Gas-dynamic nozzle of a rocket engine, in particular of a launcher engine

Abstract

The present invention relates to a gas-dynamic nozzle (1) of a rocket engine, in particular of a launcher, wherein said nozzle (1) comprises a converging portion (10), a throat portion (20), and a diverging portion (30). The main feature of the present invention lies in the fact that at least said throat portion (20) and at least said diverging portion (30) comprise an inner panel (41) and an outer panel (42), wherein said inner panel (41) and outer panel (42) define an interspace (43) in between, configured to allow a flow of coolant fluid for cooling said throat portion (20) and said diverging portion (30), and wherein said interspace (43) comprises an open-cell structure (44), in particular said structure (44) being made by additive manufacturing. Furthermore, said nozzle (1) comprises at least one optical-fibre sensor (50) housed in said interspace (43) and configured to monitor at least one physical variable of the nozzle (1) and / or of the coolant fluid flowing in said interspace (43), in particular said at least one physical variable being related to the temperature, pressure, vibration and deformation of the nozzle (1) and / or being related to the temperature and pressure of said coolant fluid.

Description

[0001] GAS-DYNAMIC NOZZLE OF A ROCKET ENGINE, IN PARTICULAR OF A LAUNCHER ENGINE

[0002] DESCRIPTION

[0003] The present invention relates to a gas-dynamic nozzle of a rocket engine, in particular of a launcher.

[0004] The present invention finds its main application in the aerospace industry, where it is known that a rocket engine is a jet engine which exploits the principle of action and reaction to generate a thrust force, ejecting gas at high speed.

[0005] In this context, space access engines are known in the art (also referred to as “launch vehicles” or “launchers”), which can launch a payload into orbit; in particular, modern launchers include engines running on liquid fuel / oxidizer, or engines running on solid fuel, or hybrid engines running on solid / liquid fuel. Based on whether the unit can be reused or not, the following launcher types exist:

[0006] - expendable launchers, which are not recovered for future missions (e.g. most modern launchers of the Ariane families);

[0007] - reusable launchers, which can be reused for future missions (e.g. the solid-fuel acceleration boosters of the Space Shuttle, the casing of which is recovered for a number of missions; the American shuttle called “orbiter”, which is a real launcher, although not fully recoverable; or the first stage of the Falcon 9 and Falcon Heavy rockets produced by American company SpaceX).

[0008] A solid-fuel launcher typically comprises two separate tanks containing fuel and oxidizer; both components are then delivered, usually by means of pumps, to a combustion chamber, the products of which then flow through an exhaust nozzle.

[0009] A solid-fuel launcher comprises a casing filled with propellant and equipped with a gasdynamic nozzle in the discharge section, where the gases produced by propellant combustion flow out.

[0010] Therefore, all launchers (whether running on solid, liquid, or hybrid fuel) comprise a gasdynamic nozzle having a converging-diverging shape, also known as “de Laval nozzle”, which accelerates the outflow of the exhaust gases to produce kinetic energy. In this context, a converging-diverging shape of the gas-dynamic nozzle allows the exhaust gases to be expelled into the atmosphere at a speed which is higher than the speed at which the rocket moves relative to the air, typically at supersonic speed. In particular, the gas-dynamic nozzle of a liquid-fuel launcher operates in the supersonic configuration to maximize the gas output velocity, with a sonic speed in the smallest nozzle section (also referred to as throat section), and has a parabolic divergent section (also referred to as “bell-shaped section” or “parabolic section”), while a gas-dynamic nozzle of a solid-fuel launcher typically has a parabolic conical divergent section (also referred to as “conical section”) due to the presence of expanding solid particulate.

[0011] However, the launchers known in the art still have a few drawbacks.

[0012] In fact, because of the high temperatures reached in the combustion chamber and in the gas-dynamic nozzle, thermal control systems need to be used in order to protect the walls from high thermal flows of burning or burned gases.

