FIG. 2 
SUPERCHARGED TOPPING ROCKET 
PROPELLANT FEED SYSTEM 
The invention described herein was made by employ- 
ees of the United States Government and may be manu- 
factured and used by or for the Government for govern- 
mental purposes without the payment of any royalties 
thereon or therefor. 
This invention relates to a rocket propellant feed 
system for advanced chemical or nuclear rocket en- 
gines. More particularly, the invention is concerned 
with a propellant feed system for large high energy 
chemical rockets where the chamber pressure is high to 
reduce the thrust chamber size and to increase the noz- 
zle expansion ratio. The invention is also concerned 
with a propellant feed system for a nuclear rocket en- 
gine in which the reactor size is as small as possible for 
improved nuclear control. 
There are several types of conventional propellant 
feed systems which include straight bleed or gas genera- 
tor systems and straight topping systems. In the bleed or 
gas generator system a fraction of the propellant is 
directed and fed into a gas generator after passing 
through propellant feed pumps. In a nuclear rocket this 
propellant is directed to the core of a reactor. The 
heated gas thus generated then passes through turbines 
used to drive propellant pumps, and the turbine flow is 
exhausted overboard. 
In the straight topping system the fuel flow, after 
being pressurized, is heated by either combustion with 
oxidizer as in a chemical system or through cooling of 
other engine components as in either a chemical system 
or a nuclear system. The fuel passes through the tur- 
bines which power the pumps where the required en- 
ergy is extracted and an appropriate pressure drop oc- 
curs. From the turbine exhaust the fuel passes into ei- 
ther the thrust chamber or the reactor. 
As engine sizes and pressure levels are increased the 
aforementioned systems become quite marginal or even 
unacceptable. As the required chamber pressure is 
raised in a bleed or generator system the amount of flow 
directed through the turbines increases proportionally. 
Because this flow bypasses the main chamber a consid- 
erable penalty in engine performance will occur. Also, 
in order to keep the turbine inlet temperatures at a level 
consistent with stress limitations, a rather large fraction 
of the fuel must be directed from the main flow system. 
As chemical engine pressures and thrust levels are in- 
creased the oxygedfuel ratio of the engine is increased 
to match the corresponding shift in that required for 
peak impulse and to maintain the required fuel tanks at 
a reasonable size. These conditions, coupled with the 
general rise in bleed rate, results in an even larger per- 
centage of fuel flow being directed to the turbines 
which results in a rather large increase in tankage vol- 
ume and weight. 
In the topping system the energy available to drive 
the required turbines is derived from the energy within 
the gas at the turbine inlet and the utilization of this 
energy through an appropriate pressure drop. As the 
required chamber pressure levels are increased, the 
energy required to power the pumps increases propor- 
tionately. In the chemical system, even at turbine limit- 
ing temperatures, this results in a considerable further 
increase in required pump pressure rise which aggra- 
vates the already serious situation. In nuclear applica- 
tions, a similar situation exists wherein the turbine inlet 
4,17 1,6 1 5 
5 
10 
15 
20 
25 
30 
35 
40 
45 
50 
55 
temperature is limited to very low values as a result of 
a desire to obtain required energy through engine com- 
ponent cooling only. Also, with topping complicated 
and difficult mechanical, packaging, and control prob- 
lems arise in the chemical system where two turbines 
must be driven either in series or in parallel by the top- 
ping gas flow. Such problems also include complicated 
ducting of a gas at its highest pressure level. 
In either the bleed system or the topping system, as 
the required pressure level is increased a corresponding 
increase occurs in already critical hydrodynamic and 
mechanical problems that occur within the turbopump 
components. As an example of a hydrodynamic prob- 
lem, increased pressure levels require considerably 
more pump staging or stage loading. Also, a high pump 
pressure rise is not compatible with low net positive 
suction pressures at the pump inlet, and this greatly 
aggravates the cavitation problem. Another hydrody- 
namic problem is that of pump stability caused by the 
increased pump loading. Among the mechanical diffi- 
culties involved is the increase in bearing problems 
caused by both power and rpm increasing. Mechanical 
problems of thrust balancing are also encountered, and 
a difficult problem of critical speed arises. 
