Patent Grant
11554882
The present invention relates to space launchers. Embodiments described herein especially relate to an attitude control and thrust boosting system and method for space launchers, in particular for one or more stages of a multi-stage space launcher.
As it is well known, the almost universally used mechanism for controlling the attitude of a space launcher is currently that of varying the thrust direction of the rocket engine of the launcher by deflecting the exhaust nozzle of the engine (a system also known as TVC-thrust launcher control), so as to produce a force component orthogonal to the axis of the launcher, which in turn generates a control torque, which is a function of the distance between the engine thrust axis and the center of gravity of the launcher.
A typical TVC system uses, in particular, two linear actuators that are coupled to the exhaust nozzle of the rocket engine, are arranged on planes orthogonal to each other, and are so actuated as to deflect said nozzle so as to vary the thrust direction, thus generating the control torque.
Conventional TVC systems have the technical disadvantages described below.
In the case of solid-propellant rocket motors (SRMs) or of Hybrid-propellant Rocket engines (HREs), the nozzle shall be provided with a flexible joint, manufacturing whereof is quite complex and which has therefore a quite high cost; moreover, the mechanical features of the single units produced may differ also significantly from the nominal value of the technical specification.
In the case of liquid-propellant Rocket Engines (LREs), the unit “nozzle/combustion chamber” shall be provided with a universal joint (“gimbal”), in this case again with an increase in complexity, weight and cost of the launcher system.
The structural engineering of the nozzle shall take into account the loads produced by the actuators while the nozzle is deflected.
In the larger engine, the nozzle movable mass may be in the order of several hundreds kilograms, that results in high inertial value of the load applied to the TVC system, significantly limiting the dynamic features thereof (for example step command response, frequency band, etc.).
Moreover, the conventional TVC systems are also subjected to a non-stationary phenomenon of gas dynamics nature, the so-called impulse load, occurring during the ignition transient of a rocket engine. This phenomenon occurs in the divergent portion of the nozzle few thousandths of a second after engine ignition (before the flow in said divergent portion is completely supersonic), and is characterized by the occurrence of strong shock waves and discontinuity in gas efflux, which in turn induce significant impulsive loads onto the set “nozzle/actuators” of the above mentioned conventional TVC systems.
FR-A-2015/011198 discloses a combustion gas discharge device for a rocket engine, wherein around a stationary nozzle, i.e. a nozzle that is fixed with respect to the aircraft on which it is installed, flaps are provided that are angularly movable and define an extension of the nozzle. The flaps are arranged in two series, one inner series and one outer series, in order to provide a substantially continuous wall without spaces, through which the combustion gas could laterally escape. Each flap of each series of overlapped flaps is provided with a respective actuator to move the flap angularly. The arrangement is very complex and expensive. FR-A-2015/011198 generically mentions the fact that the angular arrangement of the flaps can be modified according to the flight condition and level, but no practical teaching is given on how to control the flaps. The only thing mentioned is that the position of the flaps may be such as to form a cylindrical or conical convergent prolongation of the fixed nozzle. Practically, even if in very general terms, it is suggested that the flaps take a convergent inclination with respect to the opening of the nozzle.
A first object of the invention is to provide an attitude control system for space launchers, which does not have the technical disadvantages mentioned above of the current conventional TVC systems.
A second object of the invention is to provide an attitude control system that, in addition to control the thrust direction of a rocket engine of a space launcher, is also able to increase the thrust thereof at flight height above the engine ignition height.
These and other objects are achieved through the present invention as it relates to an attitude control and thrust boosting system, according to what defined in the attached claims.
In particular, the attitude control and thrust boosting system according to the invention is so designed as to be installed on a space launcher equipped with a rocket engine provided with an exhaust nozzle that does not require to be deflected as in the traditional TVC systems, but is fixed, with the longitudinal axis thereof matching the longitudinal axis of the rocket engine; wherein said exhaust nozzle comprises a divergent portion so designed as to make a supersonic gas flow exit through an exit section defined by a given angle of divergence with respect to the longitudinal axis of the rocket engine.
The attitude control and thrust boosting system is characterized by comprising a plurality of flaps that are arranged around the exit section, are shaped so as to extend the divergent portion of the exhaust nozzle, are mechanically decoupled from said exhaust nozzle and can be actuated to take different angular positions with respect to the longitudinal axis of the rocket engine. The system further comprises control means configured for: (a) receiving quantities indicative of an actual attitude of the space launcher and an ambient static pressure; (b) making the flaps take a neutral angular position where the flaps are inclined, with respect to the longitudinal axis of the rocket engine, according to an inclination angle greater than, or equal to, the given angle of divergence; (c) controlling the neutral angular position taken by the flaps according to the ambient static pressure; (d) making one or more flaps take an angular position different than the neutral angular position according to the actual attitude of the space launcher and to a required attitude for said space launcher.
