Integral booster/ramjet drive

A missile drive is provided comprising a single combustion chamber shared by a first, acceleration stage and a second, ramjet cruising stage, said chamber housing the solid propellant to be consumed during acceleration. At least one air inlet is opened at the end of the acceleration stage to allow an air flow to be introduced into the combustion chamber. Moreover, at least one additional exhaust outlet forming an additional nozzle is provided in the back of the combustion chamber and means are provided to seal off said additional exhaust outlet or outlets throughout the initial acceleration stage, such that said additional exhaust outlet or outlets contribute, together with the converging-diverging nozzle, to ejecting the exhaust gas during the ramjet cruising stage.

This invention relates to a ramjet missile propulsion system or drive with 
a built-in acceleration engine or booster, said drive comprising a single 
combustion chamber, shared by a first, acceleration stage and a second, 
ramjet cruising stage, wherein is stored the solid propellant used for 
missile acceleration, and further comprising a convergent-divergent nozzle 
optimally dimensioned for acceleration stage propulsion and at least one 
air inlet designed to open after the acceleration stage is enable enough 
air to enter the combustion chamber to at least compensate the drag force 
on the missile by the high-speed ejection of the gases resulting from the 
combustion of ram air with the missile's fuel payload. 
The invention thus concerns the propulsion of tactical-type missiles having 
a flight envelope within the atmosphere such that air can be used as one 
of the combustion agents in a ramjet engine. 
Specifically, the invention concerns rocket ramjets operating with two 
successive propulsion modes, ie. a rocket engine acceleration stage using 
a fuel and an oxidizer both of which are carried aboard the missile, and a 
ramjet-based cruising stage using air scooped from the atmosphere by the 
moving missile with only the fuel carried aboard the missile. 
Problems arise in the practical construction of such rocket ramjets due to 
the fact that the optimum nozzle for acceleration or boost must have a 
much narrower throat than the throat cross section of the nozzle most 
suitable for the subsequent ramjet cruising stage. 
Solutions involving integrated ramjet and booster stages have been 
provided, using a single combustion chamber the inside wall whereof is 
common to both propulsion stages, said designs comprising a nozzle 
optimized for the acceleration stage, which is ejected during the 
transition phase between acceleration and cruising, to produce a larger 
opening in the tail end of the combustion chamber being better suited to 
ramjet cruising. 
Unfortunately, for certain flights such as for example short, or medium 
range flights, ejection of the booster nozzle is unacceptable, since it 
constitutes a danger either for the missile operator or for friendly 
troops and populations. 
In order to avoid the latter inconvenience, some prior art designs have 
done away with the booster nozzle. In such nozzle-less power plants, 
charging of the solid propellant required throughout the acceleration 
stage is arranged by providing a center channel and a divergent rear cone 
which together ensure stable combustion and axial thrust throughout the 
acceleration stage. Nevertheless, in this type of power plant combustion 
pressure considerably decreases as the missile accelerates such that the 
mean specific impulse during the acceleration stage is substantially 
downgraded compared with that which the same propellant would have 
provided had the exhaust gas been ejected through a nozzle. 
Considering the fact that the length of the combustion chamber is generally 
dimensioned by the space required for the booster propellant and that any 
increase in the mass of this solid propellant entails an equivalent 
decrease in the mass of the ramjet fuel, any downgrading of performance in 
the acceleration stage entails a reduction of cruise time and consequently 
a proportional shortening of the missile's range. 
It is the object of the present invention to remedy the above-mentioned 
disadvantages by providing for the escape of combustion gases produced 
during the acceleration stage through an actual nozzle having been 
optimized for this stage and by avoiding ejection of said nozzle whilst 
ensuring its suitability for subsequent operation in the ramjet cruising 
stage. 