[0013] The following thermal protection systems are currently known in the art:

[0014] - regenerative cooling systems, which comprise a complex system of ducts (also known as “tube bundle system”) in which a coolant (typically fuel, e.g. hydrogen) flows, to which thermal power is transferred. The coolant in the complex network of ducts may either follow the flow of the combustion gases or flow in the opposite direction. While it has been adopted in the main engines of the Space Shuttle (SSMEs) and also in the Vulcain engine of Ariane 5, this solution is, disadvantageous^, very expensive, in that it requires accurate periodic checks of the functionality and integrity of each duct included in the tube bundle and of each connection (welded or brazed joints, etc.) to the nozzle wall. Moreover, the addition of the tube bundle inevitably increases the weight of the launcher and, thus, reduces the payload for the same amount of embarked fuel;

[0015] - liquid-film or transpiration cooling systems, in which layers of liquid coolant, typically consisting of the same fuel, flow within the convergent and divergent parts of the nozzle, thereby protecting the wall. In transpiration systems, a porous wall lets in coolant droplets that protect the wall and absorb the high thermal flows. Such a solution has the drawbacks that it does not provide accurate control over the amount of liquid coolant used for protecting the wall, and that a fraction of fuel is lost which is not burnt in the chamber, thus increasing the enthalpy of the ejected gases;

[0016] - ablative cooling systems, in which ablative material undergoes a strong endothermal reaction in contact with high thermal gradients, thus degrading itself and producing a carbon residue. This technique is not stationary, and the cooling system is not reusable;

[0017] - thermal-well cooling systems, in which a material having high thermal conductivity transports the absorbed heat towards other areas. This solution has a major drawback in that, since the melting temperature of the material having high thermal conductivity must be high, it necessarily requires the use of metals with a high melting point;

[0018] - radiative cooling systems, in which the high thermal emissivity of some materials provides a cooling effect through irradiation of other areas, which, however, are in turn at risk of overheating and jeopardizing the proper operation of the launcher.

[0019] A further drawback of the launchers currently known in the art is the absence of an active cooling control system allowing some physical variables of the nozzle, in particular in terms of temperature, pressure, vibration and deformation, to be constantly monitored to ensure effective and continuous control over the operating state and structural condition of the nozzle.

[0020] It is therefore apparent that the launchers known in the art are not in a condition to provide effective diagnostics and / or predictive maintenance of the state or condition of the nozzle, nor to provide a history of the mechanical stresses it has undergone, and it is clear that this problem is especially important for reusable launchers.

[0021] Document US3116603 A concerns a nozzle for a launcher including a cooling interspace through which a cooling gas is made to flow, which is then discharged in the terminal part through suitable nozzles to control the trajectory of said launcher.

[0022] However, the solution shown in document US3116603 A suffers from the drawback that it increases the complexity and weight of the launcher. Moreover, the solution shown in document US3116603 A is not adjustable in flight and is not feedback-controllable (in particular, by means of a “closed-loop” system).

[0023] Document CN116398320A relates to a thrust vectoring system equipped with a plurality of pipes and a control system acting upon a vortex valve. However, the solution shown in document CN116398320A has the drawback that the secondary gas is only used for controlling the thrust, and not for cooling the primary nozzle as well. Furthermore, this solution has no integrated sensors capable of monitoring both the temperature of the main nozzle and the effectiveness of the cooling and thrust control equipment.

[0024] In this frame, it is therefore the main object of the present invention to provide a gasdynamic nozzle of a rocket engine, in particular of a launcher, so conceived as to overcome the drawbacks of the solutions currently known in the art.

[0025] In particular, it is one object of the present invention to provide a gas-dynamic nozzle so conceived as to contribute to the overall performance of the launcher, without being excessively costly and without requiring time-consuming maintenance.

[0026] It is another object of the present invention to provide a gas-dynamic nozzle which comprises an integrated cooling system and which can be reused with minimal maintenance.

[0027] It is a further object of the present invention to provide a gas-dynamic nozzle so conceived that it can be manufactured without necessarily requiring the use of materials with a high melting point.

[0028] It is a further object of the present invention to provide a gas-dynamic nozzle so constructed as to ensure considerable weight savings as well as lower costs (in particular due to the use of non-high-melting materials) in comparison with prior-art solutions.

[0029] It is a further object of the present invention to provide a gas-dynamic nozzle so constructed as to allow for predictive maintenance of the state or condition of the nozzle itself, thus being particularly suitable for implementation in reusable launchers.

