Search Images Maps Play YouTube News Gmail Drive More »
Sign in
Screen reader users: click this link for accessible mode. Accessible mode has the same essential features but works better with your reader.


  1. Advanced Patent Search
Publication numberUS3486699 A
Publication typeGrant
Publication dateDec 30, 1969
Filing dateNov 21, 1966
Priority dateNov 22, 1965
Also published asDE1300355B
Publication numberUS 3486699 A, US 3486699A, US-A-3486699, US3486699 A, US3486699A
InventorsLouis Francois Jumelle
Original AssigneeSnecma
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Adjustable exhaust unit for turbojet propulsion engines
US 3486699 A
Abstract  available in
Previous page
Next page
Claims  available in
Description  (OCR text may contain errors)

Dec. 30, 1969 I... F. JUMELLE 5,

ADJUSTABLE EXHAUST UNIT FOR TURBOJET PROPULSION ENGINES Filed Nov. 21, 1966 2 Sheets-Sheet 1 Dec. 30, 1969 3,486,699


United States Patent Int. 01. B 64c 15/04 U.S. Cl. 239265.25 4 Claims ABSTRACT OF THE DISCLOSURE An adjustable exhaust unit for a jet propulsion engine, comprising at least two propulsion nozzles supplied from the tail pipe of the engine, namely: a main nozzle the throat of which is fixed and an auxiliary nozzle the throat of which is adjustable.

An almost perennial problem facing jet propulsion unit designers is the determination of a jet pipe nozzle which at one and the same time is adjustable to ensure a suitable constriction for the propulsion unit (in order to obtain optimum conditions in the thermodynamic cycle) 'and possesses good performance figures in various flight conditions and at the various operating speeds that are envisaged, without so much as possible raising too great technological difliculties in bringing this about. This problem assumes a particularly difficult character when the question at issue is the design of a jet propulsion unit with high thrust, hence having considerable dimensions (which makes the technological problems more acute), this unit being intended for a supersonic commercial aircraft (which extends the range of flight conditions over which the nozzle is to possess first rate efficiency, a basic condition in the field of economics for a supersonic civil transport aircraft).

Within this scope, the dimensions of the nozzles, the importance of varying the cross-section as an absolute value between the minimum and the maximum section required whatever the circumstances, and particularly if an afte'rburner is envisaged for certain phases of flight, the need to provide configuration suiting subsonic and also supersonic speeds, and the requirement to employ devices fulfilling a role as silencers, thrust reversers, etc. frequently lead either to deadlocks or to those conceptions that present serious disadvantages from the point of view of weight, bulkiness, drag, power output employed for control purposes and the cost of embodiments which are at the same time of large dimensions and highly complex.

In the present invention it is proposed to bring to bear on the problem under consideration a relatively simple and satisfactory solution. It relates to an adjustable exhaust outlet device for jet propulsion units the essential character of which resides in its being subdivided into at least two parallel and branching nozzles; on the one hand, a non-adjustable main nozzle and hence one which is very simple and light (or, in one modification, slightly adjustable but relatively simple) and having a section approximating to the minimum envisaged for the various operating speeds and, on the other hand, an adjustable auxiliary nozzle the section of which is variable between a zero or very small value and a maximum value corresponding to the additional amount of section or area necessary to create, along with the main nozzle, the overall section necessary to the operating speed at which the constrictive section should be at its largest.

In one preferred embodiment of the present invention, the main nozzle will possess a section equal to the minimum required by the jet engine whatever the flight conditions (or even a slightly lesser section so as to allow the jet engine to operate always at a speed subject to the adjustment of the nozzle), the adjustable auxiliary nozzle between its closed and its open position supplying the additional amount of section that makes it possible to cover (by adding together the main non-adustable section and the adjustable section of the auxiliary nozzle) all the sections required by the jet engine in every flight condition.

It should be noted that, as the two nozzles are not arranged coaxially with respect to one another, the bringing into operation of the auxiliary jet pipe gives rise to torque in respect of pitch or of yaw which it is necessary to take into account. According to one embodiment of the present invention, this disadvantage is disposed of by means of an arrangement of two or more adjustable and parallel auxiliary nozzles branching off and distributed symmetrically around the thrust axis of the jet engine, and the adjustment control of which is synchronized so as to provide equal thrusts. It will be appreciated, however, that such a symmetrical arrangement can be diverged from by so disposing the adjustable auxiliary nozzles that the torque moments generated by them in relating to the thrust axis virtually cancel each other out, as in the preceding case.