Accordingly, the invention provides a propulsion system or drive as defined 
at the beginning of the foregoing, having, besides the special air inlet 
or inlets, at least one additional exhaust outlet, in the form of an 
additional nozzle located in the downstream end of the combustion chamber, 
and means of closing said additional outlet or outlets throughout the 
initial acceleration phase, such that said additional outlet or outlets 
can also contribute, together with the main nozzle, to exhausting or 
combustion gases during the ramjet cruising phase. 
In accordance with the invention, the main nozzle designed for the 
acceleration stage is kept in place during the cruising stage, but during 
said latter stage, only part of the exhaust gas is ejected through said 
main nozzle and the remainder is ejected through one or more additional 
outlets provided at the back of the combustion chamber, said outlets, like 
the previously-mentioned chamber air inlets, being opened only once all of 
the solid propellant used for acceleration has been consumed. 
In a first embodiment of the invention, said additional outlet or outlets 
are provided in the back wall of the combustion chamber in order to 
provide a substantially axial additional exhaust. 
In another embodiment, said additional outlet or outlets are provided in an 
annular, rear lateral part of the combustion chamber. 
The propulsion system according to the invention is advantageously further 
provided with an external structure surrounding the combustion chamber 
substantially to the rear end of said chamber and extending forward of the 
chamber at least partly along the length of the missile body so as to 
provide a permanent air scoop at the front of the missile and ensure ram 
air flow to the combustion chamber via said air inlets whilst maintaining 
the casing of the combustion chamber under balanced pressure during the 
ramjet cruising stage, such that during said cruising stage, only a part 
of the air taken in by the missile's air scoops is introduced into the 
combustion chamber via the air inlet holes while the remaining air flow is 
discharged behind the missile via purposely provided outlets without 
having passed through the combustion chamber but having nevertheless been 
mixed with at least that portion of the exhaust gas exiting via the 
additional exhaust outlets. 
Said external structure can be extended backwards beyond the main nozzle to 
enable mixing of the secondary air flow with all of the exhaust gas from 
the combustion chamber. 
Splitting of the air flow into a portion .alpha.Qa which is admitted by the 
air inlets and contributes to combustion and a portion (1-.alpha.)Qa which 
constitutes the cooling flow is determined as a function of the thrust 
required and of the temperature withstandable by the combustion chamber 
wall. 
The combustion chamber is centered with the help of struts inserted between 
the casing of said chamber and said previously-mentioned external 
structure. In the embodiment with additional exhaust outlets provided in 
the shell said struts can be shaped such that the cross section of the air 
passage between the external structure and the combustion chamber casing 
provides for sonic flow of both the air upstream from the additional 
exhaust outlets and the air and combustion gas mixture downstream from the 
additional exhaust outlets. 
The combustion chamber is preferrably given a thin, heat-conducting wall. 
A number of other special features may also be advantageously provided as 
follows. 
Conical, ie. two-way tapering ramps are inserted between the external 
structure and the fuel tank upstream from the combustion chamber air 
inlets to define an annular air scoop. 
During the acceleration phase, the missile's air scoops are not sealed off 
downstream, thus allowing the ram air to be discharged through the rear to 
reduce drag. 
A substantially conical fairing connects the rear neck of the combustion 
chamber to the outlet plane of the tapered out section of the main axial 
nozzle to form a central body establishing a diffuser in which the 
secondary air flow/exhaust gas mixture can expand. 
In one specific embodiment, the air inlets and the additional exhaust 
outlets are sealed off during the acceleration stage by means of plugs 
specially designed to be ejected at the end of said acceleration stage. 
Alternatively, the combustion chamber can be designed so as to be 
automatically pushed back axially by the gas pressure at the end of the 
acceleration stage over a pre-established distance to clear the chamber 
openings while arranging, on the one hand, a space between the fuel tank 
and the chamber such as to form the air inlets and, on the other hand, a 
space between the chamber and the main nozzle such as to form the 
additional exhaust outlets. Depending on the applications envisaged, the 
combustion chamber's sliding action thus serves either to open the air 
inlets, or the additional outlets, or both the inlets and outlets 
simultaneously. 
