FR2882402A1 - Non stationarylateral force reducing method for jet nozzle, involves positioning circular body inside divergent section, so that shock wave is incident on section at axial incidence position, during starting phase of engine - Google Patents

Abstract

The method involves positioning a circular body (5) inside a divergent section along the length of a correspondence axis from an axial position of the body so that a shock wave is incident on the section at an axial incidence position, during the starting phase of a rocket engine. The wave is induced by a disturbance of a combustion gas flow produced by the body. The section produces a jet separation, in the incidence position. An independent claim is also included for a device for reducing non stationary lateral forces acting on a pipe of a rocket engine, during the starting phase of the engine.

Description

DEVICE AND METHOD FOR REDUCING SIDE FORCES

  NON STATIONARY ACTING ON A TUYERE OF A FUSEE ENGINE.

  The invention relates to a device and a method for reducing non-stationary lateral forces acting on a nozzle of a rocket motor, particularly during the starting or ignition phase of said engine.

  In particular, the implementation of the invention makes it possible to eliminate, or less substantially limit, the non-stationary lateral forces that occur in the nozzle of a rocket engine when it is ignited because of the effect known as jet separation, or delamination of the boundary layer, or separation by internal recirculation in the jet.

  It is known that the thrust of a rocket engine depends on its mass flow rate, the stop pressure ps in the combustion chamber, the expansion ratio of the nozzle, that is to say the ps / pe ratio between the stop pressure ps and the static gas ejection pressure at the outlet of the nozzle pe, and depends on the ambient pressure pa. This thrust is maximum, for given operating conditions of the chamber, when the two pressure values pe and pa are equal (the nozzle is then said to be adapted). It is also known that rocket engines are normally designed to achieve the adaptation condition pe = pa at an altitude greater than that of launch, for example of the order of 10,000 m and that, therefore, at low altitude a pe <pa (operating condition of the nozzle in overdrive). If the static gas ejection pressure at the exit of the nozzle pe is significantly lower than the ambient pressure pa (for example less than 0.2 Pa) a jet separation occurs inside the divergent nozzle . Two types of separation are known: free shock wave separation and internal recirculation separation in the jet.

  For certain nozzle geometries and / or under certain expansion ratio conditions the boundary layer of the supersonic gas flow separates from the divergent wall of the nozzle, in a separation mode called free shock wave separation mode. , and the jet is compressed by an oblique shock wave and a lateral force is applied locally at any point of the wall of said divergent downstream of the separation point. This lateral force is created by the pressure difference between the outer wall of the divergent on which the atmospheric pressure is applied and the inner wall of the divergent on which the local static pressure of the jet is applied. If the jet separation is perfectly symmetrical and stable over the entire circumference of the nozzle and at a determined axial position, the local static pressure of the jet would be uniform over the circumference of the nozzle and the resultant of these lateral forces would be zero. . In reality, the jet separation line has an irregular and highly unsteady shape. It follows that at each instant the jet separation induces a resultant non-zero force, which may present a considerable moment with respect to the throat of the nozzle, at the place where the moment of structural inertia of the engine is the weaker. It is easily understood that the most critical situation occurs when the jet separation takes place mainly on one side of the nozzle and near its outlet section. For other nozzle geometries, or with other expansion ratio conditions for the same nozzle, a different separation mode may occur, which is referred to as the internal recirculation separation mode in the jet, and which also generates unsteady side forces. In this mode of separation, as in the previous one, the boundary layer of the supersonic flow of gas separates from the wall of the divergent nozzle, but because of the level of the pressures downstream the flow is immediately recollected on the wall. The toroidal separation bulb is controlled in position by a shock which is created in the center of the flow by a large recirculation of gas which moves randomly, resulting randomly in the formation of a toroidal separation bulb. position of the central shock and the position of the toroidal separation bulb. Downstream of the toroidal separation bulb, the supersonic jet remains attached to the wall, but is compressed by an oblique shock wave whose intensity varies with the Mach number of the incident flow, and therefore with the position of the bulb. toroidal separation. As a result, the static pressures at the wall downstream of the toroidal separation bulb vary randomly. As in the case of the free shock wave separation, a lateral force applies locally at any point of the wall of said divergent downstream of the separation point. This lateral force is created by the pressure difference between the outer wall of the divergent on which the atmospheric pressure is applied and the inner wall of the divergent on which the local static pressure of the jet is applied. If the toroidal separation bulb were produced perfectly symmetrically, coaxial with the nozzle, stable with time over the entire circumference of the nozzle and at a given axial position, the recompression due to the shock it generates would be uniform and the The local static pressure of the downstream jet would be uniform over the circumference of the nozzle and the resultant of these lateral forces would be zero. In fact, for the reasons explained above, the separation line of the toroidal separation bulb has an irregular and highly unsteady shape. It follows that at each instant the internal jet recirculation separation induces a non-zero resultant force, which may present a considerable moment with respect to the neck of the nozzle, at the place where the moment of structural inertia of the nozzle engine is the weakest. It is easily understood that the most critical situation occurs when internal jet recirculation separation takes place mostly in one half in a vertical section of the nozzle and near its outlet section.

  The need to maintain the non-stationary forces induced by the jet separation to an acceptable level makes it necessary to limit the value of the ps / pe expansion ratio below its optimum value and to oversize the structure of the nozzle, which reduces the overall engine performance and thrust-to-weight ratio. Despite these precautions, the non-stationary forces caused by the jet separation determine significant vibrations that can damage the nozzle and even lead to its rupture if over time the random pressure distribution in the divergent takes a too unfavorable value.

  An in-depth analysis of the jet separation phenomena in the rocket engine nozzles and the resulting non-stationary lateral forces is carried out in the article by G. Hagemann, M. Terhardt, M. Frey, P. Reijasse, M. Onofri, F. Nasuti and J. Stlund, Flow Separation and Side-Loads in Rocket Nozzles, presented at the 4th International Symposium on Liquid Space Propulsion (4th International Symposium on Liquid Space Propulsion), Lampoldshausen, Germany, 13 - 15 March 2000 .

  Numerous devices have been proposed to control the separation of the jet inside a rocket engine nozzle in order to limit said non-stationary lateral forces. In particular; US 6,572,030 discloses the use of a releasable annular structure, extending in a radial direction, to be placed around the outlet section of the nozzle. This structure induces the formation of a low pressure zone near said outlet section, which reduces jet separation within the nozzle.

  US 5,894,723 discloses the use of ejectable inserts placed inside the nozzle. The ejection of said inserts makes it possible to increase during the ascent the ratio between the area of the outlet section of the nozzle and that of its neck, so as to allow the operation of the engine in conditions close to the adaptation throughout the climbing phase of the rocket.