It is likewise possible, on the other hand, to obtain advantage from the presence of the auxiliary nozzle by orientating it judiciously in order to correct, at least partially, torque moments which might disturb the longitudin'al stability of the aircraft (especially of a supersonic aircraft), torque moments due, for instance, to the shifting of the focal point of the wing between subsonic and supersonic flight conditions.

It is evident that one of the notable advantages achieved by the present invention refers to the relatively modest dimensions of the auxiliary nozzle in comparison with a single suitable nozzle of classic design, since the value of the section of the non-adjustable main nozzle must be deducted. It follows that the adjustment device for such an auxiliary nozzle can be effected to give but little bulk and little weight. and that its control will only necessitate servo-motors with a modest power output.

The following description with reference to the accompanying drawings and given by way of non-limitative example Will bring out how the invention .may be put into effect, features emerging both from the drawings and the text natura ly forming part of the said invention.

FIGURE 1 is a diagrammatic view end-on from the rear of an exhaust outlet device in accordance with the invention especially suitable for a subsonic aircraft.

FIGURE 2 is a fractional longitudinal section of the same through the line IIII in FIGURE 1.

FIGURE 3 is a similar view of a modified embodiment.

FIGURES 4 and 5 show, viewed from the rear as in the case of FIGURE 1, two other modifications which would appear preferable in the case of a plurality of engines grouped by nacelles each with at least two engines.

FIGURE 6 is a diagrammatic section showing an example of the adjustment means of an auxiliary nozzle especially suitable for a supersonic aircraft.

FIGURE 7 is a modification in which the primary main nozzle, instead of being non-adjustable, is partially adjustable.

FIGURES l and 2 illustrate the basic idea of the present invention. The exhaust unit of the jet propulsion engine, supplied with hot gases, comprises a tail pipe 1 which is subdivided into two nozzles which have parallel axes but which in certain cases may not be parallel, as mentioned above: a non-adjustable main nozzle 2, that is,

one with a constant section, in the extension of the tail pipe 1, and an auxiliary nozzle 3 with an adjustable section and connected to the tail pipe by a branching duct 4, the whole being housed in a fairing 5.

The branching duct may terminate upstream at a lateral portion of the tail pipe 1, as is shown in FIGURE 2, and it may there abstract a portion of the propulsive gases or, alternatively (and seemingly more advantageous), an annular collecting device 6, equipped with a grill of guide vanes 7, may be interposed in order to effect a more symmetrical abstraction, as is shown in FIGURE 3.

The non-adjustable main nozzle 2 has a section equal to the minimum section specified for every flight condition (minus a certain margin). The complementary amount of section required for every one of the flight conditions envisaged is provided by the auxiliary nozzle 3 which, in the open position, should supply the additional amount of section necessary to obtain the maximum total section desired (plus a certain margin). The auxiliary nozzle 3 will therefore be adjustable between a zero or very small section and a fixed maximum open section. By way of example and to guide ideas on the sub ject, it may be considered that in many cases the minimum total section required is about 90% of the maximum total section required and that, consequently, the nonadjustable main nozzle 2 will have a section corresponding to the said 90% whereas the adjustable auxiliary nozzle 3 will in its open position supply the of complementary section. The adjustable auxiliary nozzle 3 may in other respects be of a conventional type.

In this manner there is produced an exhaust outlet assembly in which the important basic flow representing the greater part of the total flow is delivered via a jet pipe lacking means of adjustment and in "which the necessary adjustment takes effect on a .modest complementary flow representing but a small fraction of the total flow. This adjustment may therefore be effected in the most desirable conditions as regards delicate accuracy, ease of operation, lightness, accessibility to secondary air and, in view of the fact that complications constituted by the adjustment means apply only to an assembly of small dimensions, the manufacturing costs of this adjustment means will be considerably lowered.