More specifically, the sliding combustion chamber is initially immobilized 
by means of a shearable tie and the geometrical configuration of said 
chamber is such that at the end of the acceleration stage, the direction 
of the axial resultant of the pressure forces on the chamber reverses and 
said resultant becomes greater than the resistance of the shearable tie.

This invention can be readily applied to the various configurations of ram 
rockets with integral boosters. In fact, the air scoops required for the 
ramjet stage of operation can be arranged in a number of different ways, 
including at the front, annularly, in the nose, as well as ventrally or 
laterally. Similarly, the fuel required for the cruising stage can be 
either liquid or solid. In the latter case it can be placed in an 
auxiliary self-pyrolyzing fuel generator or in contact with the inside 
wall of the combustion chamber. 
Referring now to FIG. 1, the front part of the missile equipped with the 
drive according to the invention is not shown. Only the fuel tank 8 and 
the combustion chamber 2 are shown, both surrounded by an external tubular 
structure 1 extending toward the back of the missile beyond the main 
exhaust nozzle 3. Said comustion chamber 2, attached to said external 
structure 1, contains the propellant 4 required for the acceleration stage 
(see the lower half-cross-sectional view of FIG. 1). As long as fuel is 
stored therein and throughout the acceleration stage, the chamber's air 
inlets 5 remain sealed by plugs. In the ramjet stage, the fuel stored in 
the auxiliary tank 8 located upstream from the combustion chamber 2 is 
injected into the combustion chamber 2 and the combustion air scooped by 
the moving missile is introduced into the combustion chamber 2 via the 
inlets 5 which have now been opened (see the upper half-cross-sectional 
view of FIG. 1). 
The embodiment illustrated in FIGS. 1 through 5 is associated with frontal 
air scoop designs with said air scoops being located either in the nose 
cone or, in annular form, downstream from the equipment which must be 
readily accessible, namely the charge and the instrumentation compartment. 
In both cases secondary air is taken in upstream from the combustion 
chamber 2 and allowed to circulate about the outside wall of combustion 
chamber 2 as a whole, thus providing chamber cooling and preheating of 
said secondary air at the same time. 
More specifically, the embodiment shown in FIG. 1 is provided with an 
annular air scoop 25 in line with the fuel tank 8. As can be seen from 
FIGS. 1 and 2, the air is scooped between an external tubular structure 1 
and a two-way tapering ramp 9 whose geometry defines the features of the 
air scoop 25. 
FIG. 2, which is a cross-sectional view of the missile in the plane of the 
scoop, shows how the ramp 9 is made. As shown in FIG. 2, this ramp 
consists of three segments 10 made for example of a composite casting. 
These three segments 10 of ramp 9 fulfill three functions: they define the 
opening 11 of the air scoop 25, which is bounded by said segments 10 and 
said external structure 1; secondly, they provide, together with the 
clearances 12 between the tank 8 and the ramp 9 and with the openings 13 
between said segments 10, a boundary layer trap; and thirdly, they 
contribute to centering and securing the external structure 1 on the fuel 
tank 8, by means of the projections 14 on said segments 10 serving as 
bearings thereof. 
The combustion chamber 2 is secured and centered within the external 
structure 1 at the rear of the missile by means of struts 15. As indicated 
by FIGS. 3 and 4, the width of said struts 15 is not constant, but changes 
between the axial position represented in cross section in FIG. 3, in 
which plane are contained the first additional exhaust outlets 6 provided 
in combustion chamber 2, and the axial position further downstream 
represented in cross section in FIG. 4 containing the most downstream 
additional exhaust outlets 6. Said struts 15 have a relatively substantial 
width in the plane of FIG. 3. Said width is calculated so that the 
available cross section 16 for passage of the secondary air flow running 
between the combustion chamber 2 and the external structure 1, and thereby 
already preheated, provides a sonic flow. Between the radial planes 
illustrated in FIGS. 3 and 4 several sets of additional exhaust outlets 6 
serve, during the cruising stage, to inject exhaust gas into the secondary 
air stream formed between external structure 1 and combustion chamber 2. 
