Pressurized gas injection for burning rate control of solid propellant

Heidmann, M. F.

Povinelli, L. A.

English

NASA Technical Reports Server

REPLY ro 
ATTN OF: GP 
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 
WASHINGTON. D.C. 20546 
March 27, 1971 
TO : USI/Scient i f ic  & Technical Information Divis ion 
At ten t ion :  M i s s  Winnie M. Morgan 
FROM : GF/Office of A s s i s t a n t  General 
Counsel f o r  Paterlt Matters 
SUBJECT: Announcement of NASA-Owned 
U.S. Pa ten ts  i n  STAR 
I n  accordance wi th  t h e  procedures contained i n  t h e  Code GP 
t o  Code USI memorandum on t h i s  subject, dated June 8 ,  1970, 
t h e  a t t acked  NASA-owned U.S. p a t e n t  i s  being forwarded fo r  
abstract ing and announcement i n  NASA STAR. 
The following information i s  provided: 
U.S. Pa ten t  No. Z 38345,822 
Corporate Source : Lewks Research Center 
Supplementary 
Corporate Source s 
NASA Pa ten t  Case No.: =E-03494 
Gayle Parker 
Enclosure: 
Copy of Pa ten t  
NASA-HQ 
https://ntrs.nasa.gov/search.jsp?R=19710012344 2020-03-17T02:48:56+00:00Z
Oct. 10, 1967 L, A. POVlNELLl ETAL 3,345,822 
PRO 
Filed Feb. 17, 1966 
1 4 4 -  3 
~ ~ ~ ~ ~ , ~ 2 2  United States Patent Ofice pateneed 
1 
3,345,822 
BURNING RATE CONTROL OF SOLID 
PROPELLANTS 
Louis A. Povinelli, MiddIebaug Weights, and Marcus F. 
Heidmann, Rocky River, Ohio, assignors to the United 
States of America as represented by the Administrator 
OE the National Aeronautics and Space Administration 
Filed Feb. 17,1966, Ser. No. 529,593 
7 Claims. (61. 60-251) 
ABSTRACT OF THE DISCLQSURE 
Controlling the burning rate of solid propellants by in- 
jecting pressurized hot gases into a rocket combustion 
chamber. These gases are injected tangentially around the 
longitudinal wall of the combustion chamber in a direc- 
tion transversed to the longitudinal axis of the rocket 
motor. 
The invention described herein was made by employees 
of the United States Government and may be manu- 
factured and used by or for the Government of the 
United States of America for governmental purposes 
without the payment of any royalties thereon or therefor. 
This invention relates to solid propellant rocket motors 
and more particularly to apparatus for regulating the 
mass burning rate of a solid propellant. 
Solid propellant rocket propulsion systems could be- 
come a considerably more praotical and versatile form 
of propulsion if they are provided with an accurate and 
reliable means of thrust control. The presently known 
techniques for controlling the rate of combustion of solid 
propellant rocket motors are either so complicated as to 
be impractical for reliable missile use, or are so bulky 
that the size advantages of solid propellant missiles are 
sacrificed. 
Ilt is therefore an object of the invention to provide a 
solid propellant rocket motor system wherein rocket 
thrust may be controlled by varying the rate of propellant 
combustion. 
A further object of the invention is to provide a com- 
bustion rate control apparatus for a solid propulsion 
motor that is simple in construction, inexpensive to manu- 
facture and highly efiective in operation. 
Briefly, the foregoing objects are accomplished by pro- 
viding apparatus for injecting pressurized hot gases into 
the combustion chamber of a solid propellant rocket motor 
wherein said gases are injected tangentially around the 
longitudinal wall of the combustion chamber in a direc- 
tion transverse to the longitudinal axis of the chamber to 
control the rate of combustion of the propellant charge 
and thereby control the thrust of the rocket. A typical 
solid fuel rocket motor consists of a cylindrical metal 
shell having an exhaust port and containing a shaped 
solid propellant charge of molded propellant material 
having a longitudinal bore forming a combustion cham- 
ber wherein combustion takes place. The propellani 
charge contains all the necessary ingredients for sus- 
taining chemical combustion and, once ignited, the 
charge will burn on a11 exposed surfaces. Specifically, 
the burning takes place on the longitudinal wall of the 
propellant charge bore which forms the combustion cham- 
ber. It i s  well known that the rate of combustion in a 
solid propellant rocket varies with the pressure in the 
combustion chamber so that as the chamber pressure in- 
creases, the rate of burning also increases. It has been 
found that by directing pressurized hot gases tangential- 
ly around the periphery of the bore forming the com- 
bustion chamber in a direotion transverse to the longi- 
tudinal axis of the bore a vortex flow field may be estab- 
lished within the combustor. The vortex flow with its at- 