  US 5,490,629 discloses the use of an ejectable diffuser, connected to the outlet section of the nozzle and having a constriction to re-compress the gases and thus prevent separation of the jet during the first part of the trajectory rocket.

  US 5,481,870 discloses the use of a releasable annular obstacle, connected to the outlet section of the nozzle and partially obstructing it so as to induce a stable jet separation.

  US 5,450,720 discloses the use of longitudinal notches in the downstream end portion of the nozzle to induce stable jet separation.

  All of these documents disclose solutions to the problem of eliminating or limiting the non-stationary lateral forces that occur in the nozzle of a rocket engine during the first part of its ascent phase, from takeoff to landing. altitude at which the adaptation condition is reached. However, none of the devices described therein makes it possible to limit the appearance of nonstationary lateral forces while the stop pressure in the combustion chamber of the engine has not yet reached its nominal value, that is to say before even the launch of the rocket, during the starting phase of the engine. During this phase, whose duration is of the order of one second or slightly lower, the stopping pressure ps of the gases in the combustion chamber increases rapidly from the atmospheric level to a maximum value and, consequently, the average position of the jet separation line moves towards the outlet section of the nozzle, rendering inoperative the control means known from the prior art whose geometric definition is fixed relative to the nozzle. In addition, these documents propose the use of devices which are integral with the nozzle during at least part of the rocket's ascent phase, and which therefore increase the mass, which is contrary to one of the Jet separation control objectives, which is to lighten the nozzle through the reduction of the efforts to which it is subjected.

  The only known document of the prior art which presents a solution to the problem of limiting the non-stationary lateral forces during the starting phase of the engine without making the nozzle heavier is the document FR 2 791 398, which discloses a system for stabilizing the engine. jet separation comprising a device external to the engine, integral with a ground installation, constituted by a set of injection tubes directing countercurrent fluid jets inside the nozzle towards points of impact on the wall of it. A jet separation region occurs from each point of impact and extends to the outlet section of the nozzle in a conical configuration. Such a system generally reduces the non-stationary lateral forces in the nozzle and has the advantage of being carried by a ground installation and not by the rocket itself, but it is not entirely satisfactory because it does not allow to effectively stabilize the jet separation during the entire engine start-up phase since the countercurrently directed jets impact the nozzle at fixed and non-pressure dependent positions in the combustion chamber. In addition, as shown in FIG. 1 of document FR 2 791 398, the jet separation lines starting from each point of impact intersect the edge of the exit section of the nozzle.

  However, these lines have a stable position in correspondence of the point of impact of the jet of fluid against the current which triggers, but can fluctuate downstream of this point, which induces residual non-stationary forces. This is all the more troublesome as these residual forces apply correspondingly to the edge of the outlet section of the nozzle, that is to say where they are more harmful, because their moment relative to the neck of the nozzle. the nozzle is maximum.

  An object of the invention is to obtain improved control of the jet separation during the starting phase of a rocket engine, and thereby to reduce the non-stationary forces acting on the diverging nozzle.

  Another object of the invention is to obtain such a control without increasing the mass of the nozzle or the rocket.

  Another object of the invention is to obtain such a control by means of a simpler and more economical device than the devices known from the prior art.

  Yet another object of the invention is to propose a device and a method for controlling the jet separation during the starting phase of a rocket engine that can be used together with a device known from the prior art to control the jet separation during the ascent phase of the rocket, so as to obtain such control for a large part of the operating period of the engine in the atmosphere.

  Yet another object of the invention is to make it possible, by means of improved control of the jet separation, to improve the overall performance of a rocket engine by making it possible to lighten its structure and to increase its detent ratio. .

  At least one of these objects is achieved by a method for reducing non-stationary lateral forces acting on a nozzle of a rocket motor during a starting phase of said engine, said engine comprising a combustion chamber where gases combustion are generated, a divergent in which a supersonic flow of said combustion gases occurs and a neck connecting said combustion chamber to said divergent, characterized in that it comprises the positioning of a rounded body in the interior of the diverging along its axis in correspondence of an axial position of the rounded body such that, during at least a portion of said starting phase, a shock wave, induced by the disturbance of the flue gas flow caused by said rounded body, is incident on the wall of said divergent at an axial position of incidence where it produces a jet separation or a separation in the form of a toroidal separation bulb.

  According to particular embodiments of the invention: the method may also comprise: prior to starting the engine, inserting said rounded body inside the divergent along its axis, to a first position axial; and during the start-up phase, moving said rounded body along the axis of the nozzle according to a value of the stopping pressure of the combustion gases in the combustion chamber so that during that the value of said stop pressure of the combustion gases varies during said starting phase, said shock wave continues to be incident on the wall of the divergent at an axial position where it produces a jet separation or a separation in form of toroidal separation bulb.

  The displacement of said rounded body along the axis of the nozzle as a function of a value of the stop pressure of the combustion gases in the combustion chamber can be carried out in such a way that said shock wave is incident on the wall of said divergent at an axial position corresponding to the downstream limit of the region of said divergent where spontaneous jet separation or spontaneous separation in the form of a toroidal separation bulb does not occur under the action of ambient pressure.

  - The displacement of said rounded body as a function of the stop pressure of the combustion gases in the combustion chamber can follow a setpoint which is determined using the following steps: selection of a series of discrete values of the stopping pressure of the combustion gases in the combustion chamber, between atmospheric pressure and a maximum value reached during the starting phase; selecting a series of discrete values of the position of said body of rounded shape along the axis of the divergent, between the position of the neck and that of the outlet section of said divergent; for each pair of said discrete values, determination by calculation or by testing of the static pressure value and the Mach number along the wall of the divergent and determination of the point of impact impact on the wall of the nozzle; for each pair of said discrete values, determining the axial position of the jet separation point or the toroidal separation bulb-shaped separation point using said static pressure values and the Mach number of the gas flow of combustion along the wall of the diverging; for each of said discrete values of the combustion gas pressure in the combustion chamber, determining the position of said most downstream round-shaped body such that the jet separation or the toroidal separation bulb-shaped separation is caused by said shock wave induced by the presence of said rounded body; said most downstream value being selected as the setpoint value of the position of said rounded body in correspondence of said value of the stop pressure of the combustion gases in the combustion chamber.

  - The method may also comprise an interpolation operation of said setpoint values of the position of said rounded body in correspondence of said values of the stop pressure of the combustion gases in the combustion chamber so as to determine a position command in an analytical form.