The width of said skids 15 is reduced in the plane of FIG. 4 and is 
calculated to ensure that the cross section 16 available for the 
downstream runout of the exhaust gas is such as to also have a sonic 
mixture of combusted gas and secondary air flow. 
Downstream from the plane of FIG. 4, the exhaust gas/secondary air flow 
mixture expands in a sort of central body 18. Said central body 18 forms a 
basically conical fairing connecting the rear necking of the combustion 
chamber 2 with the tapering out portion of the main axial nozzle 3 
associated with the combustion chamber 2. 
The furthest downstream part of the external structure 1 (FIG. 5) can be 
provided with casings 19 of heat insulating material to house the control 
surfaces servomotors in the case said control surfaces must be located at 
the back of the missile. 
The lower half-cross-section of FIG. 1 shows the state of the combustion 
chamber 2 prior to the acceleration stage, with the booster propellant 4 
masking the air inlets 5 as well as the additional exhaust outlets 6, said 
outlets being further sealed by means of plugs 7. The other 
half-cross-section of FIG. 1 shows the combustion chamber 2 configured for 
the ramjet cruising mode, with the air inlets 5 and the additional exhaust 
outlets 6 for injecting the combustion gases into the secondary air stream 
now open. 
During the acceleration stage, the air inlets 5 are obstructed by plugs 7 
bearing against the inside wall of the combustion chamber 2. Said plugs 
are naturally ejected in the transition between the two propulsion modes. 
Any suitable well known blow-out plug that will be released from the 
outlet 6 due to surrounding conditions may be used. The additional outlets 
6 closing plugs 7 are ejected for example, by means of pressure when the 
solid propellant in the chamber 2 has been consumed. In addition, the 
ejection of plug 7 may be facilitated for example by means of springs at 
the end off the propellant combustion tail. 
According to one specific embodiment of the invention, the combustion 
chamber 2 is fabricated by filament winding, for example of phenolic 
silica, and the fibers wind around inserts placed about the winding core 
at the location of the air inlets 5. The additional outlets 6 can be made 
in the same way. However, if the wall of the combustion chamber 2 is 
reinforced in the area 17 where the additional outlets 6 are to be 
provided, said outlets will be more easily provided by machining. 
The operation of the missile according to the invention during ramjet 
cruising will be more readily understood with the help of FIG. 6. This 
figure schematically illustrates with a lengthwise half-cross-sectional 
view, the external structure 1, the fuel tank 8, the air inlets 5, the 
exhaust outlets 6, the main nozzle 3 and the body 18. It can be clearly 
seen from the configuration in the figure that only part (.alpha.Qa) of 
the air flow Qa drawn by air scoop 25 is introduced into the combustion 
chamber 2 via said air inlets 5, .alpha. having a value less than 1. The 
remainder (1-.alpha.)Qa of the ram air Qa makes up the secondary air which 
flows between the combustion chamber wall and the external structure 1 and 
does not take part in combustion. 
Similarly, part .beta.Qb(.beta.&lt;1) of the exhaust gas Qb exhausts axially 
via the main nozzle 3, while the remainder thereof (1-.beta.)Qb exhausts 
via the additional exhaust outlets 6 and mixes with the flow of secondary 
air (1-.alpha.)Qa. If the external structure 1 is extended downstream from 
the nozzle 3 far enough, it is possible to achieve mixing of all of the 
exhaust gas with the secondary air, the flow .beta.Qb of gas issuing from 
the main nozzle being thoroughly mixed with the flow 
{(1+.alpha.)Qa+(1-.beta.)Qb} of gas running out on the outside of said 
nozzle via an annular opening 26. 
