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Патент USA US3035806

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May 22, 1962
Filed NOV. 2l, 1958
2 Sheets-Sheet 1
May 22, 1962
Filed NOV. 2l, 1958
2 Sheets-Sheet 2
United States Patent 0f ice
Patented May 22, 1962
The largest volume of sustainer propellant can be used
by employing a cylindrical internal burning propellant
grain having a central internal perforation extending be
Cecil A. Glass, China Lake, Calif., assigner to the United
tween its ends which is bonded to or rests snugly against
States of America as represented by the Secretary of
the Navy
the motor tube on its outer (non-burning) surface. This
propellant extends the full length of the rocket motor
tube. With the tandem arrangement of this device, the
“tube action” propellant takes up a substantial portion of
the available motor tube length, leaving only a portion of
Filed Nov. 21, 1958, Ser. No. 775,662
4 Claims. (Cl. 244-122)
(Granted under Title 35, U.S. Code (1952), sec. 266)
rIhe invention described herein may be manufactured 10 length of the envelope for the sustainer grain, thereby
preventing attainment of maximum power. Furthermore,
and used by or for the Government of the United States
in the present‘state of the art, this particular margin of
of America for governmental purposes without the pay
power is very critical in attaining reliable ground level
ment of any-royalties thereon or therefor.
ejection capacity. Also, the separate microswitch circuit
This invention relates to combined inner tube action
for igniting the :free ñight propellant grain increases the
possibility of malfunction or involuntary firing.
and rocket action aircraft ejection seats and in particular
to those ejection seats which incorporate the inner tube
and rocket motor into a dual thrust single assembly.
A short discussion of free volume within the internal
perforation of a cylindrical hollow grain will give a better
understanding of the present invention. The free volume
. FIG. l generally illustrates a typical combined tube
action and rocket action ejection seat. It is used princi
pally in connection with military aircraft.
The term 20 or void space within an internal burning propellant pro
“tube action” and “rocket action” are derived from the
two types of propulsion used in ejection. The seat is
propelled the few feet necessary to leave the cockpit lby
tube action, that is, contained combustion gases within a
launching tube Working against a sliding member or piston 25
which has a known area.
In’dual thrust rocket termi
nology thisís called the “booster” phase of propulsion._
After leaving the cockpit, the seat is propelled by free
flight rocket action. In dual thrust rocket terminology
vides the exit channel for the passage of the products of
combustion in reaching the nozzle. Depending upon the
particular propellant and its configurations this free vol
ume must equal 0.1 to 0.4 of the total volume of the com
bustion chamber for satisfactory performance. Anything
affecting the use of this free volume as an exit channel,
such as the presence of objects that otherwise occupy the
space, causes unsatisfactory performance. The constric
tion of the exit channel causes an increase in velocity of
this .is called the sustainer phase of propulsion. It is the 30 the hot combustion gases resulting in excessive scrubbing
and wearing of the propellant surface, which is termed
goal in the design of such ejection seats to provide free
» flight travel to an elevation permitting safe parachute
“erosive burning” and subsequent excessive motor tube
escape, even when ejection is initiated at ground level,
as in take-off and landing situations.
` 'Two‘prior art devices are known and will be described.
In one device the combustion gases for the lbooster or tube
tion comprises a simplified economical structure which
Generally stated, the improvement of the present inven
permits the combination of a booster tube or inner cham
ber, for producing the gases for the tube action, within
the internal perforation of an elongated hollow internal
burning grain, while later making effective use of the space
chamber leads into a port in the breech end of the
launcher tube so that the combustion ` gases expand 40 occupied by the booster tube as free volume during sus
action are produced in a separate combustion chamber
located externally of the launcher tube. The separate
through the launcher tube. The piston and free flight
rocket are combined in a single «assembly which slideably
engages within the launcher tube. Ignition of the free
flight rocket is effected by a separate barrier assembly
tainer grain ñring. The effectiveness of the present inven
tion is evident when it is realized that the typical internal
burning propellant grain suitable for the above referred to
envelope size has an internal perforation diameter of 11A
which is slideably engaged in the launcher tube and 45 inch. It has been found possible, as a result of this inven
which travels »outwardly with the piston. Initially in the
tion, to place a 1 inch O_D. booster tube inside this 1%
tube action, this barrier prevents the hot combustion
inch internal perforation without causing unsatisfactory
gases produced in the separate chamber from reaching the
performance through erosion burning.