  Said axial position of the spontaneous jet separation point or the spontaneous separation in the form of a toroidal separation bulb under the action of ambient pressure can be determined using a suitable empirical or semi-empirical criterion.

  - Said rounded body can be moved from said axial position to the outlet section of the divergent during the starting phase of the engine as the stopping pressure of the combustion gases in the combustion chamber increases.

  Said value of the stopping pressure of the combustion gases in the combustion chamber can be determined directly by a pressure measurement in the combustion chamber or indirectly by measuring the stopping pressure of said combustion gases. in correspondence of a vertex of said body of rounded shape.

  At least one of the aforementioned objects is also achieved by means of a device for reducing non-stationary lateral forces acting on a nozzle of a rocket motor during a starting phase of said engine, said engine comprising a combustion chamber where combustion gases are generated, a divergent in which a supersonic flow of said combustion gases occurs and a neck connecting said combustion chamber to said divergent, characterized in that it comprises: a rounded body for be positioned within the divergent along its axis; and means for positioning said rounded body within the divergent corresponding to an axial position of the rounded body such that, during at least a portion of said starting phase, a shock wave induced by the disturbance of the flue gas flow caused by said rounded body is incident on the wall of said divergent at an axial position of incidence where it produces a jet separation or a toroidal separation bulb-shaped separation .

  According to particular embodiments of the invention: the device may also comprise means for moving said rounded body along the axis of the divergent according to a value of the stop pressure of the combustion gases in the combustion chamber such that, while the value of said stopping pressure of the combustion gases varies during said start-up phase, said shock wave continues to be incident on the wall of the divergent at an axial position where it produces a jet separation or toroidal separation bulb separation.

  - Said means for moving said rounded body along the axis of the divergent may include an actuator for moving said rounded body along the axis of the diverging.

  The device may also comprise a controller for receiving from a first sensor information of value of the stop pressure of the combustion gases in the combustion chamber and to command said actuator to move said rounded body along the combustion chamber. axis of the divergent according to said value information of the stop pressure of the combustion gases in the combustion chamber.

  - Said controller may be a controller for controlling the displacement of said rounded body along the axis of the divergent nozzle according to a position setpoint determined by the following steps: choosing a series of discrete values of the stopping pressure of the combustion gases in the combustion chamber, between atmospheric pressure and a maximum value reached during the starting phase; selecting a series of discrete values of the position of said body of rounded shape along the axis of the divergent, between the position of the neck and that of the outlet section of said divergent; for each pair of said discrete values, determination by calculation or by testing of the static pressure value and the Mach number along the wall of the divergent, and determination of the point of impact impact on the wall of the nozzle ; for each pair of said discrete values, determining the axial position of the jet separation or separation point in the form of a toroidal separation bulb using said static pressure values and the Mach number along the divergent wall; for each of said discrete values of the stopping pressure of the combustion gases in the combustion chamber, determining the position of said further downstream shaped body, such as the jet separation or the separation of the separation bulb toroidal is caused by said shock wave induced by the presence of said rounded body; said most downstream value being selected as the setpoint value of the position of said rounded body in correspondence of said value of the stopping pressure of the combustion gases in the combustion chamber.

  Said position instruction may have an analytical form, determined by means of an additional operation of interpolation of said setpoint values of the position of said rounded body in correspondence with said values of the combustion gas stop pressure. in the combustion chamber.

  Said axial position of the spontaneous jet separation point or the spontaneous separation point in the form of a toroidal separation bulb under the action of the ambient pressure can be determined using a suitable empirical or semi-empirical criterion.

  Said controller may be a controller for controlling the displacement of said rounded body along the axis of the divergent nozzle from said axial position to the exit section of the divergent during the starting phase of the engine as it goes that the stopping pressure of the combustion gases in the combustion chamber increases.

  - The rounded body may have a concave surface having a vertex to be oriented towards the throat of the nozzle and a pressure sensor disposed in correspondence of said apex.

  Said means for moving the round-shaped body along the axis of the divergent as a function of a value of the stop pressure of the combustion gases in the combustion chamber may comprise a means for applying a force elastic member opposing the expulsion of said rounded body of said divergent by said combustion gases.

  - Said body of rounded shape may have axial symmetry and has a rounded surface intended to be oriented towards the neck 5 of the nozzle.

  - Said rounded body may have a cross section of between 0.5 and 2 times, and preferably between 0.8 and 1.5 times, the neck section of the nozzle.

  The device may also include a mechanical fuse for expelling the rounded body of the divergent when an axial force exerted on said body exceeds a predetermined threshold.

  Other features, details and advantages of the invention will become apparent on reading the description given with reference to the appended drawings, given by way of example, and which show: FIG. 1, a device for controlling separation of jet according to the invention disposed in operational configuration in the nozzle of a rocket engine which has a spontaneous mode of separation by free shock wave; FIG. 2 is a flowchart illustrating a method for determining a control law of the device of FIG. 1, allowing the implementation of a method according to the invention; FIG. 3, the same jet separation control device according to the invention disposed in operational configuration in the nozzle of a rocket engine which has a spontaneous internal re-circulation separation mode in the jet; - Figure 4, the same jet separation control device according to the invention which is added a protective device against mechanical overload; and FIG. 5, a simplified version of the jet separation control device according to the invention which operates in a restricted range of expansion ratio.

  A rocket engine comprises a combustion chamber 1 in which gases are generated at high temperature and high pressure (stop pressure ps), and a nozzle comprising: a convergent connected to said combustion chamber 1, a neck 2 in which the flow of said gases reaches transonic and divergent conditions 3 wherein said flow expands and accelerates at supersonic velocity. The outlet section 4 of the divergent opens out of the engine, in an environment where there is an external pressure pa, which is about 1 atmosphere at the launch altitude and decreases during the ascent of the rocket to a value negligible when it leaves the Earth's atmosphere.

In the following description:

  x indicates the distance between any point of the divergent 3 and the neck 2, measured along the axis of the nozzle; - A (x) is the effective aerodynamic area of the nozzle section at a distance x from the neck; the aerodynamic area is given by the geometric area minus the area corresponding to the displacement thickness of the boundary layer; At = A (x = 0) is the effective aerodynamic area of section 20 of collar 2; Ae is the effective aerodynamic area of the outlet section 4 of the divergent 3; p (x) is the static pressure of gases at pressure x; pe is the static pressure of the gases in correspondence of the outlet section 4 of the divergent3; it is considered that at the launch altitude the nozzle is in overdrive mode, that is to say that the value of the ambient (atmospheric) pressure pa is substantially greater than the theoretical value that pe would take if the engine was operating in the empty; - M (x) is the Mach number, equal to the ratio between the speed of the flow and the speed of sound in the gases generated in the combustion chamber; and y is the exponent of isentropic expansion of said gases.

