Fluid type thrust direction control device

Abstract

To provide a thrust direction control device that not only satisfies both a condition that a thrust is deflected when it should be deflected and a condition that the thrust is not deflected when it should not be deflected, but also satisfies a first condition suitably even with a secondary jet flow of a small flow rate.SOLUTION: In a thrust direction control device using a dual throat nozzle 10 having a first throttling part 12, an expansion part 14, and a second throttling part 15, a ratio A2* / A1* of a cross-sectional area A2* of a second throat 16 formed at a downstream end part of the second throttling part 15 to a cross-sectional area A1* of a first throat 13 formed at a downstream end part of the first throttling part 12 is set to 1.2 to 1.8.SELECTED DRAWING: Figure 1

Description

[Technical Field]

[0001] The present invention relates to a fluid thrust vector control device for controlling the direction of thrust of a rocket engine, a jet engine, or the like. [Background technology]

[0002] Many aircraft and other flying objects use aerodynamic steering to control their motion. However, aerodynamic steering may not be effective under low dynamic pressure conditions. For example, when the surrounding environment is a vacuum or low density, as in the case of space rockets and spacecraft operating in outer space or at high altitudes, or when the vehicle's speed is very slow immediately after launch, such as in the case of space rockets and ground-launched missiles, aerodynamic control (aerodynamic steering) becomes extremely difficult. To control the vehicle under these conditions, thrust vector control (TVC) is used. Thrust vector control is a technology that controls the direction of thrust by somehow changing the direction of the jet exhaust from a rocket engine or jet engine. This enables control of the vehicle's attitude and direction of travel even in situations where aerodynamic steering is ineffective. In recent years, aerodynamic steering and TVC have also been used in combination as a means of ensuring high maneuverability in aircraft.

[0003] Thrust vector control devices can be broadly divided into two types: mechanical and fluidic. Mechanical thrust vector control devices mechanically change the shape of the jet flow path. Mechanical thrust vector control devices include those that install an inclined plate inside the engine nozzle, those that change the nozzle shape, and those that change the orientation of the engine itself. On the other hand, fluidic thrust vector control devices use the action of the fluid flowing inside the nozzle to change the direction of the jet issuing from the nozzle outlet. A typical fluidic thrust vector control device injects a secondary jet into the main flow flowing inside the nozzle, causing a change in the main flow and deflecting the overall direction of the jet. Fluidic thrust vector control is superior to mechanical thrust vector control in that it is easier to reduce weight and improve response performance.

[0004] As a fluidic thrust vector control device using a secondary jet, one using a dual throat nozzle has been proposed (see, for example, Non-Patent Documents 1 and 2). Here, a "dual throat nozzle" refers to a nozzle shaped such that the outlet of a Laval nozzle, which is normally used as a nozzle for rocket engines, etc., is narrowed to provide a second throat. Figure 1 shows a conceptual diagram of a dual throat nozzle 10 that has been proposed in the past. In the figure, reference numeral "10" denotes the dual throat nozzle, reference numeral "10a" denotes the nozzle centerline, reference numeral "11" denotes the nozzle inlet, reference numeral "12" denotes the first narrowing portion, reference numeral "13" denotes the first throat, reference numeral "14" denotes the diverging portion, reference numeral "15" denotes the second narrowing portion, reference numeral "16" denotes the second throat, reference numeral "18" denotes the nozzle outlet, reference numeral "19" denotes the secondary jet inlet, and reference numeral "F IN1 " represents the main flow from the engine combustion chamber to the nozzle inlet, and "F IN2 " indicates the secondary jet injected from the secondary jet inlet, and is indicated by the symbol "F OUT " represents the jet flow coming out of the nozzle outlet, and "F a "," "F b "," "F c " and "F d " indicates a typical flow line, and the symbol "α1" indicates a separation region (the region indicated by hatching in the figure).