rocket nozzles. However, the barrier assembly contains
It is an object of the present invention to provide an
a hole with a plug which is mechanically sheared olf `just 50 upward directed aircraft ejection seat with booster tube
prior to the piston and rocket motor assembly leaving
thelauncher tube. Hot gases pass through the hole and ¿ »and free flight rocket in the same assembly, with suffi
cient power to permit reliable ground level ejection
then into the rocket motor nozzle, -igniting the free flight
propellant. In this device the separate combustion cham
It is another object of the present invention to provide
-ber takes up critical space within the aircraft and adds 55
to the cost of manufacture. The barrier assembly is a
separate moving part, which increases the possibility of
within an aircraft ejection seat system having a given
, envelope size for the booster and free flight propulsion
means, more ejection power than heretofore available.
It isa further object of the present invention to provide
In -another prior art device the com-bustion chamber for
producing gases for the tube action, the combustion cham 60 a dual thrust rocket motor wherein a booster tube may be
ber for producing gases for the free flight rocket action,
disposed within the internal perforation of an internal burn
and the piston are combined into a single assembly. In
ing propellant grain without sacrificing the use of the
other words the piston is also a dual thrust rocket. The
space occupied by the booster tube as effective free volume
two propellant charges are disposed in tandem, that is,
during sustainer grain firing.
one behind the other within the length of the rocket motor. 65
A still further object of the present invention is to pro
In this device, the free Hight propellant is ignited elec
vide within a dual thrust rocket a >simple means, requiring
trically by microswitch means, just prior to the time the
rocket motor leaves the launcher. The tandem arrange
no extra parts, for the timed ignition of the sustainer pro
,ment results in substantially less space for available rocket
Other objects and many of the attendant advantages of
motor power. It is generally recognized that the envelope 70
this invention will be readily appreciated as the saine
size available for a seat ejection power system is -a space
becomes better understood by reference to the following
approximately 31/2 xfeet long and having a 3-inch diameter.
detailed description when considered in connection with
wise cause undesired 4tumbling during free ñight and
the -accompanying drawings wherein:
secondly, yto provide a forward thrust component to soften
the rearward- impact which occurs duringa high speed
FIG. 1 illustrates the use of lan aircraft ejection seat
with the present investiert.
ejection when the pilot is abruptly catapulted into a high
FIG.. 2 is a longitudinal eentral eeetien efY the launcher
tube. reeket meter tube, au@ buuster tube ,asserublnk
PIG. 3 is an enlarged seetieu taken. .alena .line .3v-ê. _ef
FIG. 2,
ries. 4A, 4a and 4c are sesame 'a
ius Seguential conditions during eis
speed wind blast `as illustrated in FIG. 1. Nozzles 42 con
tain tightly ñtted blow out plugs `44, which remain iixed in
these nozzles until the sustainer propellant grain Z8 is
_ Ígllîîçd.- Theseplugs are made oiga relatively heat>V resistant
' material such as aluminum and prevent premature igni
ti-on of the sustainer propellant grain by entry of hot tube
V"Ille maier Components of acera> >.ined tubeaetien .aud "
gases inte Vthe main chamber through the nozzle parte
reeltet aetiuu airerait eieetieu seat will. be, apparent. fr0-ul
au examination of the illustrative example. 'shawn iu the?
