Electrothermal rocket engine using resistance heated heat exchange

Jack, J. R.

Moeckel, W. E.

English

NASA Technical Reports Server

Jan. 16, 1962 J. R. JACK ETAL 
ELECTRO-THERMAL ROCKET 
Filed Sept. 23, 1960 
- I -  - 
3,016,693 
INVENTORS 
J.R. JACK 
https://ntrs.nasa.gov/search.jsp?R=19700024045 2020-03-17T01:33:34+00:00Z
3,016,693 United States Patent Ofice Patented Jan. 16, 1962 
1 
3,016,693 
maRO-THERMAL ROCKET 
John R. Jack, North O w e d ,  and Wolfgang E. Moeckel, 
Bay Village, Ohio, assignow to the United States of 
America aa represented by the Administrator of the 
National Aeronautics and Space Administration 
Filed Sept. 23,1960, Ser. No. 58,147 
6 Claims. (Cl. 6-55) 
(Granted under Title 35, US. Code (1952), sec. 266) 
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 pertains to an electro thermal rocket 
employing an electrically-powered resistance heated heat 
exchanger. 
Analysis of space missions has indicated that a large' 
increase in payload may be achieved by using an elec- 
trical propulsion system instead of the conventional chem- 
ical system. For missions in the gravitational field of 
the earth and moon, the optimum specific impulse for an 
electric propulsion device ranges from approximately 
1500 to 5000 seconds. Adequate payloads may be han- 
dled for certain satellite missions with specific impulses 
ranging from 800 to 1500 seconds. An electric propul- 
sion device that operates well over the lower half of this 
specific impulse range-that is 800 to 2000 second-is 
the thermal arc jet or plasma jet. It is particularly de- 
sired that an electric rocket be capable of operating for 
60 days or longer with the required thrust and impulse. 
The arc jet, as well as the instant rocket, heats the propel- 
lant to a very high temperature and then expands it in a 
convergent-divergent nozzle. Propulsion is thus obtained 
by a thermal dynamic process. However, the arc jet has 
several severe penalties associated with it when operating 
times of the order of 60 days are considered. 
A disadvantage associated with the use of the electn'c 
arc rocket is that there is extreme electrode erosion when 
the rocket is used for any period of time approaching 60 
days or even much less. There is additionally the disad- 
vantage in that there is an extreme problem in the arc 
voltage regulation and difficulties are involved in the arc 
starting. 
A further disadvantage is that there is heat transfer to 
the engine. In fact it appears to be beyond existing tech- 
nology to develop an arc jet rocket without expecting 
major breakdowns in its operation for the required time. 
In the present invention the propellant is admitted to 
the rocket chamber and passes through an electrically- 
powered resistance heated heat exchanger to bring the 
propellant up to the temperature required for ejection 
through a conventional convergent-divergent nozzle at 
very high velocity. To develop the required thrust the 
hydrogen must be heated to approximately 5500" R. 
This temperature can be handled by present-day refrac- 
tory materials such as carbon, tantalum, and tungsten. 
The herein disclosed space engine is particularly designed 
for approximately one pound of thrust and a specific im- 
pulse in the region of 1000 seconds. Such a propulsion 
device would be useful for altitude control or as the final 
stage engine for a communication satellite mission. 
Thus, an object of this invention is to provide a novel 
electro-thermal rocket. 
Another object of this invention is to provide a pro- 
pulsion device utilizing an electrically-powered resistance 
heated heat exchanger. 
An additional object of this invention is to provide an 
electm-thermal rocket engine that is capable of operating 
for extended periods without major breakdown. 
A further object of this invention is to provide a de- 
vice which produces a propulsive jet through the heating 
2 
of propellant during passage through or around electrical- 
ly-heated essentially solid surfaces. 
Other objects and many attendant advantages of the 
present invention will be apparent from the following de- 
6 ,  tailed description when taken together with the accom- 
panying drawing in which: 
The single drawing is a partially-sectioned pictorial view 
of an electro-thermal rocket engine using an electrically- 
powered resistance heated porous tungsten heat exchanger. 
Referring now to the drawing, there is shown the rocket 
engine 11 having a cylindrical outer casing or shell 20 
of nickel construction or any other material having ap- 
proximately the same emissivity as nickel. The outer 
casing 20 is provided with metal bellows 22 to permit a 
16 heat exchanger 14 to expand axially without buckling. 
Two sleeves 23 and 24 are disposed at opposite ends and 
within the casing 20. The two sleeves 23 and 24 are 
principally of boron nitride construction or any high 
temperature electrical insulating material. Mounted con- 
20 centrically within the sleeves 23 and 24 are two electrical 
connectors 18 and 19. The connector 18 mounted with- 
in the sleeve 24 at the front end of the engine is a positive 
electrical connection while the connector 19 mounted 
within the sleeve 23 at the rear end of the engine is a 
26 negative electrical connection. The electrical connectors 
18 and 19 are of molybdenum construction or any suitable 
refractory metal. The electric power can be furnished 
by a nuclear turboelectric system as indicated in the fig- 
ure by the box dehed as a power source. The connec- 
30 tor 18 receives its energy through the electrical connec- 
tions 31 attached to the power source. The power source 
is additionally connected by electrical connection 32 to 
ground 33, while the connector 19 is connected to the 
ground 33 by the electrical connection 34. Situated con- 
35 centricaIly within and supported by the two electrical con- 
nectors 18 and 19 is the annular porous tungsten heat ex- 
changer 14. The heat exchanger 14 can be of "Poroloy" 
made by Bendix Filter Division of Bendk Aviation Cor- 
