System for controlling speed transition and thrust vectorisation in a multiple-shaped nozzle by secondary injection

Abstract

A mixing tube with multiple shapes is provided, allowing additional injection of gas in order to keep the flow detached from the second shape in an ascent phase and to bring about, in a descent phase, a controlled detachment as a result of the change of slope between the two shapes. A propulsion nozzle for an engine of a spacecraft or aircraft is provided including such a mixing tube and a method for controlling the speed transition of the propulsion gases in such a nozzle in accordance with the altitude. Also, a method is provided for vectorising the thrust in such a nozzle by radial and asymmetrical injection of gas and a control method which prevents re-attachment of the jet to the second shape of such a propulsion nozzle for an engine of a spacecraft when it is in the take-off or landing phase.

Claims

1. A divergent section comprising multiple curved contours for a propulsion nozzle of an engine for aircraft or spacecraft propelled by chemical reaction, the propulsion nozzle being a nozzle having rotational symmetry about an axis, the engine comprising a reaction chamber producing propulsion gases intended to be ejected outside the propulsion nozzle in a mainstream, the divergent section comprising a nozzle throat, the divergent section further comprising: a first portion comprising a first wall connecting the nozzle throat to a connecting zone, said first wall having a first surface of revolution about the axis of the propulsion nozzle, so that said first wall has a first curved profile in an axial plane, said first curved profile being suitable for generating a flow of the propulsion gases adapted to operation at low altitudes; a second portion having a total length L, which is connected to said first portion at the connecting zone defining a curved contour discontinuity, said second portion comprising a second wall extending from the connecting zone to an outlet to exhaust the flow of the propulsion gases to an ambient environment, said second wall having a second surface of revolution about the axis of the nozzle, so that said second wall has a second curved profile in an axial plane, the connecting zone defining a change in slope between the first curved profile and the second curved profile, said second curved profile being suitable for generating a flow of the propulsion gases adapted to operation at high altitudes, wherein said connecting zone extends between said first portion and said second portion such that said curved contour discontinuity promotes detachment of the flow of the propulsion gases; and a device for controlled injection of at least one additional jet of gas into the main stream of the propulsion gases exhausted by the propulsion nozzle; said divergent section further comprising said controlled injection device having one or more orifices or a group of orifices produced axisymmetrically in said second wall of said divergent section, so that the injection of the additional jet of gas is carried out radially or inclined counterflow to the main stream, each of said orifices being arranged therein at a distance d from said connecting zone comprising the curved contour discontinuity, wherein d is equal to or less than 95% of the total length L of said second portion and greater than 1% of the total length L of said second portion so that the injection of the additional jet of gas: keeps the flow of the propulsion gases detached from the second curved contour by making use of the change in slope between the first curved profile and the second curved profile when moving from operation at low altitudes to operation at high altitudes; and causes detachment of the flow of the propulsion gases from the second curved contour by making use of the change in slope between the first curved profile and the second curved profile when moving from operation at high altitudes to operation at low altitudes.

2. The divergent section according to claim 1, wherein said controlled injection device comprises a single orifice consisting of a slot having the form of a ring extending over the entire circumference of the second wall situated at the distance d from said connecting zone comprising the curved contour discontinuity, the width of the orifice making it possible to obtain a flow rate reaching up to 20% of the mass flow rate of the propulsion gases in the nozzle with an ejection speed that is subsonic or above.

3. The divergent section according to claim 1, in which said controlled injection device comprises at least three orifices having identical shape and dimensions, said orifices being regularly spaced along the circumference of the second wall which is situated at the distance d from said connecting zone comprising the curved contour discontinuity.

4. The divergent section according to claim 1, according to which said controlled injection device also comprises one or more feeding chambers, each communicating with an orifice of said injection device to homogeneously inject the additional jet of gas into said divergent section.

5. The divergent section according to claim 1, comprising several controlled injection devices situated at successive different distances d.

6. The divergent section according to claim 1, according to which the additional jet of gas injected by said controlled injection device originates from said reaction chamber of the engine, or from combustion agent or fuel tanks, or auxiliary tanks annexed.

7. A propulsion nozzle of an engine for aircraft or spacecraft propelled by chemical reaction, said nozzle comprising: a convergent portion receiving the gases produced in said reaction chamber; a nozzle throat; and a divergent section connected to the nozzle throat wherein the divergent section is as defined according to claim 1.