For better understanding the present invention, some preferred embodiments will be illustrated below with reference to the accompanying drawing, just by way of non-limiting example. In the drawing (not to scale):
The description below is provided to allow a person skilled in the art to work and use the invention. Modifications to the embodiments described herein will be immediately apparent to those skilled in the art and the generic principles disclosed herein can be also applied to other embodiments and applications without however departing from the protective scope of the invention as defined in the attached claims.
Therefore, the present invention is not limited to the embodiments described and illustrated herein; on the contrary, the scope of protection of the invention covers the principles and the features illustrated herein and defined in the attached claims.
An innovative aspect described herein is the use of a plurality of flaps (i.e. movable surfaces able to deflect fluid flows), preferably jet flaps, arranged in correspondence of an exit section of a divergent portion of an exhaust nozzle (for example a convergent-divergent nozzle, i.e. a de Laval nozzle) of a rocket engine of a stage of a multi-stage space launcher, wherein the flaps have a double function, i.e. they allow to control the attitude of the launcher and to augment the thrust of the engine at flight heights above the ignition height.
In particular, in embodiments described herein, the attitude control mechanism is based on the use of a number N (where N>3, preferably N=3) of flaps that are mechanically decoupled from the nozzle of the engine, which is kept fixed. The flaps are arranged in correspondence of the exit section of the nozzle and, when necessary, they are suitably inclined, singularly or in groups of M flaps (conveniently with M>2) so as to partially deflect the supersonic gas flow exiting from the exit section of the nozzle, thus creating a control torque.
Therefore, this attitude control mechanism is devoid of mechanical interface with the nozzle of the rocket engine.
Moreover, the attitude control mechanism described herein can synergistically combine the function of launcher attitude control with a function of thrust boosting and therefore of specific impulse of the engine for given flight phases, in an extremely advantageous manner for the low stages of a multi-stage launcher, in particular for the first stage.
In fact, the nozzles of the engines of the low stages of a multi-stage launcher have a relatively low expansion ratio due to the fact that the lower the flight height, the higher the ambient static pressure and consequently the lower shall be the expansion ratio of the nozzle to avoid the known phenomenon of over-expansion of the supersonic gas flow, with the consequent formation of shock waves and detachment of the limit layer from the inner walls of the divergent portion of the nozzle, which in turn cause a significant decrease in the engine thrust, as well as a malfunction and an irregular operation of the nozzle.
In particular, as it is well known, the expansion ratio of a convergent-divergent nozzle of a rocket engine is defined as the ratio between the total pressure at the nozzle throat (practically matching the average static pressure in the combustion chamber of the engine) and the static pressure at the exit section of the nozzle. This ratio, under conditions of supersonic flow in the divergent portion (sonic conditions in the nozzle throat) varies in homologous way (though non-linear) with the ratio of areas of the nozzle, i.e. the ratio between the area of the exit section of the nozzle and the area of the nozzle throat.
Conveniently, the flaps are so shaped as to represent, ideally, an extension of the divergent portion of the nozzle, thus allowing to have an increase in the expansion ratio of the nozzle at flight heights above the engine ignition height, with a consequent thrust and specific pulse boosting, as will be better described below.
Before describing specific embodiments of the invention, it is important to note that herein the term “neutral angular position” of the flaps means a position, which varies according to the ambient static pressure, i.e. to the flight height, as explained below, and in correspondence of which the flaps do not interfere in a counter-productive way with the expansion of the supersonic gas flow exiting from the nozzle.
In particular, in use, the flaps are so controlled as to take this neutral angular position (so as not to negatively interfere with the expansion of the supersonic gas flow exiting from the nozzle) when it is not necessary to deflect the supersonic flow, i.e., as it will be better described below, in absence of a deflection command for the attitude control of the launcher.
As, given an angular position of the flaps, the above-mentioned interference between the expansion of the supersonic gas flow exiting the nozzle and the flaps decreases as the ambient outer pressure decreases, and therefore, as the flight height increases, also the neutral angular position of the flaps varies according to the flight height.