[0030] It is yet another object of the present invention to provide a gas-dynamic nozzle so constructed as to provide control over the thrust direction of the rocket engine or launcher. Further objects, features and advantages of the present invention will become apparent in light of the following detailed description and of the annexed drawings, which are provided herein merely by way of non-limiting explanatory example, wherein:

[0031] - Fig. la is a front view of a first embodiment of a gas-dynamic nozzle of a rocket engine, in particular of a launcher, according to the present invention;

[0032] - Fig. lb is a longitudinal sectional view of the first embodiment of the nozzle according to the present invention, i.e. a sectional view along a plane A-A shown in Fig. la;

[0033] - Fig. 2a is a longitudinal sectional view of a second embodiment of the nozzle according to the present invention;

[0034] - Fig. 2b is a magnified view of detail A of Fig. 2a;

[0035] - Fig. 2c is a magnified view of detail B of Fig. 2a;

[0036] - Fig. 2d is a plan view (from below) of the second embodiment of the nozzle according to the present invention;

[0037] - Fig. 3a is a longitudinal sectional view of a third embodiment of the nozzle according to the present invention;

[0038] - Fig. 3b is a magnified view of detail C of Fig. 3a;

[0039] - Fig. 3c is a magnified view of detail D of Fig. 3a;

[0040] - Fig. 3d is a plan view (from below) of the third embodiment of the nozzle according to the present invention.

[0041] Describing now the annexed figures, reference numeral 1 designates as a whole a gasdynamic nozzle (hereafter referred to as “nozzle 1” for simplicity) of a rocket engine, in particular a launcher, according to the present invention; in this regard, it should be noted that said rocket engine or launcher is not shown in the accompanying drawings, since it is of a known type and does not need to be shown to illustrate the features of the present invention.

[0042] The nozzle 1 is of the type having a converging-diverging shape, also known as “de Laval nozzle”.

[0043] In this context, the nozzle 1 comprises, in a per se known manner, a converging portion (which may also be defined as “convergent section”, identified as a whole by reference numeral 10), a throat portion (or “throat section”, identified as a whole by reference numeral 20), and a diverging portion (or “divergent section”, identified as a whole by reference numeral 30). As is known, the nozzle 1 schematically consists of a tube having a restricted central portion (i.e. the throat portion 20) and is similar to an asymmetric hourglass.

[0044] In accordance with the present invention, at least said throat portion 20 and at least said diverging portion 30 comprise an inner panel 41 and an outer panel 42, wherein said inner panel 41 and outer panel 42 define an interspace 43 in between, configured to allow a flow of coolant fluid for cooling said throat portion 20 and said diverging portion 30.

[0045] As shown in the annexed drawings, said diverging portion 30 has a bottom edge 31 for closing said interspace 43.

[0046] Preferably, said interspace 43 comprises an open-cell structure 44 (which may also be defined as “open-cell trabecular structure 44” or “cellular core”) allowing said coolant fluid to flow within said interspace 43 in order to cool the nozzle 1.

[0047] Preferably, said structure 44 is made by additive manufacturing.

[0048] It is therefore clear that the solution proposed by the present invention, which employs a “cellular core” instead of a tube bundle (like, for example, the one shown in document US3116603 A), incorporates a cooling system into the structure of the nozzle 1 without increasing its weight, and provides coolant drain channels in one component only, resulting in simpler assembly operations and improved performance.

[0049] Furthermore, the nozzle 1 according to the present invention comprises at least one optical-fibre sensor 50 housed in said interspace 43 and configured to continuously monitor at least one specific physical variable of the nozzle 1 and / or of the coolant fluid flowing in said interspace 43, in particular said at least one physical variable being related to the temperature, pressure, vibration and deformation of the nozzle 1 and / or being related to the temperature and pressure of said coolant fluid. In this respect it should be noted that, in addition to a “raw” parameter of said at least one physical variable, it is also possible to define, based on the previously obtained physical measurement, control algorithms / smart logics / information about the behaviour of the entire system (thereby obtaining, for example, real-time attitude information as a function of the measured fuel flow / consumption, etc.)

[0050] Preferably, said at least one sensor 50 comprises (or consists of) at least one fiber Bragg grating (FBG) sensor.