  In the supersonic flow zone isentropic, upstream of any shock generated by a detachment or by the body of rounded shape, the values of A (x) and M (x) are linked by the following relation: r +, y + l A (x) (2 v2 (y ') 1 (1 + 7 -1M2 (xi 2 (rI) [1] Ai j + 1y M (x) 2 If, as a first approximation, we consider that A (x) is known and independent of the conditions of the flow, the relation [1] allows to calculate the Mach number in each point of the nozzle.The value of the pressure p (x) in the nozzle depends on the Mach number and the stop pressure in the combustion chamber, ps: rp (x) = il + Y-1M2v, -r ps \ 2 i Equations [1] and [2] express the characteristics of the flow inside the nozzle in the entire isentropic supersonic flow zone, not only for operation in a vacuum, when pa = 0, but also when the nozzle operating at altitude is adapted (pa = pe), or even in the regime of slight overdrive ( pa> pe) In the latter case, the gases are compressed at the exit of the nozzle by shock waves caused by delamination of the boundary layer and equations [1] and [2] are not sufficient to determine the speed and pressure of the gas downstream of these shock waves. It is under these conditions of overdrive that the separation effect of the jet occurs, as described by the aforementioned article by G. Hagemann et al. Various theoretical, empirical or semiempirical criteria have been proposed for determining the position x_sep at which the separation occurs. One of the simplest and most used is the Schmucker criterion: p (x sep) _ [1,88M (xsep) -11-0'64 [3] Pa where M (x) is given by the equation [ 1] and p (x) by equation [2]; in fact, the point of separation is situated at the downstream limit of the isentropic zone, which makes it possible to use the relations (1) and (2) to determine [2] the parameters p and M in criterion (3). For x> x_sep the gas pressure increases and no longer obeys equation [2].

  It must be kept in mind that the criterion expressed by equation [3] gives only an estimate of an average position of the separation point of the jet which, in fact, fluctuates over time: in practice one can instead, speak of a jet separation region centered around the position x = x_sep. NASA considers, for example, that for the nozzle geometries it has studied this region generally ranges between x_sep-20% and x_sep + 20%: such an estimate should provide a sufficiently accurate range in most cases, but in fact a more precise study may advantageously be carried out for each particular nozzle geometry.

  Of course, it is also possible to numerically simulate the flow, instead of relying on analytical models. In addition, more sophisticated models can be used, taking into account the viscosity of the gases, the non-isentropic nature of the flow, the properties of the boundary layer, and so on. Moreover, it must be considered that the exact mechanism of generation of non-stationary lateral forces is not yet fully understood (see the aforementioned article by G. Hagemann et al.).

  So far we have considered the case where ps is fixed and pa varies. In fact, when starting the motor pa is fixed and ps increases gradually from a value equal to pa up to a value of regime in an interval of the order of one second.

  Initially, ps = pa and no flow occurs. Then, to increase the ps pressure in the combustion chamber 1, a flow occurs, first under subsonic regime. At a certain value of ps, a shock wave appears at the neck 2 of the nozzle and M (x = 0) = 1; when ps increases further, the shockwave moves in the divergent 3 to the outlet section 4 and a jet separation occurs approximately in correspondence of the position x_sep which satisfies the criterion [3] (or any other suitable criterion) . To increase it by ps, the separation region moves further downstream. This has two consequences: the area over which the non-stationary lateral forces are applied increases, resulting in the increase of the intensity of said forces, and the moment of the resultant of these forces with respect to the neck also increases because moving its point of application. It is therefore towards the end of the start-up phase that the forces to which the structure of the nozzle is subjected are the most important.

  Normally, it is under these conditions that the maximum stopping pressure is reached; thereafter, after take-off, ps remains constant and ps decreases as the rocket rises. The adaptation regime is thus reached, and then that of subdelegation (pe> pa).

  The topology of nozzle flow in the overdrive regime can exhibit two types of separation during engine start-up, free-shock-wave separation and internal recirculation separation in the jet, also known as separation of the jet. jet with a limited shock wave (see the aforementioned article by G. Hagemann et al.). However, it should be noted that during the priming of a nozzle in the overdrive regime, for example during start-up of a rocket motor, the expansion ratio increases with the pressure in the combustion chamber ps ,, and the shock pattern that appears first is the simplest configuration, with free shock wave separation and detachment at the wall as described above. It is therefore possible to determine by the method described above, less in an approximate manner, in which region of a particular nozzle initially occurs the jet separation.

  The idea underlying the present invention, illustrated by Figure 1, is to induce the formation of a detached and stable predictable shock wave 8 inside the nozzle, the shock wave being incident on the wall of the nozzle at a position 9, the interaction between the shock wave 8 and the boundary layer of the flow in the vicinity of the wall of the nozzle causing the detachment of said boundary layer, that is to say the separation of the jet; in FIG. 1, the reference 10 indicates the limit of the separation region and the reference 11 the shock wave reflected by the point of incidence 9. Unlike the case of natural jet separation, induced by the ambient pressure Pa , the jet separation caused by the shock wave 8 is stable, localized and independent of any disturbance from the downstream, which allows to eliminate, or at least substantially reduce, non-stationary lateral forces. The physics of the interaction between a shock wave and a boundary layer is described in the article by J. Delery and R. Bur The Physics of Shock Wave / Boundary Layer Interaction Control: Last Lessons Learned, ECCOMAS 2000, Barcelona (Spain). ), 11-14 September 2000.

  The inventor has realized that one of the reasons why the device of document FR 2 791 398 does not make it possible to effectively stabilize the jet separation during the entire starting phase of the engine is the fact that the position points of impact of the counter-current jets which trigger the separation of the jet, remains constant. Thus, during the first phase of starting the engine, a spontaneous jet separation can occur upstream of said impact points, while in the last phase of starting these same points may be too far upstream. Indeed it has been found that there is an optimal position of the point of incidence 9 of the shock wave 8 as a function of the value of the stop pressure ps. Indeed, said point 9 must be located sufficiently upstream to determine only the separation of the jet, before the effect of the ambient pressure pa begins to be felt, but at the same time sufficiently close to the point where a spontaneous separation under the effect of atmospheric pressure to meet a weakened boundary layer, and therefore easy to take off. In conclusion, said optimal position of the point of incidence 9 is the furthest downstream position, such that the jet separation is nevertheless induced by the shock wave 8 rather than being spontaneous. For this reason, the invention makes it possible to move the point of incidence 9 of the shock wave 8 during the starting phase of the engine, so that said point of incidence 9 is constantly close to its optimum position.