In this regard, two main operational modes can be conveniently identified, i.e.: a type 1 operational mode implemented at low flight heights; and a type 2 operational mode implemented at high flight heights.
In particular, in type 1 operational mode, the neutral angular position corresponds, at the beginning, to an inclination of the flaps, with respect to the engine axis, greater than a preset reference angular position defined by the angle of divergence of the longitudinal profile of the nozzle at the exit section of said nozzle. In other words, the preset reference angular position is defined by the angle of divergence, with respect to the longitudinal axis of the rocket engine, characterizing the exit section of the nozzle.
In this way, in the initial step of engine operation, at low flight heights, it is possible to avoid the over-expansion of the supersonic gas flow exiting the nozzle, that is due to the relatively high values of the ambient static pressure at these heights.
Moreover, in the type 1 operational mode, the neutral angular position is varied gradually as the flight height increases, i.e. as the ambient static pressure decreases, up to tend to the preset reference angular position. More in particular, the corresponding inclination of the flaps with respect to the engine axis is gradually decreased. In this type 1 operational mode, as the flight height increases the thrust and specific pulse increases, the pulse being maximal when the flaps achieve the preset reference angular position.
In the type 2 operational mode, the neutral angular position substantially corresponds to the preset reference angular position, so as to have the maximal increase in thrust and specific pulse at high flight heights.
For a better understanding of the present invention,
In detail, the attitude control and thrust boosting system 100 is installed on a space launcher, preferably on a stage, for instance the first stage, of a multi-stage space launcher. The space launcher, or a stage thereof, is equipped with a rocket engine. The rocket engine may be, for instance, a solid-propellant rocket engine, a liquid-propellant rocket engine or a hybrid-propellant rocket engine. The rocket engine is provided with an exhaust nozzle. The exhaust nozzle comprises a divergent portion that is so designed as to make a supersonic gas flow exit through an exit section defined by a given angle of divergence with a longitudinal axis of the rocket engine. The longitudinal axis of the rocket engine is a central axis of symmetry of the exhaust nozzle and of the rocket engine. The exit section of the nozzle lies on a plane perpendicular to the longitudinal axis of the rocket engine.
The attitude control and thrust boosting system 100 includes a number N of flaps 110, for example jet flaps. As indicated above, the number N is equal to, or greater than, three, and preferably N=3. The flaps 110 are arranged at the exit section of the nozzle and around it. Moreover, the nozzles are so shaped as to extend the divergent portion of the exhaust nozzle. As it will be described in greater detail below, the flaps are mechanically decoupled from the exhaust nozzle. Moreover, the flaps can be actuated in order to take different angular positions with respect to the longitudinal axis of the rocket engine, according to criteria that will be described in greater detail below.
Preferably and conveniently, the flaps 110 are shaped and modeled on a curved surface ideally representing an extension of the divergent portion of the nozzle.
Moreover, the attitude control and thrust boosting system 100 conveniently includes also a number N of actuators 120. Preferably, the actuators 120 are linear actuators. Each actuator is coupled to a respective flap 110. Each actuator is controllable so as to make the respective flap 110 take different angular positions with respect to the longitudinal axis of the rocket engine.
Lastly, the attitude control and thrust boosting system 100 also includes a control unit 130 connected to the linear actuators 120 to control the operation thereof. The control unit 130 is also connected to an inertial platform 201 installed on the space launcher and configured to detect an actual (i.e. effective) attitude of the launcher and to send one or more (analogical or digital) output signals carrying one or more quantities indicative of the detected actual attitude, for example an actual attitude angle. The inertial platform may be for instance based on the use of gyroscopes. The control unit 130 is also connected to a pressure sensing device 202, for instance a piezoelectric or a potentiometric transducer, installed on the space launcher and configured to measure the ambient static pressure and to send one or more (analogical or digital) exit signals carrying one or more quantities indicative of the measured ambient static pressure.
In particular, the control unit 130 is connected to the inertial platform 201 to receive one or more signals sent by the inertial platform 201, and therefore to receive one or more quantities indicative of the actual attitude of the launcher.
The control unit 130 is also configured to store or to calculate or to receive (for example, from a flight control system of the launcher) one or more quantities indicative of an attitude required for the space launcher, for example a required attitude angle. The control unit 130 is also connected to the pressure sensing device 202 to receive one or more signals sent by the pressure sensing device 202 and therefore to receive one or more quantities indicative of the measured ambient static pressure.