The dual-flow mode of operation according to the invention, ie. involving a 
secondary air flow (1-.alpha.)Qa running between the combustion chamber 2 
and the external structure 1, which then mixes with at least a part 
(1-.beta.)Qb of the exhaust gas, provides a number of advantages over 
conventional single-flow operation. 
The primary advantage stems from the reduction of the velocity of the gases 
in chamber 2. This not only limits pressure drop and the convection 
coefficient at the wall, but, more importantly, promotes thorough 
combustion due to a longer stay in the chamber, the physical length of 
said combustion chamber 2, being defined in practice by the volume of 
booster propellant, being in fact roughly the same as for a single-flow 
ramjet design. Thus, good flame stabilization as well as greater 
efficiency of combustion result from the low velocity in the chamber. 
Another advantage is provided by the possibility of efficiently cooling the 
shell of chamber 2, for the coefficient of external convection to the 
non-combusted air is significantly greater than the coefficient of 
internal convection with the hot gases. Since the calories passing through 
the wall of the combustion chamber contribute to heating and accelerating 
the secondary air flow, nothing prevents the walls of the chamber from 
being made thin and heat-conducting, providing they are made from an 
oxidation-resistant material. This is advantageous to the extent that any 
reduction in thickness of the chamber walls enhances the missile's overall 
performance by freeing more space for booster propellant and/or ramjet 
fuel. 
Moreover, in a dual-flow type ramjet engine, the combustion chamber shell 
is under balanced pressure. The internal pressure differential in the 
cruising mode becomes nil upstream and remains small downstream. 
Furthermore, it deserves to be emphasized that during the acceleration 
stage with a conventional integral booster ramjet the combustion chamber 
air inlets are necessarily closed and consequently the missile's ram air 
augments drag, by .DELTA.F, for no useful purpose. The increase in drag on 
a dual-flow ramjet during the acceleration stage on the other hand is 
limited to .alpha..DELTA.F, .alpha. being a coefficient less than 1 
standing for the proportion of air injected through the inlets 5 during 
the cruising stage. Compared with a conventional ramjet engine therefore, 
this invention enables recovery of an amount of thrust (1-.alpha.).DELTA.F 
which is added to the thrust of the solid propellant drive during the 
missile's acceleration stage. 
To summarize, a propulsion system or drive according to the present 
invention makes it possible, as with nozzle-less booster ramjets, not to 
jettison any heavy, compact body, while at the same time enabling a 
performance comparable to that of ejectable nozzle ram rockets. 
A second embodiment of the invention is depicted in FIGS. 7 and 8 according 
to which opening of the air inlets 105 and of the additional exhaust 
outlets 106 for combustion gas is done automatically thanks to the sliding 
action imparted to the frame of the combustion chamber. 
The drive system illustrated in FIGS. 7 and 8 comprises a nozzle 103 
specifically designed for acceleration stage operation and attached to the 
missile's external structure 101, said latter structure withstanding both 
internal pressure and flight stresses. A combustion chamber 102, initially 
containing the booster propellant, is positioned with respect to said 
external structure 101 by means of struts 109 establishing a free space 
between the external structure 101 and the wall of chamber 102. Chamber 
102 can slide backwards when during the booster combustion tail the 
resultant of the pressure forces on the head of the chamber becomes 
greater than the resistance of a shearable tie. A tank 108 located 
upstream from the chamber 102 contains the solid self-combustible fuel for 
the ramjet cruising stage. Said tank or store 108 ignites from proximity 
alone and contributes to the impulsing of the booster. In ramjet-type 
operation, the gases are fired spontaneously at contact with the oxygen. 
No special ignitor is therefore required for the cruising stage. 
The booster operates in the same way as in the first embodiment described 
hereinabove. An ignitor placed in said nozzle 103 starts combustion of the 
solid propellant placed in the combustion chamber 102. 