  FIG. 1 shows that a device according to the invention comprises a body of rounded shape 5, mounted on a rod 14 which can be moved in an axial direction by an actuator 6 controlled by a controller 15. The latter receives as input a value pressure ps of combustion gases inside the chamber, obtained by a first sensor 12, and a position information of said rounded body 5, obtained by a second sensor 13 and drives the actuator 6 in such a way that the axial position of said body 5 is constantly close to a set value which depends on said pressure value ps. In operational conditions, the body 5 is disposed inside the diverging portion 3 of the nozzle, so as to be able to move along its axis.

  It is observed that the device, consisting essentially of the body 5, the rod 14, the actuator 6, the second sensor 13 and the controller 15 (the first sensor 12 is generally provided in the combustion chambers of the rocket engines) is solidary launch pad 7: therefore, it does not weigh down the rocket and can be reused at least partially, after replacing the damaged elements.

  In the case illustrated in Figure 1, the body 5 has a hemispherical shape, but it is not a choice and other forms are conceivable. This body 5 has a rounded surface oriented towards the neck 2 of the nozzle so as to generate a detached shock wave 8 of blunt form. The rounded shape of the body 5 has the advantage over a pointed form of reducing the level of flow received by the body 5; however, if a high flow level is considered acceptable, a sharp shape could also be used. What is important is the interaction of the impact with the boundary layer at the wall of the nozzle, and not the shape of the impact near the body 5. It will therefore be understood that the exact shape of the body that generates this shock is of secondary importance. However, it is preferable that, as in the case of the hemisphere of FIG. 1, a sharp line of separation separates the rounded front surface from a rear surface, which may for example be flat, to avoid the creation of a phenomenon. unsteady which would be clean to the body of rounded form. As regards the dimensions of the body 5, it has been found that they should preferably be of the same order of magnitude as the section of the neck 2; typically, the diameter of said body is between 0.5 and 2 times and preferably between 0.8 and 1.5 times that of the neck. In any case, it is important that the body 5 is not large enough to form a secondary neck, especially in its most upstream position.

  Generally it is preferable that the body 5 has an axial symmetry; however, if the nozzle has a non-circular shape, for example, an octagonal shape, it is desirable that the body 5 has a shape, which for this example will be roughly octagonal, generating a shock whose intersection with the nozzle can trigger separation of the boundary layer over the entire circumference of the nozzle prior to the local occurrence of the natural separation, the position of which in this case may depend on the location along this circumference. The optimal shape and dimensions of the body 5 can be determined for each specific case by means of numerical simulations and / or wind tunnel tests. An important optimization parameter is the angle of incidence of the shock wave 8 on the divergent wall: in fact, the greater this angle, the lower the local mechanical forces and the overheating to which said wall is subjected. ; at the same time, too oblique incidence may not allow effective separation of the jet. One of the reasons why it is important to control the position of the body 5, and therefore of the point of incidence 9, during the starting phase of the engine is, precisely, that if said point of incidence 9 is located little upstream of the point where the spontaneous jet separation would occur, the boundary layer is strongly weakened and the separation of the jet can be induced even by a shock wave 8 strongly oblique.

  The rounded body 5 must be able to withstand the jet of combustion gas for about 1 second, or less. A large choice of materials is available for its realization: metals, refractory ceramics, but also composite materials, and even the heartwood of oak, if it is a consumable renewed with each shot.

  The design of the rod 14 and the actuator 6 does not call for special remarks, except for the fact that said rod must be able to withstand without buckling a large axial force, exerted by the flow of combustion gases. on the body 5.

  If the axial force exerted on the body 5, and thus on the rod 14, is too great, the latter may bend or break, and the body 5 strike and damage the wall of the nozzle; the sudden drop in pressure that accompanies breakage of the rod can also affect the integrity of the nozzle. Such a situation may arise, for example, if due to a failure the body 5 remains stuck at a given axial position instead of progressively retracting. To avoid this kind of risk, it is advantageous to provide a mechanical fuse, that is to say a part that yields or breaks under the force allowing the expulsion of the body 5 before a dangerous accumulation of pressure occurs in malfunction of the actuator or controller. FIG. 5 shows an example of such a mechanical fuse, indicated by the reference 18. In this embodiment, the actuator 6 is connected to a support 20 in the form of a hollow cylinder via rods 18; in the event of a mechanical overload exerted in an axial direction by the combustion gases, the latter break and the assembly constituted by the body 5, the rod 14 and the actuator 6 is ejected into said hollow cylinder.

  A particularly important element of the device of the invention is constituted by the controller 15 which controls the displacement of the body 5 during the starting phase of the engine, in a direction generally going from the neck 2 to the outlet section 4 of the nozzle, in depending on the stopping pressure ps of the gases in the combustion chamber 1. In fact, the controller itself can be of a conventional type (for example a digital PID); what is really important is the determination of the setpoint x (ps) of the position of the body 5. Indeed, as it has been shown above, it is advantageous that at every moment the body 5 is close to a optimum position defined as the most downstream position, such that the jet separation is nevertheless induced by the impact of the shock wave 8 on the wall of the divergent 3, and is not caused by the realization of the condition expressed by equation [3] (or equivalent condition). Of course, the determination of said optimum position includes taking into account a sufficient safety margin, which is obtained, for example, by decreasing by 20% the value of x_sep obtained by applying equation [3].

  The flow topology that has been described with reference to FIG. 1 is known as free shock wave separation. Under certain conditions, depending on the geometry of the nozzle and the pressures involved, the flow separation induced by the shock wave 8 can not develop along the wall of the divergent, because the pressure gradient causes the immediate re-attachment of the boundary layer by creating a toroidal separation bulb on said wall. Under these conditions, a large recirculation zone is created in the central part of the jet downstream of the rounded body: the topology of the flow then becomes very similar to that described in the aforementioned article by G. Hagemann et al. under the name of restricted shock separation, also called separation by internal recirculation in the jet. This flow regime is shown diagrammatically in FIG. 3, in which the same reference numerals correspond to the same elements as in FIG. 3, the reference 16 denotes the toroidal separation bulb and the reference 17 the recirculation zone in FIG. the central part of the jet.

  As in the case of the free shock wave separation shown in FIG. 1, the detached shock wave 8 produces a stable separation bulb in a position upstream of that at which separation would occur spontaneously. The invention therefore retains all its usefulness under these conditions.