drawing,- A launcher tube 1,0y having la 1‘l1-teach .chamberY
Nozzle 46lcommunic`ates with the booster tube through an
1_2 iS Secured t0 the aircraft frame, by `any suitable trunnion
elbow euuueetiOu-‘Stltur exhausting the-eerrlbustiun eases
of the booster tube. lA conventional nozzleclosure di
aphragm 48 made et neoprene. ,is similarlyftttediuteuuz
1,1. A rocket motor Iassembly 2_0, having the, dQllble func- `
zle 46 to prevent dust or moisture from entering the booster
tion of piston and thrust rocket motor, is slideably en
¿aged within the launcher tube.' The assembly includes
the main motor tube 2,2., a smaller inner booster tube 30
and a nozzle assembly 40. The ejection seat 23 is rigidly
It is >to be understQod'that-the pilot initiates seat eiee.
tion by actuating some, conventional mechanism that tigres,
the ignite: 32 which in turn ignites thebooster propellant
attached to the motor tube 22 by means of a »suitable
strips 36, 36' withinrfthe booster. tube. Typical flow
bracket 2,4 and clamping ring 25.
paths of the combustion gases at this moment are shownn
Now referring to the rocket motor assembly 20 shown
by the arrows in FIG. 4A. = Since theapertures 34».alo1:lg>
in FIGS. 2. and 3, the Outer casing or motor Atube 22, the
' the length of the booster 30 are covered by a membrane.
nose cap 2,6, and the nozzle assembly 4t2, form a main 25 consisting of the propellant strips and since as a point of
design the web> thickness and strength of the propellant yis
designed to have an'initial Yrupture strength greater than
the booster tube combustion pressure, the gases .are C911,
28 is bonded tot or rests snugly against the insidewall of
s_trained'to ilow through-the booster tube nozzle 46 intQ
the motor tube 22y and is mechanically supported in the 30 the launcher tube breech chamber 12.. These gases will
combustion chamber to allow for thermal expansion by
work against the rocket .motor assembly 20 and provide
means of resilient spacers, springs or other structural de
the tube action while blow-out plugs 44. prevent these
vices known in the art.
gases from entering Vnozzles 42. As understood in the
A booster tube 30 is disposed within the central per
art, the cockpit canopy is cast oñ, the pilot is constrained
combustion chamber for producing the sustainer or free
flight combustion gases. The outer non-burning surface
of a hollow cylindrical internal burning '_sustainer grain
f_oration of the propellant grain 2,8, its ends being rigidly
35 to the seat, and the seat, being rigidly secured to the
secured to the nose> cap Z6 and nozzle assembly 40. VThe Y
rocket motor .assembly 20, YtravelsY upwardly, guided by
outer diameter of tube 30 is slightly smallerY than theref
fective diameter of the propellant perforation so as to
leave free volume within the main combustion chamber
all around the booster tube, The end of the booster tube
adjacent the nose cap 2.6 contains a conventional> igniterv
32 and is closed -so that the booster tube forms a small
inner combustion chamber for the production of the gases
for the ejection of the rocket rrrotor ‘assemblyV and Vman
seat out-of the launcher. Although not shown in the draw
ing, suitable igniter firing accessories such as sear' and
firing pin may be incorporated into the nose> lcap. The.
booster tube has a plurality of apertures 3.4. At- the start
o_f ejection these apertures are coveredby :booster fp_ro
v. As the rocket motor assembly travels up the launcher
tube, burning takes place normal >to the exposed burning
surfaces of the propellant strips 36, ¿36C The rocket
rnotor assembly 20 travels up the launcher tube until it
reaches a «point where it is Within a few inches of leav
ing the tube as illustrated in FIG. 4B.V At that moment
the membrane over the booster tube apertures 34, consist
ing of the propellant stripsa36, 36' rupture under the
booster tube combustion pressure as the result of reduc
tion in the web thickness of the propellantstrips 36, 36.’