These problems have been solved by the propellant 
feed system of the present invention. This system, in 
contrast with the straight bleed system or the topping 
system, has two fuel pumps operating in series. The 
pumping of fuel is the most critical. The second, or high 
pressure pump, is driven by a topping turbine and devel- 
ops the major portion of the total fuel pressure rise 
because it can be designed without compromising the 
requirements of net positive suction conditions at the 
pump inlet. This second pump is supercharged by the 
first pump which is driven by a bleed or gas generator 
driven turbine. This first pump delivers a much lower 
percentage of the required pressure rise and is designed 
to accept the net positive suction pressure condition 
imposed upon it by the fuel storage tank. 
It is, therefore, an object of the present invention to 
provide a propellant feed system in which the fuel 
pumping is divided into two parts to use a combined 
topping and bleed system wherein each pump can be 
optimized for its specific requirement. 
Another object of the invention is to provide a rocket 
propellant feed system having a minimum bleed rate to 
reduce the engine performance penalty and associated 
tank volume. 
These and other objects of the invention will be ap- 
parent from the specification which follows and from 
the drawing wherein like numerals are used throughout 
to identify like parts. 
In the drawings: 
FIG. 1 is a schematic view illustrating a propellant 
feed system constructed in accordance with the present 
invention to be used with a chemical rocket; and 
FIG. 2 is a schematic view illustrating a propellant 
feed system constructed in accordance with the inven- 
tion to be used with a nuclear Dowered rocket. 
60 Referring now to the drawkgs, there is shown in 
FIG. 1 a chemical rocket 10 having a combustion cham- 
ber 12 at one end and a convergent-divergent nozzle 14 
at the opposite end. Liquid propellants are stored in 
tanks 16 and 18. By way of illustration, the tank 16 
65 contains a fuel, such as liquid hydrogen, while an oxi- 
dizer, such as liquid oxygen, is stored in the tank 18. 
These propellants are supplied to the rocket 10 from 
the tanks 16 and 18 through a propellant feed system 
'9 
4,17 1,6 1 5 
3 
illustrated in FIG. 1. Liquid oxygen flows to an oxidizer 
pump 20 from the tank 18 through an inlet pipe 22. 
Pressurized liquid oxygen from the pump 20 is supplied 
to an injector 24 in the rocket combustion chamber 12 
through a pressurized outlet pipe 26. 
A portion of the pressurized liquid oxygen in the pipe 
26 is directed through a bleed line 28 to a gas generator 
30 which operates in a conventional manner. Gas from 
the generator 30 is carried by a line 32 to a turbine 34 
which drives the liquid oxygen pressure pump 20. After 
passing through the turbine 34 the gas is dumped over- 
board through an exhaust 36. 
Liquid hydrogen fuel is supplied to a hydrogen pump 
40 through a supply line 42 as shown in the upper left 
hand portion of FIG. 1. The pressurized liquid hydro- 
gen is transported to the nozzle end of the rocket 10 for 
regenerative cooling after which it is introduced into 
the combustion chamber 12. The regenerative cooling is 
accomplished in a conventional manner, and this cool- 
ing forms no part of the present invention. 
According to the present invention, a second or high 
pressure pump 44 is utilized on conjunction with the 
first pump 40 to pressurize the liquid hydrogen and 
transport it from the tank 16 to the combustion chamber 
12. The first stage pump 40 acts as a supercharger for 
the second stage pump 44, and the pressurized liquid 
hydrogen from the first stage pump 40 is transmitted to 
the second stage pump 44 through a line 44. Liquid 
hydrogen from the second stage pump 46 is supplied at 
a high pressure to regenerative cooling tubes 48 around 
the rocket 10 through a high pressure feed line 50. The 
liquid hydrogen becomes heated in the regenerative 
cooling tubes 48, and a portion is discharged into the 
combustion chamber 12 through a hydrogen injector 
52. 