  For particular geometries of the nozzle, and in particular in the case of the Vulcain engine, the separation of the flow can be either of free shock wave type, or of type by internal recirculation in the jet, depending on the ratio between the pressure of the stopping in the combustion chamber and the ambient pressure. As during the start phase the stop pressure increases from the value of the atmospheric pressure to the nominal operating value, the type of separation varies over time. In the case of the Vulcain engine, for example, there is a transition from a free shock wave separation regime to a separation by internal recirculation in the jet, followed by a return to free shock wave separation. These changes of regime induce much more non-stationary lateral forces than in the case of separation only by free shock wave or only by internal recirculation in the jet, and are therefore undesirable. However, the presence of the rounded body 5 tends to favor separation by internal recirculation in the jet relative to the free shock wave separation, because its drag creates by itself a central recirculation zone. As a result, the transitions from one mode to the other are suppressed, the starting of the engine being entirely in 1 o separation conditions by internal recirculation in the jet. The method of the invention therefore has a particularly beneficial effect for this type of nozzle.

  The flowchart of FIG. 2 illustrates a rational method of determining the setpoint x (ps) of the position of the body 5.

  First, step E1, a mathematical model of the rocket engine makes it possible to determine, as a function of the stop pressure ps, the temperature Ts in the combustion chamber and the exponent of isentropic expansion y of the combustion gases (in fact, it It must generally be taken into account that the composition of said gases, and therefore the value of the parameter y, varies during the start-up phase). This model results from the observation of the average operating parameters at the start of the engine during its bench qualification tests. Such tests are always performed before using a motor on a launcher, and are therefore not specific to the present invention.

  Secondly, the step E2 comprises the determination of the maximum value reached by the stopping pressure of the gases in the combustion chamber 1, psmax the minimum pressure being equal to the ambient pressure at the launch altitude pa, and the choice of a series of N discrete values pal, ps "lying between psmax and pa, the number N, for example 10, being a compromise between the desired fineness of the modeling and the accepted calculation volume.

  Thirdly, the step E3 comprises the choice of a series of M discrete values x ', ..., xM of the position of the rounded body 5, between the position of the neck 2 of the nozzle and that of its section output 4, the number M, for example 10, being a compromise between the desired fineness of the modeling and the accepted calculation volume.

  Then, step E4 comprises, for each pair of values (psi, x ') the determination, in particular by means of numerical simulations and / or by bench tests, of the pressure p (x) and the Mach number. M (x) of the flow at any point of the nozzle, as well as the point of incidence 9 of the shock wave 8 on the walls of the divergent 3, indicated by xchoc1. The simulations are performed taking into account the values of the temperature Ts 1 and the exponent y of the combustion gases determined in step E1.

  The effect of a non-zero ambient pressure pa is taken into account in step E5. During this step, it is checked whether the ambient pressure Pa causes spontaneous delamination of the boundary layer (case of FIG. 1) or separation by internal recirculation in the jet (case of FIG. 3) before the interaction of the boundary layer with the shock causes this separation or separation by internal recirculation (in the following the term separation will be used with reference to the two flow regimes). This verification is carried out either by the bench test or, for example, using the criterion of equation [3] or any other empirical criterion considered to be better adapted, the criterion being applied to the isentropic zone of the flow ( if the criterion can not be satisfied in the isentropic zone, the separation will be generated by the shock, provided that the position of the latter is sufficiently far back to interact with a weakened boundary layer, but only the most distant positions of the shock are interesting. in the following application of this process). We therefore deduce the position of the separation characterized by x_sep. The indices i and j recall that this is the value of x_sep determined for a stop pressure psi and a position x 'of the body 5. In fact, as suggested above, it is preferable not to not use for x_sep '' the value directly obtained by a criterion such as that of equation [3], but correct it by the use of a safety factor (for example, a reduction of 20%).

  In step E6, x_sep't is compared to xhhoc '(position of the point of incidence 9 of the shock wave 8 on the walls of the divergent 3). The optimum value of the position of the body 5 at the stop pressure ps ', xppt (ps'), is defined as the most downstream position of said body 5 such that x_sep '' is not less than xchoc '' . When this condition is verified, the point of incidence of the shock wave 8 on the wall of the nozzle sets the jet separation. this means that the optimum position of the body 5 for a given value of the stopping pressure ps is the most downstream position such that the separation of the jet is caused by the shock wave 8 induced by said body 5, and not by the effect of the ambient pressure pa. In other words, the optimal position of the body 5 for a given value of the stop pressure ps is the position such that the shock wave 8 is incident on the wall of said divergent at an axial position 9 corresponding to the downstream limit of the region. the divergent where does not occur a spontaneous separation of jet under the action of the ambient pressure. It is clear that, as a finite number (M) of values of the position of the body 5 has been considered, the optimal position is only defined in an approximate manner.

  In this way, a relation xopt (ps) has been determined for N values of pressure ps; this relation gives the position of the position of the body 5: x (ps) = xopt (ps). However, since this determination is made at the expense of relatively complex calculations, the value of N is necessarily limited: it is sufficient to consider that, if N = M = 10, N × M = 100 numerical simulations or tests must be carried out at step E4 , at least in principle. It is therefore advantageous to provide a further interpolation step E7, allowing for example to obtain the setpoint x (ps) in an analytical form.

  When sufficiently powerful calculation means are available, the steps E4 and E5 of the method of FIG. 2 can be grouped together by involving a numerical solution of the Navier-Stokes equations integrating from the start the effect of the atmospheric pressure and making it possible to predict the position of the flow separation zone and its type (free or restricted shock wave) without the need for empirical criteria.

  A method according to one embodiment of the invention to cause a stable jet separation inside a nozzle of a rocket engine therefore comprises the following steps: - prior to the start of the ignition phase of the engine inserting the rounded body 5 inside the diverging nozzle 3 along its axis to a first axial position; this first axial position is calculated to correspond to the optimal position of said body (in the direction defined above) when the stop pressure ps reaches a sufficiently high value that a jet separation can occur whose effects the mechanical forces generated at the neck of the nozzle at the very beginning of start-up are not structurally dimensioning as long as the separation zone remains close to the neck, as has been explained above. ); during the ignition phase, the displacement of said rounded body along the axis of the nozzle, in a direction generally from the neck 2 to the outlet section 4, so as to constantly maintain said body in a approximately optimal position.