2 
tendant radial pressure distribution causes the effective 
throat area to be substantially reduced relative to the real 
throat area. The chamber pressure is therefore increased 
and the propellant burning rate, which is sensitive to pres- 
5 sure, is increased thereby. The tangential flow will also 
give rise to increased burning at sufficiently high velocities 
due to erosive effects. The rate of combustion of the 
solid propellant may be accurately controlled, and hence, 
the thrust of the rocket is correspondingly controlled. The 
10 pressurized hot gases may be supplied by a conventional 
auxiliary solid propellant chclrge gas generator having an 
igniter, all of standard construction. As a modification, 
the gas generator may be a hydrogen peroxide generator 
with a catalyst bed. The tangential injection of the hot 
15 gases establishes a vortex flow which controls the rate of 
combustion by causing an increase in chamber presusre 
as well as by causing erosive burning. In addition, a 
slight increase in burning rate results from the adjustment 
of the chamber pressure to the new burning rate. 
Other objects and advantages of the invention will be 
apparent from the following description taken in conjunc- 
tion with the drawings wherein: 
FIGURE 1 is a side elevational sectional view of a 
solid propulsion rocket motor constructed in accordance 
FIGURE 2 is a perspective sectional view of the rear 
portion of ,the rocket motor shown in FIGURE 1, with the 
auxiliary gas generator having a cut-away portion to 
show the interior thereof; 
FIGURE 3 is a sectional view taken along the line 2-2 
of FIGURE 1; 
F I G m E  4 is a sectional view similiar to that of HIG- 
URE 3, but showing a modification of the device shown 
therein. 
Although the invention is shown and described herein 
with reference to its use in a solid propellant rocket 
motor, it will be understood that it can effectively be used 
in any gas-generating device for controlling the rate of 
combustion therein. 
Referring to the drawings, there is shown a solid pro- 
pellant rocket motor 8, constructed in accordance with 
the invention, and including a hollow cylindrical shell or 
casing 18 having an exhaust or thrust nozzle 12. The cas- 
ing 10 contains a conventional solid propellant charge 
45 14 having a bore 16 forming a combustion chamber 18 
contiguous with the nozzle 12. The propellant charge 14 
is adapted to produce combustion products in the com- 
bustion chamber 18 which discharge through the nozzle 
12 to produce thrust in the usual manner. The propeIlant 
50 charge may be a mixture of synthetic rubber and an in- 
organic oxidizer such as NH.&XO,. 
The invention is provided with a means for tangential- 
ly injecting hot gases into the combustion chamber 18 as 
will now be described. Exteriorly of the casing 10. there 
55 is provided, in the preferred form, an auxiliary gas gen- 
erator 24 including a hollow cylindrical casing 30 having 
a solid propellant charge 32 therein. The propellant 
charge 32 has a longitudinal bore 34 forming a combus- 
tion chamber 35 and may be of the same matterial as the 
60 charge 14. A conventional igniter 36 is operatively 
secured to the end of the casing and is contiguous with the 
chamber 35 for igniting the charge 32 in the usual manner, 
or, as an alternate, the generator 24 may be provided with 
a hydrogen peroxide tank 37 as shown in dotdash lines 
65 in FIGURE 1. The generator 24 may be any other suit- 
able type of gas generator such as a hydrogen peroxide 
generator with a decomposition catalyst bed to &low for 
greater frequency of ,thrust modulation. The generator 24 
contains an elongated injection tube operatively con- 
70 nected at one end to, and in communication with, the gen- 
erator combustion chamber 35, the other end of the *tube 
40 extending longitudinally into the combustion chamber 
20 
25 with the invention; 
30 
35 
40 
3,345,822 
3 4 
18 of the rocket motor 8. The tube 40 has a pluraIity of of spaced discharge ports for injecting the hot gases trans- 
spaced ports or outlets 42 for injecting the hot gases tan- versely tangentially around the longitudinal wall of the 
gentially into the chamber 18 as shown by the arrows in bore to control the rate of combustion of the propellant 
FIGURE 3. The hot gases are injected at as high a veloci- charge. 
ty as is practicable. 2. The structure of claim 1 wherein said means by 
To provide for a more continuous and even flow of the which a vortex flow is generated comprises a plurality of 
hot gases into the chamber 18, additional elongated injec- said elongated injection tubes operatively connected to 