In the embodiments described with reference to FIG- URES 1 to 3, the auxiliary nozzle 3, located above the non-adjustable main nozzle 2, will produce a nose dipping moment which is a function of the thrust exerted by this nozzle and of the leverage of said thrust in relation to the axis of the main nozzle 2. To avoid having to make arrangements aimed at neutralizing this nose dipping moment, it is suflicient to provide a symmetrical arrangement of the nozzles, as shown in FIGURES 4 and 5, one particularly recommended in the case of a plurality of jet engines grouped by nacelles each with two engines.

In FIGURE 4 there are two juxtaposed non-adjustable main nozzles 2a, 2b and, in the transverse space formed between these nozzles, two adjustable auxiliary nozzles 3a, 3b are accommodated, spaced symmetrically in relation to the longitudinal plane passing through the axes of the non-adjustable nozzles 2a, 2b.

The positioning of these two auxiliary nozzles, one at the top and the other at the bottom, may necessitate localized swellings in the fairing 5, which may in certain cases cause a slight increase in drag.

This drawback may be avoided by giving the adjustable auxiliary nozzles a prismatic form, for example a triangular one, as is shown at 3a, 3'!) in FIGURE 5.

In FIGURE 6 there is a diagrammatic illustration of one application of the present invention to a nozzle for supersonic aircraft. The main nozzle 2 is non-adjustable and the main gas flow is organized in accordance with a conventional type. For example, in the embodiment illustrated, the primary jet is recovered in a divergent recovery duct 8 the throat of which is ventilated by secondary air 10 (coming from a boundary-layer trap at the intake cowling, for example), said duct ending with adjustable shutters 11 which make it possible to suppress or to limit over-expansion during subsonic flight.

The same principle applies to the auxiliary nozzle. In the upper part of FIGURE 6 there is shown an adjustable auxiliary nozzle 12 (reduced to the plane of the figure so as to facilitate understanding), the adjustment being obtained in a conventional manner, for instance by means of multiple shutters actuated by jacks (not here shown). The auxiliary nozzle 12 also cooperates with a divergent recovery duct 13 including a ventilated throat 14 into which a small portion of secondary air 10' can pass, and terminal shutters 15 making it possible to avoid or to limit over-expansion during subsonic flight.

Other arrangements of the main nozzle may also be employed. In the lower part of FIGURE 6 there is shown a method of adjusting the section of a nozzle having a two dimensional convergent-divergent configuration and likewise reduced to the plane of this drawing in section. A portion 16 of the fairing carries interiorly a swelling 17 suitably configured to constitute one of the walls of a convergent-divergent nozzle, the other wall being formed by a central body 18. The portion 17 is able to slide longitudinally whereas the central body 18 is fixed. It will be observed that the effective section of the auxiliary nozzle can vary from a zero or very small minimum of the portion 17 (shown in chain-dotted lines) up to a considerable value for a position in which this portion has been shifted upstream (shown in a continuous line). Naturally a similar result would be obtained by inter-changing the fixed and movable portions, the swelling 17 being fixed and the central body 18 being produced in the form of a sliding tongue.

The advantages of the present invention have been set out above and apply as well to jet engines having a convergent nozzle operating dry as to jet engines having a convergent-divergent nozzle for supersonic aircraft.

In fact, even in the first-named case, it is certain than an adjustable nozzle of appreciable dimensions levies a cost in weight and in bulki'ness by reason of the internal pressures against relatively large shutters and of the jacks required to control the latter. On the other hand, within the scope of the present invention, a non-adjustable nozzle 2 backed by a nozzle 3 which allows the passage of between none and 10% of the flow from the engine, would provide a saving in weight and in bulkiness in spite of the extra ducting 4. On the other hand, too, from the point of view of bulk, a nozzle of such small dimensions may easily be accommodated in the oval formed by the nacelle or the fuselage.

Apart from this general advantage, in the case of a convergent-divergent nozzle with a ventilated throat for a supersonic aircraft, and especially for such nozzles, there are other additional advantages which derive from the employment of a non-adjustable main nozzle delivering a substantial part of the flow of gases.