by burning; The precise timing of this event is predeter
mined by the- choice of initial web thicknessfor a propelf
pellant strips 36„ 36', as best shown in FIGQ3, which have 50 lant s_triphaving a givenburning rateand given rupture
a semi-circular section and areV bondedto the linside wall
strength characteristics. The flow paths oÍf the combus
0f the beoster tube- Thusthe web. thieliuese Qi these pre
tion gases at thisv moment are shown by the arrows in
pellant strips form a membrane over the holesV and only
FIG. 4B. -A portion/of the hot booster tube gases ñoW
the inner surface of- the propellant is exposed forY burn
through holes 34 into the free volume within the main
ing within the booster tube. The choice of the particular 55 combustion chamber. This’ moment marks the begin
booster propellant is dictated by the desired burning char
acteristics. Double base formulations havinglmesa-burn
ing characteristics, i.e., uniform burning rate over a rela- „
tively wide temperature range have been found to work
elîectively. In the exemplaryl form, the strips are bonded 60
to the inner wall by a bonding Iagent 3:8V consisting of cot
ton tape, as shown in FIG. 3, that had been dipped in
polysiloxane adhesive. As is common in ftheV art, inhibit
ing tape 39, bonded to the propellant, may be used to
inhibit a portion of its exposed burning surface so as to 65
obtain particular burning characteristics.
ning of the transition between tube action Vand rocket
action. It is torbe noted at this> moment, justV prior to
ignition of- the sustainer the Vblow-out plugs 44 are still
in place and a portion of the booster gases is still exhaust
ing through'nozzle 46. .
After a very brief ignition delay, for no more than the
short time required for the piston and'rock’et motor as
sembly to travel the remaining few inches within the
launcher tube, the Ysustainer propellant, 28 ignites and
blow-outvplugs 44 are ejected from nozzles V42.V The flow
paths ¿of the sustainer propellant combustion gases an
instant after ignition are s_hown in FIG.'4C,_ thebooster
A multiport nozzle `40 containing several convergent
divergent type nozzles 4,2A communicating Vto, the main
propellant strips having burnedY away( A portion of the
combustion chamber and ya similar nozzle communicat
gases exhaust through n_ozzles 42,V passing'through the
ing to the booster tube is secured to the end of the motor 70 free >volumes between the burning surface of the propel
tube. The nozzle is canted at an angle which results in
the axes of the nozzles passing near the center of gravity
of the combined man-seat mass as shown in'FIG. 1'.> The
purpose of this is to avoid applying unwanted unbalanced
lant grain and the booster tube._; jSimultaneouslyÍ another
portion ofthe combustion gases passesthrough 'the aper
tures 34_into fthe boostentubeand exhaust y‘through
booster tube nozzle 46„ ‘thus lusing the freevolume of
torque to the man-seat combination which would other-ÍY 75 the booster tube >.during sustainer .grain tiring. This
marks the end of the transition from tube action to rocket
action and the combined pilot and seat is now propelled
as a free flight reaction rocket vehicle and the direction
of thrust has changed from the direction of arrow T to
to the piston assembly, the piston assembly being further
adapted to provide combustion gases for the ejection of
arrow T’ wherein the resultant passes near the center of
piston assembly leaves the launcher tube; the improve
gravity as shown in FIG. 1.
the piston assembly out of the launcher tube and com
bustion gases for reacting rocket free flight after the
ment wherein said piston assembly comprises, an elon
After the ejection seat has been shot high enough to
gated main combustion chamber having a first propellant
make a safe parachute escape, delay initiators release
grain therein and adapted to produce said combustion
the pilot from the seat and his attached parachute auto
gases for reaction rocket free flight, an elongated inner
matically opens by means heretofore known in the art. 10 combustion chamber having a second propellant therein
Obviously many modifications and variations of the
and adapted to produce the said combustion gases for
present invention are possible in the light of the above
the ejection of the piston assembly out of the launcher
teachings. It is therefore to be understood that within
tube, said inner chamber being disposed concentrically
the scope of the appended claims the invention may be
within the main chamber and directly communicating
practiced otherwise than as specifically described.