8. A method for controlling the transition of the propulsion gas flow regime in a divergent section comprising multiple curved contours of a propulsion nozzle of an engine for an aircraft or spacecraft as a function of whether the altitude of the aircraft or of the spacecraft increases or decreases, said method comprising, in the case in which the altitude increases, the following steps: said engine is equipped with a propulsion nozzle as defined according to claim 7; just before the natural transition altitude is reached, gas is injected radially into said nozzle via said controlled injection device; and the injection is maintained until the altitude of the aircraft or of the spacecraft reaches a transition altitude called optimal transition altitude, said method comprising, in the case in which the altitude decreases, the following steps: said engine for a spacecraft is equipped with said propulsion nozzle as defined according to claim 7; at the optimal transition altitude, gas is injected into said nozzle via said controlled injection device so as to force the detachment of the main stream from the second wall; and the injection is maintained until the altitude of the aircraft or of the spacecraft passes below said natural transition altitude.

9. A method for controlling the reattachment of the flow of the propulsion gases on the second curved contour of a divergent section comprising multiple curved contours of a propulsion nozzle of an engine of a spacecraft when the spacecraft is in the take-off or landing phase, the method comprising the following steps: said engine for a spacecraft is equipped with a propulsion nozzle as defined according to claim 7; when the engine is switched on, gas is injected radially into said nozzle via said controlled injection device until a sufficient altitude is reached guaranteeing the absence of instability; and at very low altitude, during the landing phase, gas is injected radially into said nozzle via said controlled injection device until the engine is switched off.

10. A method for thrust vectorization in a divergent section comprising multiple curved contours of a propulsion nozzle of an engine for an aircraft or spacecraft, said method comprising the following steps: said engine is equipped with a propulsion nozzle as defined according to claim 1 and further including said controlled injection device comprises at least three orifices having identical shape and dimensions, said orifices being regularly spaced along the circumference of the second wall which is situated at the distance d from said connecting zone comprising a curved contour discontinuity, and also including a convergent portion receiving the gases produced in said reaction chamber; said nozzle throat; and said divergent section connected to the nozzle throat; when it is necessary to correct the trajectory of said aircraft or spacecraft, gas is injected radially and in dissymmetric way through said orifices, while varying the injection flow rates from one orifice to another, so as to make the flow of the main stream of propulsion gas and the peripheral pressure dissymmetrical and vectorize the thrust.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) Other advantages and features of the present invention will result from the following description, given by way of non-limitative example and made with reference to the attached figures and to the corresponding examples:

(2) FIG. 1 represents a diagrammatic perspective view of a double-curved contours propulsion nozzle;

(3) FIG. 2 shows the visualization obtained by schlieren photography of a flow at the outlet of a divergent section comprising double-curved contours of a propulsion nozzle known from the prior art, for operation at low altitude (first portion 2a showing that the flow remains attached to the first curved contour and detached from the second curved contour) and operation at high altitude (second portion 2b showing that the flow remains attached to the first and second curved contours);

(4) FIGS. 3a and 3c are curves showing the evolution of thrust as a function of altitude (specific impulse as a function of the NPR, while FIG. 3b represents the evolution of the side load at the retransition then at the transition;

(5) FIG. 4 diagrammatically represents a system for propulsion by combustion equipped with a double-curved-contours nozzle equipped with a secondary injection system according to the first embodiment of the divergent section according to the invention (a single injection orifice), the secondary injection gas originating (FIG. 4a) either from the fuel tank, or from the combustion agent tank (FIG. 4b), or from an auxiliary tank (FIG. 4c);

(6) FIG. 5 diagrammatically represents a system for propulsion by combustion equipped with a double-curved-contours nozzle equipped with a secondary injection system according to the first embodiment of the divergent section according to the invention (a single injection orifice), the secondary injection gas originating from the combustion chamber without feeding chamber (a) and with feeding chamber (b).

(7) FIG. 6 represents a diagrammatic view of an example of a nozzle equipped with a divergent section comprising a curved contour discontinuity and a secondary injection system, according to the first embodiment of the divergent section according to the invention (a single injection orifice).

(8) FIG. 7 is a photograph of the nozzle illustrated in FIG. 6 with a high-pressure chamber 70 supplying the nozzle with propulsion gas, the nozzle and the high-pressure chamber being mounted on an experimental test stand;

(9) FIGS. 8a and 8b are diagrammatic representations of a nozzle equipped with a double-curved-contours divergent section and an injection system, according to the second embodiment of the divergent section according to the invention (pluralities of injection orifices): the secondary injection is carried out through four orifices (visible in FIG. 8a) each having their feeding chamber (visible in FIG. 8b);

(10) FIG. 9 shows a diagram illustrating the use of different types of orifices; the injection can be carried out through several orifices (FIG. 9a) or groups of orifices (FIG. 9b, 9c) carried out in dissymmetrical way in the second curved contour radially or slightly inclined counterflow;

(11) FIG. 10 is a diagram (with dimensions) defining the profile of an example of a double-curved contours divergent section according to the invention used for simulations and experiments;