tion tubes 40a and 40b may be longitudinally disposed in said generator, said tubes being longitudinally disposed in 
the chamber 18 and spaced circumferentially therearound the bore and spaced circumferential around the inner 
as shown in FIGURE 4, such tubes 400 and 40b be- longitudinal wall of the bore. 
ing connected, in one form of the invention, to the pipe 3. The structure of claim 1 wherein said gas generator 
40 at a point between the casings 30 and 10, or each tube comprises an auxiliary solid propellant charge having an 
40, 4Qa and 40b may be connected to a separate gas gen- igniter operatively secured thereto. 
erator. 4. The structure of claim 1 wherein said propellant 
The invention also contemplates the method of inject- 15 charge is a mixture of an inorganic oxidizer and a syn- 
ing hot gases tangentially into a solid propellant combns- thetic rubber. 
tion chamber t o  control the combustion rate therein 5. The structure of claim 4 wherein said inorganic 
by providing a solid propellant rocket motor having a oridizer is NH4C104. 
solid propellant charge with a longitudinal bore forming 6. The structure of claim 1 wherein said gas generator 
a combustion chamber, and providing a source of pres- 20 is a hydrogen peroxide generator with a catalyst bed, 
surized hot gases and injecting said pressurized hot gases 7. A solid propellant rocket motor comprising, a casing 
from said source into said combustion chamber tangen- having an exhaust nozzle, a combustible solid propellant 
tially around the longitudinal wall of the combustion charge disposed in said casing and having a longiltudinal 
chamber in a direotion transverse to the longitudiiial axis bore forming a combustion chamber contiguous with the 
of the chamber to control the burning rate of the propel- 25 nozzle, said propellant charge adapted to produce com- 
lant charge. bustion products in said combustion chamber which dis- 
With such tangential injection, burning rate increases of charge through the nozzle to produce thrust, and means 
up  t o  75% greater than previous burning rates are for injecting pressurized hot gases into said cornbustion 
achieved. chamber tangentially around the longitudinal wall of the 
The terms and expressions which have been employed 30 bore in a direction transverse to the longitudinal axis of 
are used as terms of description, and not of limitation, and the bore to control the rate of combustion of the propel- 
there is no intention, in the use of such terms and expres- lant charge by the establishment of a vortex flow field 
sions, of excluding any equivalents of the features shown within the combustion chamber; said means including a 
or described, or portions thereof, but it is recognized that plurality of gas generators each having an elongated injec- 
various modifications are possible within the scope of the 35 tion (tube operatively connected therewith and extending 
invention claimed. longitudinally into the bore, said tubes having a plurality 
of spaced discharge ports for injecting the hot gases trans- 
versely tangentially around the longitudinal wall of the 
bore to control the rate of combustion of the propellant 
Xeferences Cited 
5 
What is claimed is: 
1. A solid propellant rocket motor comprising, a cas- 
ing having an exhaust nozzle, a combustible solid propel- 
Iant charge disposed in said casing and having a longi- 40 charge. 
tudinal bore forming a combustion chamber contiguous 
with the nozzle, said propellant charge adapted to  produce 
combustion products in said combustion chamber which 
discharge through the nozzle to produce thrust, and means 
for injecting pressurized hot gases into said combustion 
chamber tangentially around the longitudinal wall of the 
bore in a direction transverse to the longitudinal axis of 
the bore to control the rate of combustion of 'the propel- 
lant charge by the establishment of a vortex flow field 
within the combustion chamber, said means including a 
gas generator, and at least one elongated gas injedion tube 
operatively connected to the generator and extending 
longitudinally into the bore, said tube having a plurality 
UNITED STATES PATENTS 
5/1962 Ledwith ______r______ 6 b 2 3 4  
1111962 Schultz _ _ _ _ _ _ _ _ _ _ _ _ _ _  60-254 
911963 Proell _ _ _ _ _ _ _ _ _ _ _ _  60-254 X 
6/1964 Avery _ _ _ _ _ _ _ _ _ _ _  60-254 X 
7/1964 Sessums _ _ _ _ _ _ _ _ _ _ _ _  60-251 
1211964 Blackman _ _ _ _ _ _ _ _ _  60-258 X 
1/1967 Greiner _ _ _ _ _ _ _ _ _ _ _ _ _  60-251 
R. CROYLE, Primary Examiner, 
3,031,842 
45 3,065,596 
3,102,386 
3,136,119 
3,142,152 
3,158,997 
60 3,298,181 
CARLTON 


Burning rate control of solid propellants Patent

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