The engine has a conventional body and nozzle configuration. The monopropellant fuel is fed into the thruster with dual injection tubes via an injector shell with dual spray jets. The spray jets are positioned generally opposed to each other. A heater screen pack combination thermally decomposes the fuel after injection into the combustion chamber of the thruster

Kuenzly, J. D.

English

NASA Technical Reports Server

United States Patent 1191 [111 4,288,982 
Kuenzly et al. [45] Sep. 15, 1981 
[54] LOW THRUST MONOPROPELLANT 
ENGINE 
[76] Inventors: 
[21] Appl. No.: 
[22] Filed: 
John D. Kuenzly, Redondo Beach, 
Calif.; Robert A. Frosch, 
Administrator of the National 
Aeronautics and Space 
Administration, with respect to an 
invention of John D. Kuenzly 
971,474 
Dec. 20, 1978 
Related U.S. Application Data 
[63] Continuation of Ser. No. 819,029, Jul. 25, 1977, aban- 
doned. 
[51] Int. Cl.3 ................................................ F02K 9/44 
[52] U.S. CI. ............................. 60/200 R 60/39.46 M 
[58] Field of Search ............. 60/39.46 M, 200 R, 203, 
60/258, 218; 239/491; 75/172, 176; 219/301 
~561 References Cited 
U.S. PATENT DOCUMENTS 
981,481 1/1911 Shoenberg .......................... 219/301 
3,054,252 9/1962 Beckett et a1 219/301 
3,236,699 2/1966 Pugh et a1 75/176 
2,378,348 6/1945 Wilrnes et al. ...................... 239/491 
3,312,539 4/1967 Marshall et al. ...................... 75/176 
3,375,109 3/1968 Peters .................................... 75/176 
3,696,616 10/1972 Lucas .................................... 60/218 
3,361,599 1/1968 Barganier .............................. 75/176 
FOREIGN PATENT DOCUMENTS 
629809 9/1949 United Kingdom ................ 239/491 
OTHER PUBLICATIONS 
“Monopropellant Hydrazine Resistojet”, TRW Report 
20266-6024-RO-00; Mar., 1973. 
Primary Examiner-Robert E. Garrett 
Attorney, Agent, or Firm-Ronald F. Sandler; John R. 
Manning; John 0. Tresansky 
[571 ABSTRACT 
A low thrust monopropellant engine having a conven- 
tional body and nozzle configuration wherein the 
monopropellant fuel is fed into the thruster with dual 
injection tubes via an injector shell with dual spray jets 
formed therein. The spray jets are positioned generally 
opposed to each other. A heater-screen pack combina- 
tion thermally decomposes the fuel after injection into 
the combustion chamber of the thruster. 
10 Claims, 3 Drawing Figures 
https://ntrs.nasa.gov/search.jsp?R=19820010440 2020-03-21T16:21:47+00:00Z
U.S. Patent Sep. 15, 1981 4,288,982 
FIG 1 
FIG 3: 
0 
29' 7- 
4.288.982 
1 
LOW THRUST MONOPROPELLANT ENGINE 
ORIGIN OF THE INVENTION 
The invention described herein was made in the per- 
formance of work under a NASA contract and is sub- 
ject to the provisions of Section 305 of the National 
Aeronautics and Space Act of 1958, Public Law 85-568 
(72 Stat. 435; 42 USC 2457). 
This is a continuation, of application Ser. No. 
819,029, filed July 25, 1977 and now abandoned. 
BACKGROUND OF THE INVENTION 
The invention relates generally to low thrust rocket 
engines. More particularly, the invention relates to low 
power thruster for extremely low temperature applica- 
tions. 
Low thrust monopropellant engines have been em- 
ployed for many years to satisfy attitude and velocity 
control requirements for satellites. The conventional 
thruster employes a catalytic bed to ignite the propel- 
lant. The use of a catalytic bed in the thruster requires 
the almost exclusive use of hydrazine as a propellant. 
Hydrazine (N2H4) contains no carbon atoms and conse- 
quently will not degrade the catalyst bed which may be, 
for example, Shell 405. The use of hydrazine, however, 
may cause problems with the propellant supply system 
because of its high freezing point which raises the dan- 
ger of the thruster propellant freezing when the satellite 
is in outer space. An attempt to substitute low freezing 
point propellants with otherwise similar characteristics 
to hydrazine has been made with the catalyst bed thrust- 
ers but such propellants have degraded the catalyst bed 
because they are universally carbonaceous and combus- 
tion has resulted in carbon deposits on the catalyst bed. 
This action substantially lowers the useful life of the 
thruster. 