with the interior of said launcher tube to pressurize same
I claim:
with gases produced by said second propellant, aperture
,Y 1,. In apparatus for use in a tube action rocket action
means extending through the wall of the inner chamber,
aircraft ejection seat having a launcher tube attached to
said aperture means adapted to be closed at the time ejec
the aircraft and rocket motor assembly attached to the
-tion starts, means for opening the aperture means after
seat and slideably engaged within the launcher tube; the 20 the piston assembly has traveled a predetermined distance
improvement wherein said rocket motor assembly com
along the launcher tube thereby permitting the combus
prises; an elongated rocket motor tube, an internal burn
tion gases produced in the inner chamber to pass into and
ing cylindrical propellant grain disposed within the ma
ignite the propellant in the main combustion chamber,
jority of the length of said motor tube having a central
said aperture means being further adapted to permit a
perforation extending between its ends, one end of said 25 portion of the gases produced in the main chamber to
motor tube having a first nozzle means for exhausting
pass into and exhaust through the inner chamber; and
combustion gases produced by said propellant grain when
nozzle means to exhaust said combustion gases for free
ignited, said first nozzle means having vblow-out means
fiight constructed and arranged to exhaust angularly to
which remain therein until said grain is ignited, the
the longitudinal axis of said elongated combustion cham
other end of said motor tube being closed, a booster tube 30 ber to produce a component of thrust passing through
disposed Within said perforation, said booster tube having
said seat.
an igniter at the end thereof adjacent the closed end of
3. A piston assembly as described in claim 2 wherein
said motor tube, and having at the other end thereof
said aperture means comprises a plurality of apertures
second nozzle means for exhausting combustion gases,
and the means for opening the aperture means comprises
said second nozzle means being adjacent said iirst nozzle 35 said second propellant being formed from strip propel
means, said first and second nozzle means being con
lant having a predetermined initial web thickness mounted
structed and arranged to exhaust angularly to the lon
over the apertures on the inside surface of .the inner
gitudinal axis of the` motor tube to produce a com
chamber, the strip propellant being adapted to act as a
ponent of thrust passing through the seat, said booster
membrane in closing the apertures to passage of com
tube having a plurality of apertures along the length 40 bustion gases produced in the chamber into the main com
thereof and extending through its walls, said apertures
bustion chamber at the start of ejection and being further
being closed with strips of propellant fixed to the inner
adapted to rupture at a predetermined time after the
surface of said booster tube, said igniter being operable
start of ejection as a result of the decrease in membrane
to ignite said strips of propellant for producing combus
web thickness and strength resulting from burning.
tion gases, said combustion gases produced by said strips 45
4. A piston assembly as described in claim 3 wherein
of propellant being operable to cause said rocket motor
assembly to travel out of the launcher tube, said strips
of propellant being further adapted to form a membrane
said nozzle means comprises a nozzle communicating with
the main chamber for exhausting combustion gases pro~
duced therein, said nozzle initially closed by a blow-out
over said apertures and prevent combustion gases pro
plug to prevent premature ignition of the main chamber
propellant by entry of gases produced in the inner cham
ber therethrough, and a nozzle communicating with the
inner chamber for exhausting the gases produced therein
and exhausting said portion of the gases produced in the
main chamber which exhaust through the inner chamber.
duced by said strips from passing through said apertures
until said rocket motor assembly has moved to a prede
termined position along said launching tube at which po
sition said strips of propellant are adapted to rupture
and permit hot combustion gases produced by said strips
to pass through said apertures, said hot gases passing 55
through said apertures being operable to effect the igni
tion of said propellant grain, the combustion gases pro
duced by said propellant grain being operable to pro
duce thrust to project the seat away from the aircraft
after said rocket motor assembly leaves the launcher tube, 60
a portion of said combustion gases produced by said pro
pellant grain after ignited adapted to pass through the
apertures in the booster tube whereby gas is simultaneous
References Cited in the file of this patent
Hickman ____________ .__
Hickman ____________ _Edelman et al. _______ _..
McKnight ___________ _..
Fox ________________ _...
Hirt et al. ___________ .__
Jan. 20, 1948
Nov. 22, 1955
ly exhausted through said first and second nozzle means.
2. In an aircraft seat ejection system having a launcher 65
tube secured to the aircraft, a piston assembly slideably
Aviation Week Magazine, Nov. 12, 1956, vol. 65, No.
engaged in the launcher tube and a seat fixedly attached
20, pages 71, 72, 74 and 77 relied upon.
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