[0051] The nozzle 1 is preferably associated with an interrogator (not shown in the annexed drawings) connected to said at least one sensor 50 and connected to a control unit (e.g. consisting of an electronic board, also not shown in the annexed drawings), preferably positioned at a considerable distance from said at least one sensor 50 (and, hence, from the nozzle 1). It is therefore apparent that said at least one sensor 50, interrogator and control unit belong to (or constitute) an active control system for controlling the cooling of the nozzle 1 according to the present invention, wherein said active control system is configured to continuously monitor some physical variables of the nozzle 1, in particular in terms of temperature, pressure, vibration and deformation; this also provides effective and constant control over the state or condition of the nozzle 1, as well as effective diagnostics and / or predictive maintenance of the nozzle 1, while also making it possible to reconstruct a history of undergone mechanical stresses.

[0052] Said at least one sensor 50 may be positioned in contact with the inner panel 41, i.e. in contact with a side of the inner panel 41 delimiting the interspace 43 (i.e. that side of the inner panel 41 which faces the interspace 43). As an alternative, said at least one sensor 50 may be positioned within the structure 44, which is preferably made by additive manufacturing (as previously explained); in this context, the fact that the structure 44 is made by additive manufacturing facilitates the positioning of said at least one sensor 50 within the structure 44, in particular said additive manufacturing making it possible to create at least one path in which the optical fiber of said at least one sensor 50 can be inserted (or incorporated during the additive manufacturing process).

[0053] As far as said at least one sensor 50 is concerned, it should be noted that it is shown in the accompanying figures as a single continuous line, also considering the negligible thickness of the optical fiber and of the fiber Bragg gratings; it is however clear that the nozzle 1 according to the present invention may be so constructed as to comprise a plurality of sensors 50 positioned at different points within the interspace 43. For example, the nozzle 1 according to the present invention may comprise a plurality of sensors 50 arranged in series with one another and configured to operate at different wavelengths to detect different wavelengths of variation of the measured thermomechanical quantities, and hence so configured as to detect different physical variables or parameters of the nozzle 1 and / or of the coolant fluid.

[0054] Figures la and lb concern a first embodiment of the nozzle 1, wherein said nozzle 1 is comprised in a rocket engine or launcher running on liquid fuel.

[0055] In the first embodiment, also the converging portion 10 of the nozzle 1 comprises said inner panel 41 and outer panel 42 that define the interspace 43.

[0056] In this context, the outer panel 42 of the nozzle 1 comprises at least one first aperture 42A for the entry of the coolant fluid into the interspace 43 and at least one second aperture 42B for the exit of said coolant fluid from said interspace 43. As a consequence, said at least one first aperture 42A and said at least one second aperture 42B are respectively connected (e.g. by means of pipes not shown in the accompanying drawings) to a tank and a combustion chamber of the launcher. Note that each one of said apertures 42A, 42B may be provided with a respective valve, not shown in the accompanying drawings.

[0057] Preferably, in the first embodiment the coolant fluid is the fuel of the launcher, which enters the interspace 43 in liquid form (through said at least one first aperture 42A), and then exits said interspace 43 (through said at least one second aperture 42B) in gaseous form due to the heat exchange undergone by the fuel as it flows within the interspace 43. In a preferred embodiment (like the one shown in Figures la and lb), said at least one first aperture 42A is located in proximity to a bottom edge 31 of the nozzle 1, and said at least one second aperture 42B is located in proximity to a top edge 32 of the nozzle 1; it is therefore evident that, in this preferred embodiment, the coolant fluid flows within the interspace 43 in the direction opposite to the flow of the combustion gases. It is however clear that the nozzle 1 may also be constructed differently: for example, said at least one first aperture 42A may be located in proximity to the top edge 32, and said at least one second aperture 42B may be located in proximity to the bottom edge 31. In addition, the nozzle 1 according to the present invention may be so constructed as to comprise a plurality of first apertures 42A, in particular distributed around the outer panel 42 and in proximity to the bottom edge 31, and / or a plurality of second apertures 42B, in particular distributed around the outer panel 42 and in proximity to the top edge 32.

[0058] It should also be noted that, in accordance with the present invention, the nozzle 1 of an rocket engine or launcher running on liquid fuel (like the one shown in Figures la and lb) may also be conceived in such a way that the coolant fluid is not the fuel of the rocket engine; therefore, the nozzle 1 of a rocket engine or launcher running on liquid fuel (like the one shown in Figures la and lb) may also be constructed as described below (and shown in Figures 2a to 3d) with reference to a nozzle 1 of a rocket engine or launcher running on solid fuel.