OBJECTS OF THE INVENTION 
Accordingly, it is an object of the present invention 
to provide a low-level monopropellant thruster employ- 
ing low freezing point propellants. 
Another object of the invention is to provide a low- 
level monopropellant thruster with thermal decomposi- 
tion of the propellants employed. 
A further object of the instant invention is to provide 
a low-level monopropellant thruster without a catalyst 
bed. 
Yet another object of the present invention is to pro- 
vide a low-level monopropellant thruster which em- 
ploys low freezing point propellants while retaining a 
long useful life. 
SUMMARY OF THE INVENTION 
According to the present invention the foregoing 
objects are attained by a low-level thruster employing 
dual injector tubes for carrying a monopropellant 
which tubes are connected to an injector shell shaped to 
form spray jets for atomizing the propellant. 
BRIEF DESCRIPTION OF THE DRAWING 
FIG. 1 is a cross-sectional view taken along the major 
axis of the thruster. 
FIG. 2 is a cross-sectional view of the atomizing 
spray injector configuration of the thruster taken along 
its major axis. 
5 
10 
15 
20 
25 
30 
35 
40 
45 
50 
55 
60 
65 
2 
FIG. 3 is a front view of the injector shell of the 
thruster taken through the center lines of the injector 
tube recesses. 
DETAILED DESCRIPTION OF THE 
PREFERRED EMBODIMENT 
Referring to the drawing wherein reference numerals 
identify like parts throughout the several views, FIG. 1 
depicts a thruster body which may be mounted on a 
thruster valve and fed a propellant by a pressurized 
diaphram or bladder. If, in turn, the thruster is con- 
tained in a spinning satellite, it may be gravity fed. Both 
of these operations may be accomplished in a conven- 
tional fashion. 
A thruster body 11 is formed of a Mo-50% Re alloy 
and defines a combustion chamber 12. The thruster 
body surrounds and is attached to a platinum screen 
pack 13, which, in turn, is attached to and surrounded 
by a heater 14. A nozzle 15 also formed of Mo-50% Re, 
is welded to the rear end of thruster body 11. A barrier 
tube 16, constructed of Haynes alloy 25, is disposed 
intermediate of a seal plate 17, constructed of 304 series 
stainless steel, and the frontal area of thruster body 11. 
The barrier tube and the seal plate function to offset the 
thruster body 11 from, typically, a thruster valve (not 
shown). An injector shell 18 is secured to the front end 
of thruster body 11 by brazing and functions to atomize 
the propellant before it enters the head space of the 
thruster chamber 12. Dual injector tubes 19, 20, com- 
posed of Mo-50% Re or pure rhenium, run along the 
longitudinal axis of seal plate 17 and barrier tube 16 and 
serve as lines for propellant. The ends of these tubes are 
attached in a generally opposed fashion to the injector 
shell 18. 
It is to be understood that screen pack 13 materials, 
packing density and characteristic length play a secon- 
dary-but very important-role with respect to propel- 
lant dispersal. More specifically, short screen packs 
greatly benefit pulse-mode operation but hamper 
steady-state operation. High packing densities allow the 
use of shorter screen packs, but increase residence times 
which, in turn, affect response time. Increased residence 
times promote endothermic, metal-catalyzed reactions. 
As a result, a rather standard packing density of 118 
screendcm is employed. 
It has been found that screen pack materials which 
are catalytically active for carbonaceous, low freezing 
point propellants are rapidly degraded in a decompos- 
ing hydrazine type propellant environment. Typical of 
these materials are stainless steels, Inconels and Haynes- 
25. The preferred material for the screen pack 13 is 
platinum. The best results are obtained when not more 
than one-half of the screen pack area is impinged on by 
the atomized propellant spray. 
The thermo-electric heater 14 provides heat to the 
screen pack 13 which, in turn, provides for the thermal 
decomposition of the propellant as an alternative to 
decomposition provided by a catalyst bed. The dis- 
closed heater has a virtually unlimited life. The heater is 
annular and fabricated completely from refractory met- 
als and oxides. A heater retainer tube 21, composed of 
Mo-50% Re, encloses an electrical insulator 22 which is 
composed of high purity alpha-alumina fabricated from 
alpha-alumina powder sintered at 2200" K. A heater 
wire 23, made of W-3% Re, is embedded in the insulator 
22. Electrical power is applied to the heater wire 23 by 
lead wires 24 which are formed of rhenium and at- 
tached by a Mo-Ru braze. The heater 14 is also affixed 
4,288,982 
3 
to the thruster body 11 by brazing. The thermal capabil- 
ities of the disclosed heater configuration are about 
1000” K. higher than prior art heaters commonly used. 