A portion of the pressurized liquid hydrogen in the 
line 46 between the first stage pump 40 and the second 
stage pump 44 is directed through a bleed line 54 to the 
gas generator 30. The gas produced by the generator 30 
is directed through a gas line 56 to a turbine 58 which 
drives the first stage hydrogen pump 40. After passing 
through the turbine 58 the gas is dumped overboard at 
a discharge 60. 
Another important feature of the invention is a provi- 
sion of a topping turbine 62 which drives the second 
stage liquid hydrogen pump 44. A portion of the liquid 
hydrogen in the cooling tubes 48 surrounding the 
rocket 10 is directed to the topping turbine 62 through 
ducts 64. After passing through the topping turbine 62 
the gaseous hydrogen leaves through a discharge line 66 
and is directed into the combustion chamber 12. 
By utilizing the supercharged topping feed system of 
the present invention instead of the straight bleed or 
topping systems, the critical fuel pumping operation is 
divided into two pumps operating in series. The second, 
or high pressure, pump 44 develops 60 to 70 percent of 
the total pressure rise because it can be designed with- 
out the compramising requirements of net positive suc- 
tion conditions at the pump inlet. In this manner, the 
pump 44 is supercharged by the first pump 40 which is 
driven by the bleed or gas generator drive turbine 58. 
The pump 40 delivers between 30 to 40 percent of the 
required pressure rise and is designed to accept the net 
positive suction conditions imposed on it by the tank 16. 
The utilization of the turbopumps 40 and 44 in a su- 
percharging manner to pressurize the hydrogen from 
the tank 16 reduces the required bleed rate to the lines 
28 and 54 from that of a conventional bleed system. 
4 
This, in turn, greatly reduces the attendant engine per- 
formance penalty and reduces the associated tank vol- 
ume and weight problems. 
Another feature of the invention is that the power 
5 required of the topping turbine 62 is reduced considera- 
bly from that required for a straight topping system. 
Therefore, the required pressure drop across the turbine 
62 is drastically reduced which, in turn, lowers the 
overall pump pressure rise. 
The use of a gas generator 30 to power the drive 
turbine 58 for the supercharger pump 40 and the turbine 
34 for the oxidizer pump 20 enables only one turbine 62 
to be driven in a topping fashion thereby eliminating 
many of the mechanical, packaging, and control prob- 
l5 lems associated with additional turbines in the topping 
portion of the cycle. It is contemplated that the turbines 
34 and 58 may be powered in series by the gas generator 
30. 
As the pressure required by the rocket 10 and the 
2o combustion chamber 12 increases, the two fuel pumps 
40 and 44 can accommodate additional staging much 
more efficiently than the single pumps utilized by the 
prior art. Each pump 40 and 44 can be optimized for its 
specific requirements. The topping driven pump 44 can 
25 be optimized for pressure production without consider- 
ation of cavitation. The supercharger pump 40, which is 
now a rather low pressure producing pump, can be 
optimized to meet the cavitation requirements imposed 
3o upon it by the tank 16, either through appropriate de- 
sign of the pump itself, or under more severe conditions, 
by the incorporation of a geared, hydraulic, or gas 
driven boost pump stage. Inasmuch as each pump 40 
and 44 can be loaded to a smaller extent, the stability 
35 problems are reduced. Also, the three major mechanical 
problems of bearing, thrust, and critical speed are 
greatly reduced because each turbine pump unit is much 
more conservative. 
Referring now to FIG. 2, there is shown a nuclear 
powered rocket 70 having a nuclear reactor 72 for heat- 
ing the propellant and a convergent-divergent nozzle 
74. A propellant, such as liquid hydrogen, stored in a 
tank 76 is supplied to the rocket 70 through a propellant 
feed system embodying the features of the present in- 
45 vention. Liquid hydrogen in the tank 76 at a pressure of 
about 30 psia is fed to a geared inducer 78 through a 
propellant feed line 80. The liquid hydrogen leaves the 
inducer and enters a supercharger pump 82 which raises 
its pressure to about 630 psia. 