In detail, the control unit 130 is configured to control the linear actuators 120, conveniently by sending suitable control instructions for controlling the neutral angular position of the flaps. In particular, the control unit 130 is so configured as to make the flaps 110 take a neutral angular position where the flaps 110 are inclined, to the longitudinal axis of the rocket engine, at an inclination angle greater than, or equal to, the given angle of divergence. Moreover, the control unit 130 is so configured as to control the linear actuators 120 in such a manner as to control the neutral angular position taken by the flaps 110 according to the measured ambient static pressure and, therefore, to the flight height corresponding to the measured ambient static pressure.
The control unit 130 is also so configured as to control the linear actuators 120, by sending suitable control instructions for controlling the attitude of the launcher, so as to make one or more flaps 110 take an angular position different than the neutral angular position, according to the actual attitude of the space launcher and to the required attitude for the space launcher.
Preferably, the control means 130 are so configured as to control the neutral angular position taken by the flaps 110 by decreasing the inclination angle as the ambient static pressure decreases.
In some embodiments, the control unit 130 is so configured as: to reduce the inclination angle as the ambient static pressure decreases until said inclination angle corresponds to said angle of divergence of the nozzle; and then to keep the inclination angle equal to the given angle of divergence independently of the ambient static pressure.
In some embodiments the control unit 130 is so configured as:
Preferably, the control unit 130 is so configured as to make one or more of the flaps 110 take an angular position different than the neutral angular position by comparing the actual attitude of the space launcher and the attitude required for the space launcher, for instance, by checking if the actual attitude angle and the required attitude angle are equal to each other or if they differ more than a preset threshold. If the actual attitude and the required attitude differ (or differ more than a preset threshold), based on said actual attitude and said required attitude, the central control unit 130 can determine an angular position where one or more of the flaps deflect the supersonic gas flow exiting the exit section so as to bring the actual attitude towards the required attitude. Once this angular position of one or more flaps has been determined, the central unit 130 can control the actuators of the flap(s) in order to make them take the given angular position.
More in particular, according to advantageous embodiments, the control unit 130 may be configured so that, if the actual attitude of the launcher and the required attitude differ, the following steps are performed:
As it is clearly apparent from the description above, the control of the position taken by the flaps 110, implemented by the control unit 130 through the linear actuators 120, have the double function of controlling the attitude of the launcher based on the actual attitude and the required attitude; and of controlling the neutral angular position, based on the ambient static pressure, i.e. on the flight level.
Moreover,
Conveniently, when the launcher is at low flight heights, i.e. below a preset threshold height, the control unit 130 implements the type 1 operational mode and performs the control of both the attitude of the launcher and the neutral angular position, varying the neutral angular position as the ambient static pressure decreases, i.e. as the flight height increases, up to achieve a preset reference angular position corresponding to an inclination of the flaps 110, with respect to the longitudinal axis of the rocket engine, substantially equal to that of the given angle of divergence characterizing the exit section of the nozzle. This preset reference angular position is also associated with the preset ambient static pressure threshold corresponding to the above mentioned preset threshold height.
Conveniently, when the launcher is at high flight heights, i.e. above the preset threshold height, the control unit 130 implements the type 2 operational mode, and performs only the control of the attitude of the launcher, whilst, in the absence of attitude control commands, the flaps 110 are kept in the preset reference angular position.
In this regard it should be noted that the law based on which, in the type 1 operational mode, the neutral angular position of the flaps 110 is determined according to the ambient static pressure depends on the features of the specific fluid-dynamic field existing inside and outside the nozzle and can be therefore conveniently defined case by case based on CFD (computational fluid dynamics) simulations and/or experimental tests.
The control unit 130 may be configured so as to determine the neutral angular position of the N flaps 110, i.e. the inclination angle to the longitudinal axis of the rocket engine, by executing a first preset calculation function, or using a first preset lookup table, where values are stored of the inclination angle associated with respective height values, i.e. values of ambient static pressure.
In the same way, the control unit 130 can be conveniently configured to determine the angular positions of the flaps 110 for the attitude control by executing a second preset calculation function, or using a second preset lookup table.
In some embodiments, the control of the attitude of the launcher and that of the neutral angular position may be implemented by a single processing and control unit, as shown in
An exemplary embodiment of the flaps 110 and the linear actuators 120 is shown in
In the example of
In the example of
The jet flaps 111, 112, 113 are arranged at 120° from one another in a plane orthogonal to the longitudinal axis AL of the rocket engine and are advantageously provided with a suitable thermal insulation, both on the inner surfaces and on the outer surfaces.