As the missile accelerates, the ram air pressure increases and when the 
direction of the resultant of the booster's internal pressure and of the 
external air pressure on the body of the combustion chamber 102 reverses, 
during the combustion tail, said resultant eventually becoming greater 
than the strength of a shearable tie serving to initially immobilize said 
combustion chamber 102, said justmentioned chamber slides back to open 
annular spaces 105 and 106 at the front and the back respectively, so 
respectively allowing air entry to the chamber 102 and exhaust of the 
spent mixture. The transition between the two propulsion modes (rocket to 
ramjet) thus differs from that of the first-mentioned embodiment. During 
the cruising stage, the very rich combustion products generated in the 
chamber 102 are partly mixed with the secondary air flow in the free space 
126 forming an open annular chamber between the external structure 101 and 
the main nozzle 103 of combustion chamber 102, downstream from the annular 
space 106 forming an additional nozzle. The servomotors driving the 
control surfaces 122 can be housed in add-on casings 120 of the external 
structure 101, in the space forming a mixing chamber. 
The self-pyrolyzing solid propellant tank 108 communicates with the 
combustion chamber 102 via an axially located outer opening 121 provided 
in the rear end of said tank 108. The nose cone 123 located ahead of tank 
108 includes a point 124 which, together with the front end of the 
external structure 101, establishes an annular front air scoop 125. This 
configuration has the advantage of ensuring a good aerodynamic performance 
and preserving the missile's symmetry of revolution. The air inlet 105 
into the combustion chamber 102 which is formed during the cruising stage 
is annular. The cross-sectional area of air inlet 105 is roughly half that 
of the combustion chamber 102. The widening of cross-sectional area 
accordingly formed has the effect of recycling air to stabilize 
combustion. 
The main nozzle 103, the rear end fitting 110 of the combustion chamber 102 
which bears against the nozzle 103 during the acceleration stage and the 
respective end fittings 111 and 112 of the front of the combustion chamber 
102 and the rear of the tank 108 which cooperate to seal the combustion 
chamber 102 during the acceleration stage, are fabricated with short 
fibers, such as for example, molded short silica fibers in a phenolic 
resin matrix, thus avoiding any problems or differential expansion as 
would be likely to occur with metal parts. 
In the embodiment of FIGS. 7 and 8, the rear end fitting 110 of the 
combustion chamber has a cylindrical section 115 of revolution about the 
axis of the missile. In starting position, said section 115 surrounds a 
matching cylindrical section 116 of the nozzle 103 and applies to said 
section 116 in an airtight manner, a seal 113 being furthermore provided 
between the coaxial cylindrical sections 115 and 116 such that said two 
sections form a means 107 of sealing the additional annular outlet 106. In 
the transition from the acceleration stage to the cruising stage, said 
outside cylindrical section 115 simply slides over said inside cylindrical 
section 116, thus clearing an additional annular exhaust outlet 106. At 
the front of combustion chamber 102, said end fitting 111 is provided with 
a cylindrical flange of revolution 117, which in starting position engages 
in a mating slot 118 of the rear end fitting 112 of tank 108, a seal 114 
being arranged between the flange 117 of end fitting 111 and the slot 118 
of end fitting 112. The slot 118 and flange 117 have opposite faces 
parallel to the missile centerline so as to enable smooth sliding in the 
transition from the acceleration to the cruising mode. 
For example, the ratio of the cross sectional area of the air scoop 125 to 
that of the reference cross section consisting of the cross sectional area 
of the combustion chamber 102 can be in the neighborhood of 0.4, whereas 
the ratio of the cross section of the air inlet 105 to that of said 
combustion chamber 102 is about 0.5. The ratio of the throat cross section 
of main nozzle 103 to same said combustion chamber cross section can be 
about 0.06 and the cross section of the additional exhaust outlet or 
outlets 106 can be about twice that of said nozzle 103 throat, whereas the 
cross-sectional area for the flow of secondary air between the external 
structure 101 and the chamber 102 can be on the order of 1.4 times the 
cross-sectional area of the main nozzle 103.