(a) Accessibility to the secondary airflow, often poor because of the multiplicity of jacks, pipes and nozzleshutter actuating rods, is greatly improved by reason of the fact that, instead of having to make a large-dimension nozzle vary, the jacks, pipes and linkage rods are only applied to an auxiliary nozzle on a much reduced scale, hence are themselves of reduced size and so of reduced obstructiveness. For this reason the degradation in energy of the secondary air circulating around the primary hot jet is greatly reduced, hence a considerable gain in efficiency for the exhaust assembly.

(b) One of the difliculties in solutions as at present employed is that the adjustable primary nozzle has a variable section whereas the throat of the divergent recovery duct is non-adjustable which imposes a compromise that is likewise harmful to the efliciency of the main nozzle by reason of the fact that, in order to prevent shocks on the recovery duct when the nozzle is fairly open, one is reduced to give the recovery duct too large a cross-sectional area, which causes it to deviate from the optimal crosssectional area. With the solution proposed, the major portion of the gas flow passes by way of a non-adjustable nozzle, which allows the achieving of the optimum match for the dimension and the position of the throat of the recovery duct.

(c) If a silencer is to be incorporated (this is not shown in the figures), variations in the section of the jet are also generally a cause of complexity. In the case in which the present invention is utilized, the silencer is simpler since it has to be matched only to a jet of a nonadjustable section; the small-dimension adjustable auxiliary nozzle can be left without a silencer without any serious disadvantage.

The operating balance-sheet in respect of the main nozzle therefore shows a considerable gain in respect of weight, a very noticeable gain in accessibility to secondary air, hence in performance figures, a better match in the recovery throat and improved effectiveness in the silencer. Further, the complex portion being of small dimensions, the combination of simple main nozzle and complicated but small auxiliary nozzle has a lower manufacturing cost than that of a complicated large-dimension nozzle.

Moreover, the complementary nozzle, by reason of its modest dimensions, can be of the best possible type without so much importance being attached to problems of weight and of complexity.

It must be especially emphasized that nozzles with multiple adaptation means, when intended for use with the large engines required in supersonic aircraft, are of such dimensions that, on the one hand, their shapes must of necessity be curved and preferably of revolution shape (flat panels cannot, due to temperature, withstand the pressures involved with such engines, except accompanied by a prohibitive increase in weight), and on the other hand compromises have to be accepted in respect of a solution in principle, for the selfsame reasons of complexity and weight. On the contrary, in small nozzles, the weight and complexity required to obtain maximum suitability in all flight conditions (for example, by means of a nozzle with a two-dimensional tongue) are admissible because the main nozzle is, on the other hand, simplified and made lighter and because complexity and extra weight apply only to a minor proportion of the whole.

Further, it should be noted that such auxiliary nozzles, once specified and fabricated, are applicable to different engines, or to diiferent developmental stages of the same engine, the main nozzle alone being changed, this latter, however, being of great simplicity since it is nonadjustable, and so this solution provides great flexibility in engine development.

Moreover, it should be observed that a great number of engines for supersonic aircraft have characteristics of a kind that the outlet section of the exhaust unit increases continuously with the Mach number, generally in a minor way (and, precisely, of the order of between takeoff and supersonic cruise). This means that take-off will take place with the auxiliary nozzle practically closed (except for the slight aperture to retain a margin for adjustment purposes) and that the auxiliary nozzle will open gradually as the Mach number increases until it is practically at maximum aperture in supersonic cruise.

This observation may be utilized in two ways: on the one hand, the suitability of the auxiliary nozzle may be brought about by choosing, by way of compromise, an adaptation closely approximating to that which corresponds to the maximum expansion rate corresponding to supersonic cruise, since the air flow passing along the nozzle for take-off or subsonic cruise is zero or very small. It is therefore easier to accept relatively large losses on a very minor fraction of a flow which is itself small (about 10% of the flow from the jet engine).

On the other hand, use may likewise be made of the difference in possible total section between the section required at take-01f and the total section available when the adjustable auxiliary nozzle is completely open, through burning, particularly in the main nozzle (this would appear a preferable solution), a certain amount of fuel in an afterburner. The appropriate constriction of the engine, in order to achieve the optimal thermodynamic cycle, will thus be maintained by thermal constriction of the main nozzle, with an additional amount of thrust provided roughly by the ratio of the total sections, with auxiliary nozzle open, to the main nozzle section, i.e., in the example selected, an added amount of thrust of the order of 10%.