poration, which is made by winding h e ,  flattened wire 
40 into tubular shapes. At the front end of the heat ex- 
changer 14 are three radiation shields 29 to prevent heat 
leakage. The shields are preferably of tungsten construc- 
tion and are supported by a ring member 17 situated with- 
in the electrical connector 18. Coolant coils 16 are pro- 
vided to direct coolant to the face of the radiation shield 
46 29. At the rear end of the heat exchanger 14 is situated 
the convergent divergent nozzle 21 which, in this em- 
bodiment, is of graphite for test purposes mainly. In 
an actual space flight application, the nozzle would prob- 
ably be constructed of a refractory material that does not 
50 react with hydrogen. High temperature O-ring seals 25 
are used throughout the rocket engine to prevent any leak- 
age of the propellant between the various parts of the 
engine. Coolant coils 15 and 26 surround the outer cas- 
65 ing 20 and serve to cool the engine. Water can be used 
as the coolant in this experimental prototype. Of course, 
10 
60 
65 
70 
in actual space application a regenerative type of coolant 
system would be used whereby the propellant acts as the 
coolant and cools the nozzle, as well as the outer casing 
before entering the heat exchanger. 
In operation of the engine the propellant, which is hy- 
drogen, is admitted through the inlets 12 and passes and 
flows in a direction as indicated by the arrow heads be- 
tween the radiation shields 13 which are made of tungsten 
and then on through the walls of the electrically-powered 
rdistance heated porous tungsten heat exchanger 14 
whereby the hydrogen is raised to a temperature of ap- 
proximately 5500" R. and is then exhausted through the 
nozzle 21. The radiation shields 13 serve to prevent heat 
leakage from the heat exchanger and are separated from 
one another by support means 27 and 28, as shown in the 
figure. By having the electrical connections in contact 
, 3,016,085 
4 
3. An electro thermal rocket engine comprising an out- with the walls of the heat exchanger 14, it becomes &re- 
eistance to the electrical current and the material is raised er cylindrical shell, an annular porous tungsten heat ex- 
to an exceedingly high temperature so that when the hy- changer mounted concentrically within said outer shell, 
drogen passes through the porous walls it is raised to the cylindrical radiation shields situated between said outer 
desired approximately 5500" R. and, of course, expands 6 shell and said heat exchanger, propellant inlet meam at- 
upon being heated to such a temperature, forcing its way tached to said outer shell for admitting hydrogen propel- 
out through the .convergent-divergent nozzle 21. &- lant to said heat exchanger whereby the hydrogen propel- 
though the particular embodiment shown utilizes a porous lant flows around said radiation shields and then through 
heat exchanger, it should not be concluded that this is said porous heat exchanger, and a convergent-divergent 
the only type possible. Any engine employing an elec- 10 exhaust nozzle rigidly attached to said outer shell. 
Uically-powered resistance heated heat exchanger is suit- 4. The porous tungsten heat exchanger of claim 3 which 
able. Thus, many heat exchanger designs utilizing this is an annulus of wound flattened tungsten wire. 
feature are feasible for the present invention. As an ex- 5. An electro thermal rocket engine comprising: an 
ample, a staggered tube heat exchanger using refractory outer cylindrical shell; an annular porous metal heat ex- 
wire filaments is an .alternative embodiment of this in- 16; changer mounted within said outer shell; electrical con- 
vention. Thus, this invention embraces all electro-ther- nection means operably coupled to opposite ends of said 
mal space engines which utilize the principle of heating heat exchanger whereby an electrical current may be 
the propellant during passage through or around elec- caused to pass through said heat exchanger; cylindrical 
trically-heated solid surfaces. radiation shields positioned within said outer shell be- 
Obviously, many modifications and variations of the 20 tween said heat exchanger and said shell; propellant inlet 
present invention are possible in the light of the above means attached to said outer shell for admitting propel- 
teachings. It is, therefore, to be understood that within lant to said heat exchanger whereby the propellant ad- 
the scope of the appended claims the invention may be mitted flows around said radiation shields and through 
practiced other than as specifically described. said porous heat exchanger; means surrounding said outer 
What is claimed: 26 shell to provide passage for a coolant to cool said shell; a 
1. An electro thermal rocket engine comprising an Outer convergentdivergent exhaust nozzle rigidly attached to 
shell, an annular porous metal heat exchanger situated said outer shell. 
within said shell, electrical connections operably associ- 6. The engine set forth in claim 5 wherein said outer 
ated with the opposite ends of said heat exchanger where- shell is provided with bellow means to permit axial ex- 
by an electrical current may be caused to pass through 30 pansion thereof. 
said heat exchanger, propellant inlet means operably as- 
sociated with said outer shell for admitting propellant to References Cited in the file of this patent 
said heat exchanger whereby the propellant admitted UNITED STATES PATENTS 
passes through and is heated by said porous heat ex- 
changer, and a convergent-divergent exhaust nozzle rigid- 811 
ly attached to said outer shell. 
2. The rocket engine of claim 1 wherein said annular 
porous metal heat exchanger is an annulus of porous 
tungsten. 20,697 Great Britain _ _ _ _ _ _ _ _ _ _ _  Sept. 17,1907 
3 
1,687,550 
2,837,623 
Ehret ______________--- Oct. 16, 1928 
Judson _____________---_ June 3, 1958 
FOREIGN PATENTS 


Electro-thermal rocket Patent

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