  If the engine nozzle is dimensioned so that when the stopping pressure ps of steady state is reached, the re-compression of the gases is entirely effected outside the divergent, the problem of the jet separation does not occur. arises only during the start-up phase. At the end of this phase, the body 5 will have reached a position so downstream that it will no longer have any influence on the flow in the nozzle, and can therefore be completely retracted. If, on the other hand, the operating conditions under regime are such that a jet separation can occur even after the end of the start-up phase, it may be necessary to combine the invention with one of the known devices of the invention. prior art to suppress non-stationary lateral forces also during the first part of the rocket's ascent phase.

  A variant of the embodiment described above allows the device according to the invention to operate autonomously, without the need to receive pressure information from the combustion chamber. Indeed, even if a pressure sensor 12 is generally provided in most modern launchers, the transmission of information to the controller 15, via a launcher-ground connector or telemetry, can cause reliability problems . If these problems are considered particularly important, it is possible to replace the direct measurement of the pressure ps in the combustion chamber by a psb pressure measurement at the point of stagnation of the body 5, that is to say at the top of the latter . The pressure psb is connected to the pressure ps in the combustion chamber by a simple relation of proportionality: Ps = 2y M2 y 1 r 'x Psb y + 1 y + 1 r 2 x 1 + y 1 r-' y + 1 M2 y + 1 [4] where M is the Mach number upstream of the detached shock wave 8, calculated by the numerical methods of fluid mechanics. This embodiment of the invention requires a slight increase in complexity and cost of the rounded body 5, which must be provided on its top with a pressure sensor (reference 12 'in FIG. measure of psb.

  The embodiments of the invention which have just been described make it possible to eliminate the non-stationary lateral forces during the entire starting phase of the engine. However, these forces are particularly harmful when they apply to the region of the divergent close to the opening, ie towards the end of the start: in some applications it will be sufficient to suppress them only during this final phase, which makes it possible to implement a simplified embodiment of the invention, not requiring the use of an actuator controlled by an actuator to retract the body of rounded shape. If the method of the invention is to be implemented only during the terminal phase of the start, the rounded body 5 will be initially placed at an axial position such that the impact point of the shock 8 on the divergent will be relatively close to the outlet section 4 of the nozzle, and its displacement will take place over a relatively small range of axial positions, corresponding to a restricted range of values of the stop pressure ps in the combustion chamber. Under these conditions, the setpoint x (ps) can reasonably be linearized. The axial force exerted by the flow of combustion gases on the body of rounded shape and tending to expel the divergent nozzle is proportional to Psb and, as shown in equation [4] above, Psb is proportional to ps and depends on M which itself depends on the position x of the body 5 and the shape of the nozzle. In a limited domain of position variation of the body 5, it is possible to approximate this last implicit relation by a linear relationship between Psb and x. It follows that the relationship between the set position of the body 5 and the axial aerodynamic force applied to this body can be approximated by a linear relationship. This approximate linear dependence of the setpoint position x of the body 5 at the pressure psb can then be obtained by means of a simple means of applying an elastic force opposing the expulsion of said body 5. This is what is shown in Figure 5, where the controller 15 and the actuator 6 are replaced by the spring 19. It is easily understood that this simple assembly provides a displacement of the rounded body 5 which is proportional to the axial force which is exercised on the latter. The choice of a spring having appropriate stiffness and prestressing makes it possible to comply with the linearized setpoint law. The spring 19 can be of any type adapted to the specific application considered: helical, pneumatic, etc. The embodiment described above is not only much simpler than those described with reference to Figures 1 and 3, it is also more reliable because fully passive.

  Another embodiment of the invention pushes the simplification even further: in certain applications, the range of displacement required for a satisfactory suppression of the non-stationary lateral forces is sufficiently narrow to be considered punctual: in this case, it is sufficient to arranging the rounded body 5 at the desired location using a suitable means such as an actuator, without the need to move it during the ignition phase of the rocket engine. The fixed position of the body 5 is determined using the method described with reference to FIG.

Claims (23)