In some embodiments, the jet flaps 111, 112, 113 are hinged to a support structure 140 extending around the exit section of the divergent portion 302 of the nozzle substantially on a plane orthogonal to the axis AL. In the example of
The support structure 140 is fixed to the outer structure 301 of the launcher stage, for example to an engine flange or an inter-stage flange, as in the case of
Each jet flap 111, 112, 113 can be connected to the support structure 140 by means of a respective pair of traditional hinges 150, as shown in the example of
Moreover, in the example of
In the example of
It should be noted that the example shown in
Moreover, the actuators 120 may be designed using different technologies. For example, electro-mechanical actuators can be used, or hydraulic actuators, pneumatic actuators etc., of the linear or rotary type or of any other type.
In any case, it should be noted that the support structure 140, on which the jet flaps 111, 112, 113 are hinged, practically avoids any mechanical interface between the nozzle and said jet flaps 111, 112, 113, that are therefore mechanically decoupled from the nozzle. In this way, in use, the jet flaps 111, 112, 113, the support structure 140 and therefore the outer structure 301 of the launcher stage are subjected to structural loads, but not the nozzle.
In the example of
It should be understood that the attitudes of the jet flaps 111, 112, 113 shown in
It shall be noted that in
In advantageous embodiments, in order to minimize the side leakages of gas through the gaps between jet flaps 111, 112, 113 and therefore to increase the efficiency of the system, inter-flap panels can be used arranged at said spaces.
To this end,
In particular, in the preferred embodiment of
The inter-flap panels 160 are conveniently provided with suitable thermal insulation and have curved shape adapted to the shape of the jet flaps 111, 112, 113. In the example illustrated in
It is clearly apparent from the description above that the present invention have many advantages, among which the following are worth mentioning:
In
To better understand what illustrated in
T={dot over (m)}Ve+Ae(Pe−Pa), (1)
where {dot over (m)} indicates the mass flow of the gases exiting from the nozzle, Ve indicates the speed of the gases at the exit section of the nozzle, Ae indicates the area of the exit section of the nozzle, Pe indicates the static pressure of the gases at the exit section of the nozzle, and Pa indicates the ambient static pressure depending on the flight level, i.e. Pa=Pa(h), where h=flight height.
To better understand the meaning of the equation (1), it should be recalled that, as the expansion ratio of the nozzle varies, Ve and Pe vary in opposite manner (i.e. as one increases the other decreases and vice versa) and that, in any case, for a given value of ambient static pressure Pa (where Pa decreases as the flight height increases), the maximal thrust value is obtained at that specific expansion ratio for which Pa=Pe (matching condition of the nozzle). On the contrary, if the expansion ratio of the nozzle increases behind the value of the matching condition, the phenomenon of over-expansion of the supersonic flow occurs, and in this condition the thrust and the specific impulse decrease in monotonic way as the expansion ratio increases, due to the formation (as mentioned above) of shock waves (that are as more intense as higher the over-expansion level is) and the consequent detachment of the limit layer from the inner walls of the divergent portion of the nozzle, wherein this latter event induces, beyond a given over-expansion level, an anomalous or irregular operation of the nozzle.
Therefore, with reference to
Once arrived at higher heights (typically in the order of 10/15 km), the flaps 110 are deflected to the reference angular position, thus practically extending the divergent portion of the nozzle and therefore increasing the expansion ratio thereof, without in this case incurring in the problems due to the over-expansion of the supersonic flow that would occur at lower heights, as mentioned above. Consequently, the increase in said expansion ratio of the nozzle entails a thrust and specific impulse boosting with respect to the case in which the invention is not used.
A further particular aspect of the structure of the exhaust nozzle and of the flaps is shown specifically in
Just by way of example, in
In
As can be clearly understood from
Therefore, in particular when the angle of divergence of the flaps 111, 112, 113 is greater than the angle of divergence a (angle between the engine axis AL and the tangent to the exit section of the nozzle) of the divergent portion 302, as shown in particular in
This is beneficial as the risk of flow detachment from the inner walls of the flaps is reduced also in solid-propellant rocket engines, where e separate gas flow, to be conveyed towards the coupling edge of the flaps, is not available (i.e. a flow not coming from the combustion chamber).
The air sucked from the outside thanks to the ejector effect through the gap 360 between each flap 111, 112, 113 and the divergent portion 302 has also the advantage of reducing the temperature on the inner wall of the flaps.