This pressurized liquid hydrogen is fed to a main 
uumu 84 at the uressure of 630 usia through the high 
10 
50 
- -  
pressure line 86. -The main pump84 raises the pressure 
of the liquid hydrogen to 1680 psia and feeds this pres- 
surized propellant to regenerative cooling tubes 88 
The hydrogen in the cooling tubes 88 is heated to a 
temperature of 300" R and is fed into a topping turbine 
92 at pressure of 1430 psia. About 20 percent of the 
hydrogen fuel is bypassed around the turbine 92. The 
60 hydrogen leaves the topping turbine 92 at a discharge 
94 and is fed to the core of the nuclear reactor 72. At the 
intake side of the core 72 the hydrogen has a pressure of 
lo00 psia and at the discharge side of the core adjacent 
the nozzle 74 the hydrogen is at a pressure of 800 psia. 
Hydrogen at a temperature of 1500" R is fed to a 
bleed turbine 96 through a bleed line 98. The bleed rate 
is about 3 percent. The bleed turbine 96 drives the su- 
percharger pump 92. After passing through the bleed 
55 around the rocket 70 through a line 90. 
65 
4,17 1,6 15 
5 b 
pumps and the second pump delivers a major portion of 
the pressure rise, 
said first pump being adapted to accommodate a net 
positive suction pressure condition imposed upon it 
3. A feed system as claimed in claim 1 wherein the 
rocket engine is of the chemical type having a combus- 
tion chamber at one end and a convergent-divergent 
nozzle at the opposite end for receiving fuel and oxidant 
'the first and second pumps being connected to the 
another pump connected to the oxidant storage tank, 
a gas generator connected to the bleed turbine for 
means for directing a portion of the fuel and oxidant 
from the outlets of the first pump and other pump 
to said gas generator, 
means for directing pressurized fuel from the outlet of 
5 by the storage tank. 
10 from a pair of propellant storage tanks, 
fuel storage tank, 
15 driving the first pump, 
turbine 96 the gaseous hydrogen is discharged over- 
board through a discharge 100. 
All of the aforementioned pressures and temperatures 
are given by way of illustration and not by way of limi- 
tation. By using the feed system of the present invention 
with the nuclear rocket 70 the reduced energy required 
permits the topping turbine 92 to operate with reason- 
able pressure drops under the temperature restrictions 
providing the system with some margin. 
While several embodiments of the invention have 
been shown and described, it will be appreciated that 
various structural modifications may be made without 
departing from the spirit of the invention or the scope of 
the subjoined claims. 
What is claimed is: 
1. A feed system for supplying a liquid propellant 
from a storage tank to a rocket engine comprising 
a first Pump having an thereof connected to the 20 second pump to regenerative cooling tubes about 
said rocket engine, and 
cooling tubes through the topping turbine. 
storage tank and an outlet directed toward the 
means for directing a portion of the fuel from the rocket engine, 
a second pump having an inlet thereof connected to 
the outlet of said first Pump SO that said first Pump 4. A feed system as claimed in claim 3 including an- 
supercharges said second pump, said second pump 25 other bleed turbine for driving the other pump for 
having an outlet in communication with the rocket pumping the oxidant. 
engine, 5. A feed system as claimed in claim 1 wherein the 
rocket engine is of the nuclear type having a reactor 
core at one end and a convergent-divergent nozzle at 
means for directing pressurized propellant from the 
outlet of the second pump to regenerative cooling 
tubes about said convergent-divergent nozzle, and 
a bleed turbine for driving said first pump, 
means for removing a portion of the propellant from 
said feed system to drive said bleed turbine, 
a topping turbine for driving said second pump, 
means for directing a portion of the propellant from 
30 the opposite end, 
said feed system through said topping turbine for means for directing said propellant from said cooling 
driving the same, and 35 tubes to the topping turbine. 
means for directing said last named portion Of the 
Propellant to said rocket engine after Passing 
through said topping turbine. 
2. A feed system as claimed in claim 1 wherein the 
pressure of the liquid propellant is raised by both the 40 
6. A feed system as claimed in claim 5 including 
means for directing a portion of the gaseous propel- 
lant in the convergent-divergent nozzle to the 
bleed turbine. * * * * *  
45 
50 
55 
65 


Supercharged topping rocket propellant feed system

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