After the aircraft has taken off under these conditions while using moderate re-heat and with the auxiliary nozzle open, and after the aircraft has reached a safe altitude and speed, the extinction of the afterburner can be effected, the adjustable auxiliary nozzle then closing to approximately zero section. It should liksewise be noted in this case that, on the one hand, the use of a silencer made to suit the main nozzle is not affected by varying the cross-sectional throat area of the exhaust unit and that, on the other hand, in the preferred solution of re-heat in the main nozzle only, moving parts (always delicate) are not subjected to an increase in temperature by reason of the afterburner.

The same basic idea of dividing the flow of gases into a main part and an auxiliary part can be applied to main nozzles the cross-sectional throat area of which is adjustable in certain operating conditions of the engine and non-adjustable in other operating conditions. It has, in fact, been proposedthat the main primary nozzle should not only be convergent but should include a divergent portion. The aim of this proposal is to bring the pressure of primary gases down to a level of expansion sufiicient for them to be mixed with secondary air, after use has been made of the additional amount of thrust that corresponds to a controlled expansion in the divergent portion, an expansion taking place from the pressure prevailing at the throat to the pressue of the secondary air. The pressure of the latter is, in fact, generally distinctly lower than the total pressure of the primary jet but, in spite of all that, it remains higher than the outside pressure, so that the combination of the two jets, the primary and the secondary, reduced to the same pressure as stated above, is afterwards expanded in a divergent recovery duct of conventional type.

The results of wind-tunnel tests confirm the above reasoning, and calculations show that, in certain conditions, improved efficiency in the exhaust outlet assembly can be obtained by the use of a primary nozzle compris: ing a continuous throat and slight divergence behind the throat.

Nevertheless, the complexity of the sealing devices for an adjustable nozzle with a convergent-divergent contour and the power required to adjust shutters that perforce are longer in practice prevent the application of such a solution, if only in the light of the increase in weight. The employment of a semi-adjustable nozzle, of the type shown in the upper part of FIGURE 7, has therefore been envisaged. This nozzle comprises a nonadjustable convergent portion 19, at the downstream end of which movable shutters 20 are hinged, these being able to assume positions ranging'from a divergent position 20" to a convergent position20' in which they practically form an extension of the non-adjustable portion 19. As in the case of FIGURE 6, the expulsion of gases ensues in a conventional manner, for example in a divergent recovery duct having a throat 22 which is supplied with secondary air 10, and a divergent portion which comprises a non-adjustable portion 21 and adjustable terminal shutters 23.

At take-ofi speeds and in all subsonic conditions, the primary nozzle then operates as an adjustable convergent unit of classic type between a cylindrical position and the position 20', this thus effecting a suitable constriction of the jet and maintaining the optimal thermodynamic cycle for all subsonic conditions.

On the other hand, in supersonic cruise conditions, the primary nozzle is open at position 20", effecting a preliminary expansion of the primary flow down to a suitable pressure approximating to that of the secondary air 10 Theoretically, the terminal diameter of the non-adjustable convergent portion 19 should be determined to provide in supersonic cruise a throat section 24 corresponding to the theoretical construction section required in supersonic cruise. Nevertheless, it will be observed without diflicult that in supersonic cruise the primary throat is non-adjustable, which constitutes a grave drawback by reason of the fact that the constrictive section required in supersonic cruise may vary either from one engine to another of the same type, as a consequence of manu facturing tolerances, or as a function or transient ambient conditions that are met with. For this reason, in the solution proposed above, the drawback of the lack of adaptation in the adjustment section 24 in supersonic cruise balances, and in the majority of cases even outweighs, the advantage of preliminary expansion behind the non-adjustable nozzle. This lack of adaptation in fact makes it necessary to dc-rate the basic engine, both as regards its rotational speed and its turbine entry temperature, in order that, even when taking into account engine scatterings and scatterings appertaining to transient conditions, values in respect of the rotational speed and of the turbine entry temperature never exceed the limiting values acceptable for mechanical behaviour or endurance.