[0059] As shown in Figures la and lb, in the first embodiment of the nozzle 1 the interspace 43 is essentially defined by the inner panel 41, the outer panel 42, the bottom edge 31 and the top edge 32.

[0060] Figures 2a to 2d show a second embodiment of the nozzle 1, in particular said nozzle 1 belonging to a launcher running on solid fuel (identified by reference CS in Fig. 2a and positioned in a casing 2).

[0061] In the second embodiment, shown in Figures 2a to 2d, the nozzle 1 according to the present invention comprises a cooling system (identified as a whole by reference SR in Figures 2a and 2b) comprising a tank 60 containing a coolant fluid (which may be either in gaseous form or in liquid form, such as liquid carbon dioxide), wherein said tank 60 is positioned within the interspace 43, in particular in a position corresponding to the throat portion 20 of the nozzle 1. Said tank 60 may be integrated into the nozzle 1 during the construction of the latter by additive manufacturing.

[0062] As clearly shown in Figures 2a and 2b, said tank 60 is located in a region between the converging portion 10 and the throat portion 20 of the nozzle 1. In this regard, the inner panel 41 and the outer panel 42 may be configured or shaped in such a way that they join each other upstream (in particular, with reference to the direction of the exhaust gases) of the tank 60, i.e. said inner panel 41 and outer panel 42 comprise a union tract 40 (shown in Fig. 2b and Fig. 3b) which is common to both the inner panel 41 and the outer panel 42 and which is located upstream of the tank 60 (still with reference to the direction of the exhaust gases); as a result, the interspace 43 is defined by the inner panel 41, the outer panel 42, the bottom edge 31, and the union tract 40 of said panels 41, 42 upstream of the tank 60.

[0063] In an alternative embodiment (not shown in the accompanying drawings), the tank 60 may be so configured as to constitute an element for closing the interspace 43; as a result, in such alternative embodiment the interspace 43 is essentially defined by the inner panel 41, the outer panel 42, the bottom edge 31, and the tank 60.

[0064] Moreover, said tank 60 is toroidal in shape, i.e. substantially annular and internally hollow (that is to say, said tank 60 is internally empty or hollow to be able to contain the coolant fluid).

[0065] In accordance with the present invention, the tank 60 comprises injection means 61 (e.g. comprising a plurality of valves or injectors) configured to inject the coolant fluid into the interspace 43.

[0066] The cooling system SR of the nozzle 1 further comprises draining means (identified as a whole by reference numeral 62) for draining the coolant fluid from the interspace 43.

[0067] In particular, said draining means 62 may comprise a plurality of passages 62A preferably formed in the bottom edge 31 of the diverging portion 30 (as shown in Figures 2c and 2d).

[0068] It is therefore apparent that, in the embodiment shown in Figures 2a to 2d, the coolant fluid flows in the interspace 43 in the same direction as the exhaust gases of the nozzle 1. In this context, the draining means 62 may be configured to direct the thrust of the nozzle 1 in a direction other than parallel to a longitudinal axis of said nozzle 1, thus providing a thrust vectoring function.

[0069] As a result, the solution proposed by the present invention makes it possible to add an integrated monitoring network through said at least one sensor 50, in particular of the FBG type, which provides real-time on-board or remote control over both the cooling system and the thrust vectoring function. In addition to this, the solution of the present invention gives the possibility of integrating the optical fiber of said sensors 50 directly into the cellular structure 44. Moreover, the properties of the cellular structure 44 allow different physical parameters to simultaneously overcome the cross sensitivity of the fiber Bragg gratings, thus considerably simplifying the installation process. Note also that, in accordance with the present invention, the plurality of sensors 50, in particular of the FBG type, make it possible to monitor the whole system, not just the physical characteristics of the coolant fluid.

[0070] Figures 3a to 3d illustrate a possible third embodiment of the nozzle 1 according to the present invention; in this regard, it should be noted that the nozzle 1 shown in Fig. 3a belongs to a launcher running on solid fuel (identified by reference CS in Fig. 3a, and positioned in a casing 2, as already shown in Fig. 2a), but the characteristics of the nozzle 1 shown in Figures 3a to 3d can also be implemented in a nozzle 1 (like the one shown in Figuresla and lb) of a launcher running on liquid fuel.