(12) FIG. 11 shows the Mach fields obtained by numerical simulation inside a nozzle according to the invention for different NPRs in the case of an inactive (a) and active (b) secondary injection the flow rate of which corresponds to 2% of the main mass flow rate;

(13) FIG. 12 shows the evolution of thrust Fx as a function of the NPR, for an increasing NPR, for different secondary injection pressures;

(14) FIG. 13 shows the evolution of thrust Fx as a function of the NPR, for a decreasing NPR, for different secondary injection pressures;

(15) FIG. 14 shows the evolution of the secondary injection pressure Pi (measured in the feeding chamber in kPa), of the thrust Fx (in N), of the side load Fy (in N) as well as of the NPR over time;

(16) FIG. 15 shows the visualization obtained by schlieren photography (a) and by numerical simulation (b), of a flow at the outlet of a divergent section comprising a single curved contour of a propulsion nozzle known from the prior art with a radial secondary injection carried out through a single circular orifice.

(17) The identical elements shown in FIGS. 1 to 15 are identified by identical numerical references.

(18) FIGS. 1 to 5 are explained in the description of the invention corresponding to the prior art.

(19) FIGS. 6 to 9 are described in more detail in the following part of the description, relating to the detailed description of the double-curved-contours divergent sections according to the invention (three embodiments) and of a nozzle according to the invention.

(20) FIGS. 10 to 15 are described in more detail in the following examples, which illustrate the invention without limiting the scope thereof.

DETAILED DESCRIPTION

(21) FIG. 6 represents a diagrammatic view of an example of a nozzle 4 equipped with a divergent section 1 comprising a curved contour discontinuity according to the first embodiment according to the invention (a single injection orifice), forming part of a propulsion system (shown on its test stand in FIG. 7) for an aircraft or spacecraft propelled by chemical reaction comprising a high-pressure chamber 70 (upstream) coupled with an injection nozzle 4. The nozzle 4 is a nozzle having rotational symmetry about an axis 41 (FIG. 6). The high-pressure gas is expanded and ejected towards the outside through the nozzle 4 essentially in a main stream 42 (FIG. 6). The divergent section 1 comprises from the nozzle throat 3: a first portion 11 comprising a first wall 110 having a first surface of revolution about the axis of the nozzle 41, so that the first wall 110 has a first curved profile (or first curved contour) in an axial plane, a second portion 12 having a total length L, is connected to said first portion 11 at a connecting zone 13 defining a curved contour discontinuity, the second portion 12 comprising a second wall 120 having a second surface of revolution about the axis of the nozzle 41, so that the second wall 120 has a second curved profile (or second curved contour) in an axial plane, and also a device 14 for controlled injection of at least one additional jet of gas into the main stream of gas 42 exhausted by the nozzle 4.

(22) The injection of the fluid may be carried out by an orifice (141, FIG. 6) or more orifices or group of orifices (142, 143, 144: FIG. 8a) each having their feeding chamber (146, 147, 148 to 89: FIG. 8b), or also by several groups of orifices (FIG. 9b, 9c) produced axisymmetrically in the second curved contour 120, so that the additional injection of the additional jet of gas is carried out radially or slightly inclined counterflow. The orifices are arranged, for the first two embodiments (illustrated in FIG. 6 on the one hand, and FIGS. 8, 9a and 9b on the other hand) at a distance d from the connecting zone 13 comprising a curved contour discontinuity, in such a way that the injection allows, in the case of the high altitude operation regime (attached to the second curved contour) to the shock of detachment to stabilize over the connecting zone 13.

(23) In the third mode of the present invention (plurality of groups of orifices: FIG. 9c), several injection systems may be used so as to prevent reattachment on the second curved contour 120 downstream of the primary injection.

EXAMPLES

(24) A nozzle according to the invention (as shown in FIGS. 6 and 7) is used, produced as a function of the experimental capabilities of the test stand. The initial conditions (upstream of the nozzle throat) are 3.5 bar and 290 K, the nozzle throat has a radius r.sub.t of 8.5 mm, which gives a theoretical flow rate of 0.2285 kg/s. The profile of the nozzle is obtained using the inverse characteristics method initiated by the Sauer method.sup.[5] applied to transonic flow. The ideal nozzle is truncated (length L.sub.b) to obtain an output angle of 4.4°. The extension (length L.sub.e) is also determined by the method of characteristics so as to have a constant pressure at the wall. The angle at the connecting zone 13 is 8°. These data are summarized in FIG. 10.

(25) FIG. 11 illustrates the operating principle of the injection system according to the present invention in the case of an increase in the altitude (increasing NPR), without secondary injection (FIG. 11a) and with active secondary injection (FIG. 11b).

