Ignition system for rocket engines

An ignition system for a rocket engine which improves the reliability of engine re-ignition by supplying additional oxygen to the ignitor directly from the rocket engines main fuel/oxygen injector elements, thereby eliminating the need for a supplemental oxygen supply system.

TECHNICAL FIELD 
This invention relates to ignition systems for liquid fueled rocket 
engines, and more specifically, to an efficient system which increases the 
reliability of engine re-lights on multiple ignition launches. 
BACKGROUND OF THE INVENTION 
Liquid fueled rocket engines are commonly used as upper stage propulsion 
systems on multiple stage launch vehicles. During a typical launch, for 
example, the placing of a satellite in near Earth orbit, an upper stage 
rocket engine may fire briefly, then coast, then fire again. Multiple 
firings of an engine during a single launch requires a highly reliable 
ignition system that is capable of multiple engine re-lights. 
Ignition systems of the prior art typically include a supplemental oxidizer 
supply line to provide additional oxidizer to the region surrounding the 
engine's ignitor during ignition to ensure proper lighting for re-lighting 
of the fuel/oxidizer mixture in the rocket engines combustion chamber. 
While supplying additional oxidizer to the ignitor has proven to produce a 
desirable fuel/oxidizer ratio at ignition, the supplemental oxidizer 
supply line requires a shut-off valve to avoid over-heating. Combustion 
products back flow into the valve during the start-up pressurization of 
the engine, causing the valve to freeze closed. With the valve frozen 
closed for subsequent re-light attempts, the engine will not light 
reliably. Attempts to accommodate and/or eliminate back flow have proven 
to be ineffective at preventing freezing of the shut-off valve. 
What is needed is an ignition system for liquid fueled rocket engines that 
increases the reliability of engine re-light on launches requiring 
multiple firings of the engine, while reducing the cost and weight of 
ignition systems of the prior art. 
SUMMARY OF THE INVENTION 
It is therefore an object of the present invention to provide an ignition 
system for liquid fueled rocket engines that increases the reliability of 
engine re-light on launches requiring multiple firings of the engine. 
Another object of the present invention is to reduce the weight of the 
ignition systems for liquid fueled rocket engines. 
Another object of the present invention is to reduce the cost of the 
ignition systems for liquid fueled rocket engines. 
Accordingly, an ignition system for a rocket engine is disclosed having an 
ignitor, a faceplate having a centrally located ignitor port and first and 
second annular sections located radially outward of the port, the first 
annular section located radially inward of the second annular section, the 
ignitor mounted to the ignitor port, a plurality of first injector 
elements located in the first annular section of the faceplate and a 
plurality of second injector elements located in the second annular 
section, each of the injector elements having a first orifice defining a 
first flow area, and a second orifice defining a second flow area, the 
second orifice concentric with and radially outward of the first orifice 
relative to the first orifice and having a radially inner half and a 
radially outer half, the radially inner half located between the radially 
outer half and the ignitor port, each of the first injector elements 
located equidistant from the ignitor port, and, at least one insert 
located in the second orifice of one of the first injector elements, 
thereby reducing the second flow area of the second orifice in which the 
insert is located. 
The foregoing and other features and advantages of the present invention 
will become more apparent from the following description and accompanying 
drawings.