[0071] The nozzle 1 shown in Figures 3a to 3d is made substantially in the same way as the one of the second embodiment (see Figures 2a to 2d), since said nozzle 1 comprises a cooling system SR comprising the tank 60, the injection means 61 and the draining means 62; in turn, said draining means 62 comprise a plurality of passages 62A formed in the bottom edge 31 of the diverging portion 30 and preferably configured to direct the thrust of the nozzle 1 in a direction other than parallel to a longitudinal axis of said nozzle 1, thus providing a thrust vectoring function. In particular, this control can advantageously be achieved by varying the flow rate of each duct 72; in particular, such flow rate variation can be selectively controlled.

[0072] In addition, the interspace 43 of the nozzle 1 shown in Figures 3a to 3d comprises a plurality of septa 71 configured to divide said interspace 43 into a plurality of ducts 72 for channelling the coolant fluid injected into the interspace 43 by the injection means 61. In substance, each septum 71 develops in a substantially radial direction within said interspace 43 and connects the inner panel 41 to the outer panel 42; moreover, each duct 72 is comprised between the tank 60, two septa 71, and the inner surfaces of the inner panel 41 and outer panel 42.

[0073] In the third embodiment of the nozzle 1 according to the present invention, said draining means 62 comprise at least one drain port 62B per duct 72, wherein said at least one drain port 62B is formed in the inner panel 41, in particular of the diverging portion 30. As can be observed in Figures 3a-3d, said draining means 62 comprise also at least one passage 62A per duct 72.

[0074] It is therefore apparent that, in the embodiment shown in Figures 3a to 3d, the coolant fluid flows in the interspace 43 in the same direction as the exhaust gases of the nozzle 1. In this context, also the drain ports 62B may be configured in such a way as to direct the thrust of the nozzle 1 in a direction other than parallel to a longitudinal axis of said nozzle 1, thus providing a thrust vectoring function.

[0075] In this regard, it should be noted that the fact that the interspace 43 is divided into a plurality of ducts 72 by a plurality of septa 71 results in the coolant fluid flowing in channels along a number n of directions, so that flow rates can be selectively increased in order to obtain a “directional thrust”.

[0076] In addition to this, the fact that the nozzle 1 is provided with drain ports 62B in its expansion section (i.e. the section running from the throat portion 20 to the bottom edge 31) advantageously permits the formation of a film of cold gas in the boundary layer.

[0077] As previously explained, it must be pointed out that also the nozzle 1 of a rocket engine or launcher running on liquid fuel (like the one shown in Figures la and lb) may comprise a cooling system SR (like the one shown in Figures 2a to 3d) and may further comprise the plurality of septa 71 dividing the interspace 43 into a plurality of ducts 72 (i.e. it may be constructed like the third embodiment shown in Figures 3a to 3d).

[0078] The advantages of the present invention are apparent from the above description.

[0079] Indeed, the peculiar provisions of the present invention result in a gas-dynamic nozzle of a rocket engine, in particular of a launcher, so conceived as to attain considerable technical and performance-related advantages over the solutions currently known in the art.

[0080] In particular, the nozzle 1 according to the present invention offers punctual and distributed control over some specific physical variables of the nozzle 1 and / or of the coolant fluid flowing in the interspace 43, in particular said physical variables concerning the temperature, pressure, vibration and deformation of the nozzle 1 and / or the temperature and pressure of said coolant fluid.

[0081] In this frame, the provision of at least one optical-fiber sensor 50 housed in said interspace 43, in particular at least one fiber Bragg grating sensor 50, makes it possible to attain precise monitoring with much shorter response times than possible with prior-art solutions (in particular comprising pressure switches or thermocouples), thus ensuring accurate control even under extreme operating conditions.

[0082] In addition, the monitoring of the coolant fluid flowing within the interspace 43, which may consist of the fuel of the launcher (as in the first embodiment shown in Figures la and lb, wherein the nozzle 1 belongs to a launcher running on liquid fuel) offers accurate information about the coolant fluid / fuel during the preheating phase, thereby providing real-time closed-loop control and contributing to improving the performance of the launcher. Clearly, the possibility of monitoring in real time the operating conditions of the coolant fluid during the pre-heating phase makes for increased safety because action can be readily taken in the event of a problem.