The applicant has verified that the cost of this derating in engine thrust and in the expenditure of fuel per kilometer flown was much higher than the slight gain obtained by preliminary expansion.

The use of the principle of the present invention does away with the drawbacks of this solution involving a semi-adjustable nozzle (adjustable in subsonic flight, nonadjustable in supersonic).

Referring again to FIGURE 7, it is sufficient, in accordance with the present invention, to add an adjustable auxiliary nozzle of small dimensions to a main nozzle with preliminary expansion, hence including a divergent portion, so as to benefit from the gain due to preliminary expansion, without in any way giving up adjustment during supersonic flight. The auxiliary nozzle might, for example, be of the same type as that shown in the upper part of FIGURE 6, and might include a primary portion 12 adjustable from a zero section to a specific maximum section said primary portion cooperating with a ventilated throat 14 appertaining to a recovery duct which is supplied with a small amount of secondary air 10', expansion taking place in a divergent portion 13 of said recovery duct with movable shutters 15 at the downstream end thereof.

The operation of the auxiliary nozzle is unchanged in relation to what has been stated previously; the operation of the main nozzle is indicated below.

At take-off and in subsonic flight, engine constriction is effected by the shutters 20, the primary auxiliary nozzle 12 being in the closed position and the shutters adopting a position lying between the cylindrical position and the position 20' and between these two positions ensuring the most efficient adjustment of the engine load, and hence the optimal thermodynamic cycle. In supersonic cruise, the shutters 20 open into the divergent position 20 the auxiliary nozzle 12, as appears preferable, having the intermediate section between zero and its maximum section. Through this fact, differences in the total constriction section as required for the engine in supersonic cruise and due either to scatterings from the engine or scatterings from ambient conditions may be taken into account by adjusting the auxiliary nozzle 12 which is adjustable in either direction. From this, the advantage is gained of efficiency in the pro-divergent portion 20", without surrendering the requisite adjustment of the overall cross-sectional throat area of the exhaust unit, said adjustment being effected by the auxiliary nozzle 12 in relation to the theoretical section of an average engine flying in standard atmospheric conditions for which the required constriction would be represented by the sum of the non-adjustable throat section 24 and of one half of the maximum section of the adjustable nozzle 12. The section 24 will preferably have been designed to satisfy the relationship cross-sectional area of throat section 24 half of cross-sectional area of section 12 the required cross-sectional area of the constriction for an average engine flying at cruise speed in standard conditions.

It should be noted in this case that the margin of variation required for the adjustable auxiliary nozzle is greatly decreased in relation to the case in which the main nozzle is of the nonadjustable type, as shown in FIG- URE 6. In fact, the variations in the overall cross-sectional throat area to be effected by the adjustable auxiliary nozzle, in the case of FIGURE 7, only corresponds to scatterings from the engines and to scatterings from ambient conditions in supersonic cruise, the margin of adjustment to take account of variations in the operating conditions of the engine between supersonic cruise and subsonic cruise or take-off being effected by the main nozzle 20 operating as an adjustable convergent unit. In the case of FIGURE 6, the adjustment margin to be covered by the adjustable auxiliary nozzle would, moreover, correspond to engine operating conditions from zero speed (take-off) to supersonic flight.

In particular, it has been found that an adjustment margin in the auxiliary nozzle of 5%, to take account of scatterings from the engines or from ambient conditions, would be broadly sufficient (by reason of the fact that the main nozzle 20 makes it possible in subsonic conditions to effect the requisite constriction adjustment for the engines from its position 20' up to the cylindrical position).

What has been stated previously with reference to non-adjustable main nozzles is obviously applicable to the semi-adjustable nozzle shown in FIGURE 7; the employment of an afterburner, especially, with the shutters 20 located in the cylindrical position (or very slightly convergent), is applicable in this case, the adjustable auxiliary nozzle 12 in this event being even left closed, but equally being capable of being placed in a position of full aperture (a preferred solution, since it allows the additional amount of thrust at take-off due to the afterburner inside the main nozzle to be increased by an extra 5% It will be appreciated that the embodiments described are only examples and that they may be modified, more especially by substituting technical equivalents, without however thereby going beyond the scope of the invention.