(26) This calculation is carried out using the FASTRAN commercial software program written specifically for fluid mechanics applications in respect of fluids compressible at high Mach numbers. It solves Reynolds-averaged Navier-Stokes equations (RANS) by the finite volumes method, the non-viscous terms being calculated using the Van Leer.sup.[6] scheme with a third-order Osher flux limiter. The turbulence model used is the K-Omega SST-Menter. Multi-domain meshing is used. The initial conditions are: temperature 290 K, 3.5 bar. The injection is situated at x/r.sub.t=6.483, x being the distance between the injection and the nozzle throat and r.sub.t being the radius of the nozzle throat. The injection orifice has a width of 0.5 mm, the injection pressure is 0.4 bar, and the injected flow rate is approximately 2% of the main stream rate of the nozzle (flow rate of the propulsion gas).

(27) In FIG. 11a showing the field of Mach numbers (speed/speed of sound) for different NPRs (altitude), it can be seen that, without secondary injection, the detachment point denoted 80 (FIG. 11a) displaces downstream as the NPR increases. From NPR=11, the shock of detachment is attached at the second curved contour, while at NPR=14 it has practically reached the outlet lip of the second curved contour. FIG. 11b shows the field of Mach numbers for different NPRs in the case of an active secondary injection 81 (2% of the primary mass flow rate). It shows that it is possible to push back the reattachment altitude onto the second curved contour. In fact, the detachment point 82 is in this case maintained on the connecting zone 13 comprising a curved contour discontinuity up to an NPR of 14, while without injection, this detachment point 80 is practically situated at the outlet of the second curved contour 120. For an NPR of 16, the secondary injection is no longer significant enough (in terms of flow rate) to continue to maintain the detachment, the flow thus reattaches behind the injection 81.

(28) This predictive calculation is verified experimentally using the set-up shown in FIG. 7 as illustrated in FIGS. 12 and 13, which show respectively the evolution of thrust Fx as a function of the NPR in the case of an increasing (FIG. 12) and decreasing (FIG. 13) NPR for different secondary injection pressures (measured in the feeding chamber 145):

(29) In FIG. 12 it can be seen that a secondary injection delays the transition to a higher NPR (83 with a secondary injection at 50 kPa and 84 with a secondary injection at 77 kPa), compared with the natural transition NPR 80.

(30) In FIG. 13 it can be seen that a secondary injection triggers the retransition to a higher NPR (93 with a secondary injection at 50 kPa, and 94 with a secondary injection at 77 kPa), compared with the natural retransition 90, which is the advantageous effect sought. The secondary injection induces not only a more favourable transition and retransition, but also a reduction, or even an elimination of damaging side loads.

(31) FIG. 14 shows the evolution of the secondary injection pressure Pi (measured in the feeding chamber), of the thrust Fx, of the side load Fy as well as of the NPR over time during an experiment. A drop 140 in thrust at the increase of NPR, and an increase 141 in thrust at the decrease of NPR is observed, which indicates that the injection pressure is not high enough (curve 142) to reach the optimal operation described previously. Despite this, the side load peak is no longer observed as in the case without injection (peaks referenced 36 in FIG. 3b), which is the advantageous effect sought.

(32) FIG. 15 shows the visualization obtained by schlieren photography (portion a in the Figure) and by numerical simulation (portion b in the Figure), of a flow at the outlet of a divergent section comprising a single curved contour of a propulsion nozzle known from the prior art with a radial secondary injection carried out via a single circular orifice with a diameter of 5.8 mm. The injection is situated at 88% of the total length of the divergent section, the mass flow rate of the secondary injection is 8% of the main stream rate of propulsion gas. In FIG. 15 it is observed that the secondary injection causes a partial detachment of the flow in the nozzle. This injection thus causes a dissymmetry in the flow and in the pressure distribution on the wall. The thrust force obtained is thus no longer aligned with the axis of the nozzle. Because of this, it becomes possible to control the trajectory of an aircraft or of a spacecraft without other specific trajectory control systems.

REFERENCE LIST

(33) [1] European patent application EP 2103799 by MITSUBISHI HEAVY INDUSTRIES. [2] French patent application FR2618488 by Société Européenne de Propulsion (SEP). [3] European patent application EP2137395 by MOSCOW AVIATION INSTITUTE. [4] American patent U.S. Pat. No. 3,394,549 by AMERICAN ROCKWELL CORPORATION. [5] R. Sauer, “Dreidimensionale Probleme der Charakteristikentheorie partieller Differential-gleichungen”, Zeitschrift für angewandte Mathematik und Mechanik, Vol 30, pp 347-356, November-December 1950. [6] Van Leer, B. (1979), “Towards the Ultimate Conservative Difference Scheme, V. A Second Order Sequel to Godunov's Method”, J. Com. Phys., 32, 101-136.