[0083] Another advantage of the nozzle 1 according to the present invention lies in the fact that the provision of said at least one sensor 50, in particular of the fiber Bragg grating type, makes it possible to take measurements by means of a pure optical system, without either undergoing or generating any electromagnetic noise and without the risk of triggering a short circuit or a fire. In addition, said at least one optical-fiber sensor 50 is minimally invasive, generates minimal measurement errors, gives the possibility of multiplexing or simultaneously interrogating any number of sensors 50 over the same optical communication line, and cannot generate sparks / short circuits or, most importantly, chemical inertia. Furthermore, said at least one sensor 50, in particular of the fiber Bragg grating type, allows monitoring multiple physical parameters while using the same hardware, in particular using multiple sensors 50 made of the same material.

[0084] A further advantage of the nozzle 1 according to the present invention lies in the fact that the provision of the open-cell structure 44 (which may also be defined as an “open-cell trabecular structure 44” or “cellular core”), in particular formed by additive manufacturing, contributes to optimizing both the weight of the nozzle 1 and the heat exchange, thereby allowing the use of materials having a lower melting point than those required by prior-art solutions. This technology makes it possible to integrate the sensors 50 into the physical component, thus improving the accuracy and response times thereof while reducing the total weight of the nozzle 1. Moreover, the provisions of the present invention make it possible to geometrically locate the measurement or reading taken. It should also be noted that the use of a “cellular core” (like the one of the present invention) instead of a tube bundle (like, for example, the one shown in document US3116603 A) makes it possible to integrate the cooling system into the structure of the nozzle 1 without increasing its weight, as well as to provide coolant discharge ports in a single component, thus improving the performance and not requiring complex assembly operations.

[0085] Another advantage of the nozzle 1 according to the present invention lies in the fact that, since it can acquire and process very important data relating to safety-critical systems, it will provide information about the thermal stress and structural state of the nozzle 1, resulting in simpler, less time-consuming and less expensive maintenance; this feature is very interesting for reusable launchers, the development of which has rapidly grown in recent years. In this regard, it should be noted that a considerable advantage of the nozzle 1 according to the present invention lies in increased launch safety and fewer “aborted” or unsuccessful launches, as well as in the small volumes of said at least one sensor 50, resulting in the possibility of sensorizing even small commercial launchers or small defence launchers.

[0086] It is therefore apparent that the solution of the present invention provides a considerable reduction in the costs incurred for manufacturing the nozzle 1, being well suited to fulfil the propulsion requirements of the Space Economy paradigm.

[0087] A further advantage of the solution proposed by the present invention lies in the fact that it makes it possible to add an integrated monitoring network through said at least one sensor 50, in particular of the FBG type, which provides real-time on-board or remote control over both the cooling system and the thrust vectoring function.

[0088] In addition to this, the solution of the present invention gives the possibility of integrating the optical fiber of said sensors 50 directly into the cellular structure 44. Moreover, the properties of the cellular structure 44 allow different physical parameters to simultaneously overcome the cross sensitivity of the fiber Bragg gratings, thus considerably simplifying the installation process. Note also that, in accordance with the present invention, the plurality of sensors 50, in particular of the FBG type, make it possible to monitor the whole system, not just the physical characteristics of the coolant fluid.

[0089] The nozzle 1 described herein by way of example may be subject to many possible variations without departing from the novelty spirit of the inventive idea; it is also clear that in the practical implementation of the invention the illustrated details may have different shapes or be replaced with other technically equivalent elements.

[0090] Among the various possible variants, for example, the nozzle 1 may be so constructed as to comprise:

[0091] - both at least one first aperture 42A for the entry of the coolant fluid (consisting of fuel) into the interspace 43 and at least one second aperture 42B for the exit of said coolant fluid (consisting of fuel) from said interspace 43 (as shown in Figures la and lb), wherein the nozzle 1 preferably comprises a channel configured to connect said first aperture 42 A and second aperture 42B;

[0092] - both the cooling system SR, in particular comprising the previously described tank 60, injection means 61 and draining means 62 (and optionally also the plurality of septa 71 dividing the interspace 43 into the plurality of ducts 72).

[0093] It is therefore clear that, in this variant, the nozzle 1 according to the present invention has a dual cooling system, i.e. a cooling system where the coolant fluid is fuel, and also an additional cooling system comprising coolant fluid contained in the tank 60, injected into the interspace 43 by the injection means 61, and drained from the interspace 43 through the draining means 62.