  A method for reducing non-stationary lateral forces acting on a nozzle of a rocket motor during a starting phase of said engine, said engine comprising a combustion chamber (1) where combustion gases are generated, a divergent (3) in which a supersonic flow of said combustion gases occurs and a neck (2) connecting said combustion chamber to said divergent, characterized in that it comprises the positioning of a body (5) of rounded shape to the interior of the divergent portion (3) along its axis in correspondence of an axial position of the rounded body such that, during at least a portion of said starting phase, a shock wave (8), induced by the disturbance of the flow of the combustion gases caused by said rounded body (5) is incident on the wall of said divergent (3) at an axial position of incidence where it produces a jet separation or a bulb-shaped separation of s toroidal repair.   2. Method according to claim 1 comprising: - prior to starting the engine, the insertion of said body (5) of rounded shape inside the divergent (3) along its axis, to a first axial position; and during the start-up phase, the displacement of said rounded body (5) along the axis of the nozzle as a function of a value of the stop pressure (ps) of the combustion gases in the chamber of combustion (1) such that, while the value of said stopping pressure (ps) of the combustion gases varies during said starting phase, said shock wave (5) continues to be incident on the wall of the diverge at an axial position where it produces a jet separation or a toroidal separation bulb shaped separation.   3. Method according to claim 2 wherein the displacement of said body (5) of rounded shape along the axis of the nozzle according to a value of the stop pressure (ps) of the combustion gases in the chamber of combustion (1) is carried out in such a way that said shock wave (8) is incident on the wall of said divergent at an axial position (9) corresponding to the downstream limit of the region of said divergent where spontaneous separation does not occur of a jet or a spontaneous separation in the form of a toroidal separation bulb under the action of ambient pressure.   4. Method according to one of the preceding claims wherein the displacement of said body (5) of rounded shape as a function of the stop pressure (ps) of the combustion gases in the combustion chamber (1) follows a setpoint which is determined by the following steps: -selection of a series of discrete values of the stopping pressure (ps) of the combustion gases in the combustion chamber (1), between atmospheric pressure and a maximum value reached during the start-up phase (E2); - choosing a series of discrete values of the position of said body (5) of rounded shape along the axis of the divergent, between the position of the neck (2) and that of the outlet section (4) of said divergent (E3); for each pair of said discrete values, determination by calculation or by testing of the static pressure value and the Mach number along the wall of the divergent (E4) and determination of the impact point of the impact on the wall the nozzle; for each pair of said discrete values, determining the axial position of the jet separation point or the toroidal separation bulb-shaped separation point using said static pressure values and the Mach number of the gas flow combustion along the wall of the divergent (E5); for each of said discrete pressure values (ps) of the combustion gases in the combustion chamber (1), determining the position of said body (5) of the most downstream rounded shape, such as the jet separation or the toroidal separation bulb shaped separation is caused by said shock wave (8) induced by the presence of said body (5) of rounded shape (E6); said most downstream value being selected as a set value of the position of said rounded body (5) in correspondence of said value of the stopping pressure (ps) of the combustion gases in the combustion chamber (1).   5. The method of claim 4 also comprising an interpolation operation of said setpoint values of the position of said body (5) of rounded shape in correspondence of said values of the stop pressure (ps) of the combustion gases in the chamber of combustion (1) to determine a position setpoint in an analytical form (E7).   The method of claim 4 or 5 wherein said axial position of the spontaneous jet separation point (9) or the spontaneous separation in the form of a toroidal separation bulb under the action of ambient pressure is determined using an adapted empirical or semiempirical criterion.   7. Method according to one of the preceding claims wherein said body (5) of rounded shape is moved from said axial position to the outlet section (4) of the divergent during the starting phase of the engine as and when the stopping pressure (ps) of the combustion gases in the combustion chamber (1) increases.   The method of one of the preceding claims wherein said rounded body (5) is axially symmetrical and has a rounded front surface facing the neck (2) of the nozzle.   9. The method of claim 8 wherein said body (5) of rounded shape has a cross section of between 0.5 and 2 times, and preferably between 0.8 and 1.5 times, the section of the neck (2). of the nozzle.   Method according to one of claims 2 to 9, wherein said value of the stopping pressure (ps) of the combustion gases in the combustion chamber (1) is determined indirectly from a measurement of the pressure of stopping (Psb) said combustion gases in correspondence of an apex of said rounded body.   Apparatus for reducing non-stationary lateral forces acting on a nozzle of a rocket motor during a starting phase of said engine, said engine comprising a combustion chamber (1) where combustion gases are generated, a divergent (3) in which a supersonic flow of said combustion gases occurs and a neck (2) connecting said combustion chamber to said divergent, characterized in that it comprises: - a body (5) of rounded shape intended to be positioned at the interior of the divergent (3) along its axis; and means for positioning said rounded body (5) inside the divergent portion (3) in correspondence of an axial position of the rounded body such that, during at least a portion of said starting phase, a shock wave (8), induced by the disturbance of the flow of combustion gases caused by said body (5) of rounded shape, is incident on the wall of said divergent (3) at an axial position of incidence where it produces a jet separation or a separation in the form of a toroidal separation bulb.   12. Device according to claim 11 also comprising means (6, 19) for moving said body (5) of rounded shape along the axis of the divergent (3) according to a value of the stop pressure ( ps) of combustion gases in the combustion chamber (1) such that, while the value of said stop pressure (ps) of the combustion gases varies during said start-up phase, said shock wave ( 5) continues to be incident on the divergent wall at an axial position where it produces a jet separation or a toroidal separation bulb form separation.   13. Device according to claim 12 wherein said means (6, 19) for moving said body (5) rounded along the axis of the divergent (3) comprises an actuator (6) for moving said body of rounded shape along the axis of the divergent (3).   14. Device according to claim 13 also comprising a controller (15) for receiving from a first sensor (12, 12 ') information of value of the stop pressure (ps) of the combustion gases in the combustion chamber ( 1) and controlling said actuator to move said rounded body (5) along the divergent axis (3) according to said value of the stopping pressure (ps) of the combustion gases in the chamber of combustion (1).   15. Device according to claim 14 wherein said controller (5) is a controller for controlling the displacement of said body (5) of rounded shape along the axis of the divergent (3) of the nozzle according to a set position instruction to using the following steps: - choosing a series of discrete values of the stopping pressure (ps) of the combustion gases in the combustion chamber (1), between atmospheric pressure and a maximum value reached during the start-up phase (E2); - choosing a series of discrete values of the position of said body (5) of rounded shape along the axis of the divergent (3), between the position of the neck (2) and that of the outlet section (4) ) of said divergent (E3); for each pair of said discrete values, determining by calculation or by testing the static pressure value and the Mach number along the divergent wall (E4), and determining the point of impact impact on the nozzle wall; for each pair of said discrete values, determining the axial position of the jet separation point using said static pressure values and the Mach number along the divergent wall (E5); for each of said discrete values of the stopping pressure (ps) of the combustion gases in the combustion chamber (1), determining the position of said body (5) of rounded shape furthest downstream, such as the separation of jet or toroidal separation bulb shaped separation is caused by said shock wave (8) induced by the presence of said body (5) of rounded shape (E6); said most downstream value being selected as a set value of the position of said rounded body (5) in correspondence of said value of the stopping pressure (ps) of the combustion gases in the combustion chamber.   16. Device according to claim 15 wherein said position setpoint has an analytical form, determined by means of an additional operation (E7) of interpolation of said setpoint values of the position of said body (5) of rounded shape. matching said values of the stopping pressure (ps) of the combustion gases in the combustion chamber (1).   17. Device according to one of claims 15 or 16 wherein said axial position of the spontaneous jet separation point or the spontaneous separation point in the form of a toroidal separation bulb under the action of the ambient pressure is determined by using a suitable empirical or semi-empirical criterion, 18. Device according to one of claims 14 to 17 wherein said controller (15) is a controller for controlling the displacement of said body (5) of rounded shape along the axis of the divergent (3) of the nozzle of said axial position towards the outlet section (4) of the divergent portion (3) during the starting phase of the engine as the stopping pressure (ps) of the combustion gases in the combustion chamber (1) increases.   19. Device according to one of claims 11 to 18 wherein the body (5) of rounded shape has a concave surface having an apex intended to be oriented towards the neck (2) of the nozzle and a pressure sensor (12 ' ) arranged in correspondence of said vertex.   20. Device according to claim 12 wherein said means for moving the body (5) of rounded shape along the axis of the divergent (3) according to a value of the stop pressure (ps) of the gases of combustion in the combustion chamber (1) comprises means (19) for applying an elastic force opposing the expulsion of said body (5) of rounded shape of said divergent (3) by said combustion gases.   21. Device according to one of claims 11 to 20 wherein said body (5) of rounded shape has an axial symmetry and has a rounded surface intended to be oriented towards the neck (2) of the nozzle.   22. Device according to one of claims 11 to 21 wherein said body (5) of rounded shape has a cross section of between 0.5 and 2 times, and preferably between 0.8 and 1.5 times, the section neck (2) of the nozzle.   23. Device according to one of claims 11 to 22 comprising a mechanical fuse for the expulsion of the body (5) of rounded shape of the divergent (3) when an axial force exerted on said body (5) exceeds a predetermined threshold.