FIG. 2 shows more clearly the injector tubes 19, 20 
entering the injector shell 18. The front end of shell 18 
has an injection area 25 into which propellant is fed by 
the injector tubes. The front end of the injection area is 
sealed by an injector plug 26 while the rear end of the 
injection area 25 is open to the thruster chamber 12 via 
aperture 27. 
FIG. 3 shows more clearly the two recesses 28 
formed in shell 18 for receiving injector tubes 19, 20. 
The center lines of the recesses may be generally paral- 
lel to a line going through the major axis of the shell 18 
in a perpendicular fashion and offset therefrom. Con- 
nection lines 29 are formed in shell 18 from the recesses 
28 in an essentially tangential fashion into the injection 
area 25. Aperture 27 is centrally located with respect to 
the injection area 25. The tip regions of the recesses 28 
form the spray jets 30. It is important that the spray jets 
are generally opposed to each other. 
It should be noted that the off-set spray jet configura- 
tion of this invention may be employed with catalyst 
bed thrusters as well as with thermal decomposition 
thrusters. Single stream injection has been found to 
cause severe bed channeling and ultimate flooding with 
a low thermal mass screen pack. Longer screen pack 
assemblies have allowed sustained operation but at the 
cost of holding power and pulsed-mode performance. 
An injector therefore is required which would (1) pro- 
vide necessary pressure drop to limit the instantaneous 
flow rate immediately after valve opening, (2) decouple 
the propellant supply system from chamber pressure 
oscillations, (3) provide a small dribble volume, and (4) 
prevent two phase flow conditions which may result in 
propellant decomposition in the injector. In effect, sin- 
gle tube injectors could not produce a thruster using 
carbonaceous, low freezing point propellants capable of tubes* 
meeting auxiliary propulsion requirements regardless of The principlesy preferred embodiment$ and modes Of 
changes in head space and Screen pack configuration. operation of the present invention have been described 
Satisfactory operation could be obtained for hereinbefore. The invention which is intended to be 
steady-state operation, not pulse-mode operation. Vari- protected herein, however, is not to be construed as 
ous prior art of fuel injection such as multiple limited to the particular forms disclosed since these are 
injectors, drilled hemisphere ‘‘shower heads”, concen- to be regarded as illustrative rather than restrictive. It is 
tric cylinders, swirler tubes and tangential injection 45 to be understood that variations may be made by those 
misters, are unable to satisfy the required performance skilled in the art without departing from the spirit of the 
characteristics. present invention. 
While specific materials are identified for the dis- What is claimed is: 
closed embodiment, other materials may be useful. The 1. A thruster for orienting spacecraft, comprising: 
embodiment disclosed avoids the use of the previously 50 a thruster body; 
used Haynes-25 alloy except for the barrier tube 16 a screen pack surrounded by and connected to said 
where the material is not critical. The criteria for the 
successful application of a given material for thruster means for injecting a propellant into said thruster 
use is (1) must not nitride, (2) must allow easy fabrica- 
tion, (3) must exhibit low catalytic activity with the 55 a plurality of injector tubes discrete from the thruster 
propellant employed, both prior to and after combus- 
tion, and, (4) must be easily joined. Three materials an injector shall having an infundibuliform cavity 
successfully meet these criteria; namely, platinu- with a discharge aperture, receptive means for 
m-lO% iriduim (Pt-lO%Ir); zirconia grain stabilized connecting said injector tubes to said shell and a 
platinum (ZGS-Pt); and, molybdenum-50% rhenium 60 plurality of opposed limitation means disposed to 
(Mo-50% Re). atomize and tangentially discharge a propellant 
The disclosed embodiment successfully allowed the directly into the cavity. 
use of low freezing point, carbonaceous propellants 2. The thruster of claim 1 wherein said injector tubes 
without requiring a catalyst bed which would be sub- and said injector shell are formed of molybdenum rhe- 
ject to degradation. Exemplary propellants which may 65 nium. 
be used in the thruster according to this invention in- 3. The thruster of claim 1 wherein said receptive 
clude (1) monomethyl-hydrazine (MMH), (2) 50% means include dual recesses formed in said injector 
MMH, 35% hydrazine (N2H4), 15% ammonia (NH3), 
thruster body; and 
body, including: 
body; and 
shell. 
4 
(3) 50% MMH, 50% N2H4, (4) 80% N2H3, 20% NH4, 
(5) Aerozine-50 (50% N2H4-50% unsynmetrical di- 
methylhydrazine (UDMH)) and, (6) 23% hydrazine 
azide-77% N2H4 (designated as HA). All of these fuels 
5 exhibit freezing points well below the freezing point of 
hydrazine of 274.5” K. and, therefore, eliminate the 
inherent restrictions placed on spacecraft when using 
hydrazine and avoid complications related to spacecraft 
thermal control systems. 