What I claim is:

1. In a jet propulsion unit having a jet pipe supplied with a stream of motive gas under pressure, a propelling nozzle system having an adjustable effective area for the exhaust of said gas into the atmosphere, said effective gas exhaust area being continuously variable between a minimum cross-section value and a maximum cross-section value, said propelling nozzle system comprising: two propelling nozzles connected in parallel-flow relationship directly and substantially unobstructedly to said jet pipe to be supplied therefrom through respective motive gas tapping intakes of fixed cross-section, namely a primary propelling nozzle having a motive gas tapping intake of relatively large constant cross-section for deriving a major fraction of said motive gas stream from said jet pipe and having further a substantially constant effective gas exhaust area which is a major fraction of said maximum cross-section value and which is close to said minimum cross-section value, and

a secondary propelling nozzle having a motive gas tapping intake of relatively small constant cross-section for deriving a minor fraction of said motive gas stream from said jet pipe and having further movable sidewall means peripherally bounding said minor fraction stream of motive gas in the process of exhausting into the atmosphere, said movable sidewall means defining therefor both a gas exhaust outline of variable geometrical shape and an effective 'gas exhaust area of variable magnitude which is continuously adjustable between a substantially zerocross-section wherein said sidewall means is in a substantially closed position of high degree of convergence and a maximum cross-section wherein said sidewall means is in a fully open position of low degree of convergence, said maximum cross-section being substantially equal to the difference between said maximum value and said minimum value.

2. Nozzle system as claimed in claim 1, wherein said primary propelling nozzle has an effective area short of said minimum value, and said secondary nozzle, when in said substantially closed position, has a reduced effective area which adds up to that of said primary propelling nozzle to make for said minimum value.

3. Nozzle system as claimed in claim 1, further comprising a recovery duct spaced downstream of said primary propelling nozzle to collect the jet issuing therefrom and induce ambient fluid from around said primary propelling nozzle, said duct having a minimum crosssectional area greater than said substantially constant effective area of said primary propelling nozzle.

4. Nozzle system as claimed in claim 3, wherein said primary propelling nozzle has an efiective area short of said minimum value, and said secondary nozzle, when in said substantially closed position, has a reduced efiective area which adds up to that of said primary propelling nozzle to make for said minimum value.

References Cited UNITED STATES PATENTS OTHER REFERENCES Flight: Jan. 13, 1961, pp. 42-43, Solid-Propellant Motors.

EVERETT W. KIRBY, Primary Examiner U.S. CL. X.R.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2692800 *Oct 8, 1951Oct 26, 1954Gen ElectricNozzle flow control
US2825205 *Jul 12, 1954Mar 4, 1958SnecmaCombustion devices especially suitable for gas turbine engines and propulsion units for aircraft
US2928238 *Jun 8, 1953Mar 15, 1960Lockheed Aircraft CorpJet deflector and orifice control
US3167912 *Jan 4, 1960Feb 2, 1965United Aircraft CorpThrust control for solid rocket
US3245620 *Dec 13, 1961Apr 12, 1966Gen Motors CorpMissile steering control
US3304723 *Aug 10, 1965Feb 21, 1967Gaura Robert EProportional thrust valve for space vehicle control system
US3349563 *Nov 15, 1965Oct 31, 1967Thiokol Chemical CorpThrust control apparatus for plastic propellant rocket motor
USRE26177 *Jul 28, 1961Mar 21, 1967 Automatic steering of space craft
GB999581A * Title not available
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US6845606 *Jun 13, 2002Jan 25, 2005Snecma MoteursVariable cycle propulsion system with gas tapping for a supersonic airplane, and a method of operation
US7162859Nov 23, 2004Jan 16, 2007Snecma MoteursVariable cycle propulsion system with gas tapping for a supersonic airplane, and a method of operation
U.S. Classification239/265.25, 239/265.27, 60/263, 239/265.17, 239/265.19
International ClassificationF02K1/40, F02K1/52, F02K1/06, F02K1/15
Cooperative ClassificationY02T50/671, F02K1/40, F02K1/15, F02K1/06, F02K1/52
European ClassificationF02K1/40, F02K1/06, F02K1/52, F02K1/15