BEST MODE FOR CARRYING OUT THE INVENTION 
As shown schematically in FIG. 1, a typical rocket engine 10 includes a 
thrust chamber 12, a combustion chamber 14 upstream of the thrust chamber 
12, a faceplate 16 which contains a plurality of fuel/oxidizer injector 
elements 18 and an ignitor port 20, and an ignitor 22 mounted to the 
ignitor port 20, a fuel supply line 24 and an oxidizer supply line 26 for 
supplying fuel and oxidizer to the fuel/oxidizer injector elements 18, and 
a supplemental oxidizer supply system 28 for supplying additional oxidizer 
to the ignitor 22 during ignition. 
As shown in FIG. 2, the faceplate 16 of the present invention has a 
centrally located ignitor port 20 and the ignitor 22 is mounted to the 
ignitor port 20. Radially outward of the port 20 is a first annular 
section 30 of ignitor elements 18, and radially outward of the first 
annular section 30 is a second annular section 32 of ignitor elements 18. 
A plurality of first injector elements 18' are located in the first 
annular section 30 of the faceplate 16 and a plurality of second injector 
elements 18' are located in the second annular section 32. Each of the 
injector elements 18 has a first orifice 34 defining a first flow area, 
and a second annular orifice 36 defining a second flow area. The second 
annular orifice 36 is concentric with and radially outward of the first 
orifice 34 relative to the first orifice 34. In the ignition system of the 
present invention, the first orifice 34 of each injector element 18 
provides oxidizer to the combustion chamber 14, and the second annular 
orifice 36 provides fuel to the combustion chamber 14. 
Each of the first injector elements 18' is located equidistant from the 
ignitor port 20, and as shown in FIG. 4, the second annular orifice 36 of 
each of the first injector elements 18' has a radially inner half 38 and a 
radially outer half 40. The radially inner half 38 is located between the 
radially outer half 40 and the ignitor port 20. At least one, and 
preferably several of the first injector elements 18' includes an insert 
42 located in the second orifice 36 thereof. As shown in FIGS. 4 & 5, the 
insert blocks a portion of the second annular orifice 36, thereby reducing 
the second flow area of the second annular orifice 36 in which the insert 
42 is located. 
As shown in FIGS. 5 & 6, the insert 42 comprises a semi cylindrical wall 44 
having a thickness 46 that is only slightly less than the width 48 of the 
second annular orifice 36 in which the insert 42 is located. As used 
herein, the term "semi cylindrical wall" means a body generated by 
rotating a rectangular area having a thickness "T" and a height "H" at a 
distance "R" about a reference axis "A" for less than one complete 
revolution. For the insert 42 of the present invention, the semi 
cylindrical wall 44 results from one half of a rotation about the 
reference axis A, and the thickness 46 is slightly less than the radial 
width 48 of the second annular orifice 36. As used herein, the term "width 
of the second annular orifice" means the distance 48 between the radially 
inner surface 50 of the second annular orifice 36 and the radially outer 
surface 52 of that annular orifice 36, as shown in FIG. 5. A flange 54 
which is integral with one end 56 of the semi cylindrical wall 44 extends 
radially outward therefrom in a direction that is substantially 
perpendicular to the reference axis "A". As shown in FIG. 6, the length 58 
of the flange 54 is preferably at least half of the height "H" of the semi 
cylindrical wall 44. The insert 42 is preferably made of the same material 
as the faceplate 16, but may be made of a similar material having similar 
material properties as the material from which the faceplate 16 is made. 
As shown in FIG. 4, in the preferred embodiment of the present invention 
there are six injector elements 18' in the first annular section of the 
faceplate 16, and the injectors 18' are spaced equally about the 
circumference of the ignitor port 20. Inserts 42 are located in the 
radially inner half 38 of the second orifice 36 of at least half of the 
first injector elements 18'. As those skilled in the art will readily 
appreciate, the insert 42 reduces the second flow area of the second 
orifice 36 in which the insert 42 is located by approximately 50 percent, 
and since fuel is delivered through the second annular orifice 36 of each 
of the injectors 18, fuel flow through each of the first injector elements 
18' in which the insert 42 is located is reduced by approximately 50 
percent. Consequently, the inserts 42 produce a fuel-lean (oxidizer-rich) 
zone immediately adjacent to the ignitor elements 18' in which the inserts 
42 are located. Preferably, each of the first injector elements 18' in 
which an insert 42 is located is immediately adjacent another first 
injector element 18' in which an insert 42 is located and immediately 
adjacent another first injector element 18' in which an insert 42 is not 
located. Preferably, the first injector elements 18' in which inserts 42 
are not located are immediately adjacent two first injector elements 18' 
in which inserts 42 are located. The inserts 42 are fixedly secured to the 
faceplate 16, preferably by welding the flange 54 thereto. 
Immediately prior to engine ignition, fuel and oxidizer are fed to all of 
the injector elements 18 by the fuel supply line 24 and the oxidizer 
supply line 26, respectively. Due to the presence of the inserts 42 in the 
radially inner half 38 of the second annular orifice 36 of some of the 
first injector elements 18', fuel spray from the second annular orifice 36 
of those injector elements 18' is directed away from the ignitor port 20, 
while a portion of the oxidizer spray from the first orifices 34 of the 
same injector elements 18' is directed toward the ignitor port 20. The 
injector elements 18' in which the inserts 42 are located therefore 
produce an oxidizer rich zone on that side of the injector element closest 
to the ignitor port 20. 
On the side of the injector element 18' furthest from the ignitor port 20, 
the injector elements 18' in which the inserts 42 are located produce a 
core spray 66 of oxidizer wrapped in a curtain of fuel 60, as shown in 
FIG. 7. The shear surface between the oxidizer core spray 66 and the fuel 
curtain 60 produce a mixing zone 68 which provides a combustible 
fuel/oxidizer mixture. At engine ignition, the ignitor injects fuel 70 
into the oxidizer rich zone 64 created by the first injector elements 18' 
in which the inserts 42 are located, as shown in FIG. 7. Then, following a 
spark from the ignitor 22, the fuel is energized to 5000 degrees R, 
igniting the fuel in the oxidizer rich zone 64, which then propagates to 
the mixing zone 68 and on to the other injector elements 18. 
As those skilled in the art will readily appreciate, the first injector 
elements 18' in which the inserts 42 are located provide the oxidizer rich 
zones 64 which are necessary for reliable ignition without the need for 
the supplemental oxidizer supply system 28 of the prior art. Accordingly, 
rocket engines incorporating the ignition system of the present invention 
avoid both the cost and ignition problems associated with the supplemental 
oxidizer supply system 28 of the prior art, while the eliminating the 
weight of the supplemental oxidizer supply system 28 from the rocket 
engine. 
Although this invention has been shown and described with respect to a 
detailed embodiment thereof, it will be understood by those skilled in the 
art that various changes in form and detail thereof may be made without 
departing from the spirit and scope of the claimed invention.