Low power level thrusters typically exhibit 0.1 to 0.2 
pounds of thrust. The disclosed embodiment may well 
be useful on thrusters exhibiting thrust in the order of 
magnitude of 10 pounds. Physically, a typical low level 
thruster may be about g’’ in diameter and two inches 
l5 long. Injector tubes may have a five to six mil I.D. and 
a 14 mil O.D. For a thruster in the 0.1 to 0.2 pound 
range an injection area diameter of 0.0240” and an aper- 
ture diameter of 0.0075” have proved satisfactory. 
Spray jet regions may be offset from the center of the 
The disclosed injection technique and composite 
heater-screen pack combination allow production of a 
low level thruster which may use low freezing point, 
carbonaceous propellants and yet have a potentially 
25 unlimited life. Such thrusters are viable alternatives to 
an entire generation of catalytic hydrazine thrusters. 
Several cautions must, however, be observed. In partic- 
ular, joining and fabrication must be performed with the 
3o complete absence of any material containing silicon, 
phosphorous or titanium. Even trace quantities of these 
elements migrate to noble metal grain boundaries and 
cause severe embrittlement. Also, the introduction of 
air with C02 in seemingly insignificant quantities during 
35 propellant handling attacks the propellant system lines 
and causes the generation of high metallic ion concen- 
trations which may clog the small diameter injector 
10 
2o shell by about 0.0090”. 
4. The thruster of claim 3 wherein said dual recesses 
are formed generally parallel and opposed to each 
other. 
5. The thruster of claim 1 wherein said thruster body 
6. A thruster for maneuvering satellites in space, 
a thruster body; 
a screen pack formed of platinum and surrounded by 
a heater surrounding and connected to said thruster 
a heater retainer tube surrounding said thruster body 
a electrical insulator surrounding said thruster body 15 
a heater wire encapsulated within said electrical insu- 
means for injecting a propellant into said thruster 
dual injector tubes discrete from the thruster body, 
for conducting said propellant parallel to an axis 
and through single arcs to diametrically opposed 
terminals offset from the axis; and 
an injector shell having a pair of substantially parallel 25 
and opposed recesses therein for individually re- 
ceiving different ones of said opposed terminals, an 
infundibuliform right circular cavity formed within 
said shell coaxial to the axis and in communication 
with said recesses via a pair of generally opposed 30 
spray jets for atomizing said propellant, said spray 
jets formed by tip regions of the recesses disposed 
to discharge tangentially into the circular cavity, 
and an aperture centrally located at an apex of the 
cavity between said spray jets, the aperture cou- 35 
pling said cavity with the interior of said thruster 
body. 
is formed of molybdenum-50 percent rhenium. 5 
comprising: 
and connected to said thruster body; 10 
body, including: 
forming a space therebetween, 
within said space, and 
lator surrounding said thruster body; and 
body, said injection means including: 20 
7. A low power thruster, comprising: 
a combustion chamber; 
structural means adjoining the chamber, and provid- 40 
ing an interior cavity having an interior perimeter 
4,288,982 
45 
50  
55 
60 
65 
6 
and a centrally located aperture opening into the 
chamber; 
a plurality of recesses supplying said cavity, oriented 
tangentially to said interior perimeter; 
an equal plurality of atomizing means formed in gen- 
eral opposition by said recesses and adjoining the 
cavity; 
an equal plurality of conduit means distinct of said 
structural means, communicating directly into said 
atomizing means; and 
catalytic means for causing decomposition of a pro- 
pellant inside said chamber. 
8. A low power thruster, comprising: 
a combustion chamber; 
an injector shell with an interior cavity positioned 
adjoining said chamber and having centrally lo- 
cated perpendicular to its longitudinal axis an inte- 
rior perimeter describing the base of a concial por- 
tion of the cavity with an aperture at its apex open- 
ing into said chamber; 
a plurality of injector shell recesses oriented tangen- 
tially to said interior perimeter and off-set from the 
longitudinal axis; 
an equal plurality of atomizing means formed by tip 
regions of said recesses and arranged to discharge 
into said cavity tangentially to the interior perime- 
ter; 
an equal plurality of injector tubes discrete from the 
injector shell and leading directly into the recesses; 
and 
means for decomposing a propellant inside the cham- 
ber. 
9. The thruster set forth in claims 7 or 8 wherein said 
means for decomposing the propellant comprises screen 
pack consisting essentially of platinum and having a 
surface area exposed to said aperture. 
10. The thruster set forth in claim 9 wherein the pro- 
pellant passing through the aperture impinges upon 
approximately one-half of said screen Dack surface area. * * * * *  


Low thrust monopropellant engine

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