[0094] It can therefore be easily understood that the present invention is not limited to the abovedescribed nozzle 1, but may be subject to many modifications, improvements or replacements of equivalent parts and elements without departing from the inventive idea, as clearly specified in the following claims.

Claims

CLAIMS1. A gas-dynamic nozzle (1) of a rocket engine, in particular of a launcher, wherein said nozzle (1) comprises a converging portion (10), a throat portion (20), and a diverging portion (30), said nozzle (1) being characterized in that at least said throat portion (20) and at least said diverging portion (30) comprise an inner panel (41) and an outer panel (42), wherein said inner panel (41) and outer panel (42) define an interspace (43) in between and configured to allow a flow of coolant fluid for cooling said throat portion (20) and said diverging portion (30), wherein said interspace (43) comprises an open-cell structure (44), in particular said structure (44) being made by additive manufacturing, and wherein said nozzle (1) comprises at least one optical-fibre sensor (50) housed in said interspace (43) and configured to monitor at least one physical variable of the nozzle (1) and / or of the coolant fluid flowing in said interspace (43), in particular said at least one physical variable being related to the temperature, pressure, vibration and deformation of the nozzle (1) and / or being related to the temperature and pressure of said coolant fluid.

2. Nozzle (1) according to claim 1, characterized in that said at least one sensor (50) is positioned within said structure (44).

3. Nozzle (1) according to claim 1, characterized in that said at least one sensor (50) comprises at least one fiber Bragg grating sensor.

4. Nozzle (1) according to one or more of the preceding claims, characterized in that it comprises a plurality of sensors (50), in particular arranged in series with each other, wherein said sensors (50) are configured to operate at different frequencies, in particular for detecting different wavelengths of variation of the measured thermomechanical quantities.

5. Nozzle (1) according to one or more of the preceding claims, characterized in that also said converging portion (10) comprises said inner panel (41) and outer panel (42), which define said interspace (43).

6. Nozzle (1) according to claim 5, characterized in that said outer panel (42) comprises at least one first aperture (42A) for the entry of the coolant fluid into the interspace (43) and at least one second aperture (42B) for the exit of said coolant fluid from said interspace (43).

7. Nozzle (1) according to claim 6, characterized in that said coolant fluid consists of a fuel of the rocket engine, wherein said fuel enters the interspace (43) in liquid form through said at least one first aperture (42A) and exits said interspace (43), in particular in gaseous form, through said at least one second aperture (42B).

8. Nozzle (1) according to one or more of claims 6 and 7, characterized in that said at least one first aperture (42 A) is located in proximity to a bottom edge (31) of the nozzle (1), and said at least one second aperture (42B) is located in proximity to a top edge (32) of the nozzle (1).

9. Nozzle (1) according to one or more of the preceding claims, characterized in that said nozzle (1) comprises a cooling system (SR) comprising a tank (60) containing a coolant fluid, wherein said tank (60) is positioned within the interspace (43), in particular in a position corresponding to the throat portion (20) of the nozzle (1).

10. Nozzle (1) according to claim 9, characterized in that said tank (60) comprises injection means (61) configured to inject the coolant fluid into the interspace (43).

11. Nozzle (1) according to claim 10, characterized in that said interspace (43) comprises a plurality of septa (71) configured to divide said interspace (43) into a plurality of ducts (72) for channelling the coolant fluid injected into the interspace (43) by said injection means (61).

12. Nozzle (1) according to one or more of claims 9 to 11, characterized in that said cooling system (SR) of the nozzle (1) comprises draining means (62) for draining the coolant fluid from the interspace (43).

13. Nozzle (1) according to one or more of claims 9 to 11, characterized in that said draining means (62) comprise a plurality of passages (62 A) formed in a bottom edge (31) of the diverging portion (30) and / or at least one drain port (62B) per duct (72).

14. Nozzle (1) according to one or more of claims 12 and 13, characterized in that said draining means (62) are configured to direct the propulsion of the nozzle (1) in a direction other than the one parallel to a longitudinal axis of said nozzle (1), and to permit actions useful for controlling the attitude of the rocket engine.

15. A rocket engine, in particular a launcher, comprising a nozzle (1) according to one or more of claims 1 to 14.

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