In
Here on the method for controlling the angular opening of the flaps 111, 112, 113, which has been described above in principle with reference to
For each flap 111, 112, 113 the control algorithm comprises the two following distinct functions:
In some embodiments, the sensors used to obtain the input data for running the algorithm may comprise the following:
The input data required for the calculation of the desired static pressure, at the trailing edge of the flap, i.e. for the calculation of [(PS)FTE]Target are the following:
The input data for the calculation of the neutral angular position of the flap βDAP for the calculated value [(PS)FTE]Target are:
The calculation of the target static pressure inside the trailing edge of the flap, i.e. of [(PS)FTE]Target requires, as a prerequisite, the prediction of the corresponding flow separation static pressure on the inner side of the trailing edge of the flap indicated with (PSEP)FTE. The separation static pressure can be suitably determined by means of a criterion of flow separation for rocket nozzles, using for example: the ambient static pressure and the values of the local Mach number of the flow and/or of any other flow parameter required. Once the actual flow separation static pressure (PSEP)FTE has been obtained, it is possible to calculate the target static pressure by simply adding, to the value obtained from the calculation, a safety margin (ΔPS)Margin taking into account leakages, calculation uncertainties and other factors. Practically, therefore:
[(PS)FTE]Target=(PSEP)FTE+(ΔPS)Margin
It should be taken into account that the value of the neutral angular position βDAP obtained through the above described calculation is generally different from flap to flap. However, as it is necessary to adopt the same angular position for all flaps 111, 112, 113, as angular position a position will be adopted, which is defined for example through one of the following criteria:
The value of the sound speed α, for calculating the Mach number, is determined as follows:
α=√{square root over (γR(TS)FTE)}
where (TS)FTE is the static temperature (absolute temperature, i.e. in Kelvin degrees) of the gas at the trailing edge of the flap, γ is the ratio of the specific heats at constant pressure and constant volume of the gas, R is the constant of the gas.
It should be taken into account that these parameters generally depend on the degree of mixing between combustion gas from the exhaust nozzle 352 and air sucked through the gap 360 at the leading edge of the flap 111, 112, 113. This can be taken into account by means of functions comprised in the algorithm, which take into account the following variables:
The functions comprised in the algorithm can be determined experimentally and/or through Computational Fluid Dynamics (CFD) simulations.
In
The block diagram of
The set of blocks in
The attitude control is practically performed as follows. The inertial platform 473 measures the attitude angles of the launcher and more precisely the pitch angle and the yaw angle. The on-board computer (block 465) compares the measured attitude angles and the target ones and sends to the attitude control unit 471 the target deflection values for the various flaps (three in the illustrated example). These angles take into account the neutral angular position calculated according to the flight height (
The invention has been described with reference to various embodiments; however, it will be clearly apparent to those skilled in the art that modifications, variants and omissions can be done to the invention, without however departing from the scope of protection thereof as claimed in the attached claims. Furthermore, if not otherwise stated, the order or sequence of any step of method or process may be changed according to alternative embodiments.
| Number | Date | Country | Kind |
|---|---|---|---|
| 102017000063102 | Jun 2017 | IT | national |
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/IB2018/054092 | 6/7/2018 | WO |
| Publishing Document | Publishing Date | Country | Kind |
|---|---|---|---|
| WO2018/224998 | 12/13/2018 | WO | A |
| Number | Name | Date | Kind |
|---|---|---|---|
| 3229457 | Rowe | Jan 1966 | A |
| 3304722 | Culpepper | Feb 1967 | A |
| 10316796 | Dutheil | Jun 2019 | B2 |
| 20020079404 | Schroeder | Jun 2002 | A1 |
| 20100320329 | Boelitz | Dec 2010 | A1 |
| 20160177875 | Dutheil | Jun 2016 | A1 |
| Number | Date | Country |
|---|---|---|
| 106 593 696 | Apr 2017 | CN |
| 2015011198 | Jan 2015 | FR |
| H02 99751 | Apr 1990 | JP |
| H06 257512 | Sep 1994 | JP |
| 02092988 | Nov 2002 | WO |
| 2015011198 | Jan 2015 | WO |
| WO-2015011198 | Jan 2015 | WO |
| Entry |
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| Benson, Tom, Nozzle Design, Jan. 18, 2008 (Year: 2008). |
| Wikipedia, ST-124-M3 inertial platform, Apr. 25, 2012 (Year: 2012). |
| Number | Date | Country | |
|---|---|---|---|
| 20210147099 A1 | May 2021 | US |