[0005] As shown in FIG. 1, a secondary jet F is injected from a secondary jet injection port 19 provided on the lower nozzle wall near the first throat 13. IN2 When injected, the mainstream F IN1 The flow of the streamline F a ,F b ,F c ,F d (Especially the lower streamline F d ), the main flow F separates from the lower inner wall surface of the dual throat nozzle 10, and a separation region α1 is formed between the secondary jet injection port 19 and the second throat 16 on the lower inner wall surface. IN1 is pushed up significantly upward and is pushed back in the opposite direction by the second throttle portion 15 immediately before the nozzle outlet 18. OUT The thrust generated as a reaction to this deflection is deflected to the upper left rather than the direction parallel to the nozzle center line 10a. In a dual throat nozzle 10 using a secondary jet, the thrust direction is controlled by the above mechanism (hereinafter sometimes referred to as the "flow separation mechanism"). [Prior art documents] [Patent documents]

[0006] [Patent Document 1] Patent No. 7197895 [Non-patent literature]

[0007] [Non-Patent Document 1] Flamm, J.D., Deere, K.A., Mason, M.L., Berrier, B.L., and Johnson, S.K., “Experimental Study of an Axisymmetric Dual Throat Fluidic Thrust Vectoring Nozzle for Supersonic Aircraft Application”, AIAA Paper, 2007-5084, 2007. [Non-Patent Document 2] Shin, C.S., Kim, H.D., Setoguchi, T. and Matsuo, S., “A Computational Study of Thrust Vectoring Control Using Dual Throat Nozzle”, J. Thermal Science, 19, 6(2010), pp. 486-490. [Non-Patent Document 3] Fearn, R.M., Mullin, T. and Cliffe, K.A., “Nonlinear Flow Phenomena in a Symmetric Sudden Expansion”, J. Fluid Mech., 211(1990), pp. 595-608. [Non-Patent Document 4] Durst, F., Pereira, J.C.F. and Tropea, C., “The Plane Symmetric Sudden-Expansion Flow at Low Reynolds Numbers”, J. Fluid Mech., 248(1993), pp. 567-581. [Non-Patent Document 5] Nakanishi Suketsugu, Sakurai Motoyasu, Oda Shingo, “Numerical Study on Asymmetric Flow in a Two-Dimensional Symmetric Sudden Expansion Flow Path”, Hiroshima Institute of Technology Research Bulletin, 30(1996), pp. 37-43. [Non-Patent Document 6] By Liepmann and Roshko, translated by Tamada Kou (Kou is Wang biased with Guang), “Gas Dynamics”, Yoshioka Shoten, 1960. [Summary of the Invention] [Problem to be solved by the invention]

[0008] Incidentally, when examining the practicality of a thrust vector control device, including a dual throat nozzle using a secondary jet, it is necessary to confirm the condition that "thrust is deflected when it should be deflected" (hereinafter, sometimes referred to as the "first condition") and the condition that "thrust is not deflected when it should not be deflected" (hereinafter, sometimes referred to as the "second condition"). However, while the first condition has already been confirmed in Non-Patent Document 1 and Non-Patent Document 2, the second condition has not yet been confirmed. For this reason, the inventors conducted a simulation to confirm whether a dual throat nozzle using a secondary jet satisfies the second condition. Specifically, in the dual throat nozzle 10 shown in FIG. 1, the secondary jet F IN2 Mainstream F without injection IN1 The flow field and thrust vector angle in the dual throat nozzle 10 when only the air flows in were calculated by numerical simulation.

[0009] The simulation analysis method used was the Direct Simulation Monte Carlo (DSMC) method. The molecular model was a hard sphere, which was assumed to be specularly reflected by the inner circumferential wall of the dual throat nozzle 10. The thrust (vector quantity "f") was calculated using the total momentum (vector quantity "Δp") of the molecules ejected from the nozzle outlet 18 during a short time (Δt) using the following equation 1. The thrust deflection angle (δ) was calculated using the following equation 2.

[0010]

number

number

[0011] In this simulation, the dual throat nozzle 10 was defined as a two-dimensional nozzle with the same shape in any cross section perpendicular to the y-axis direction, similar to that shown in FIG. 2. The dimensions of each part of the dual throat nozzle 10 in a cross section perpendicular to the y-axis direction (the dimensions of each part relative to the opening width D1 of the nozzle inlet 11) were set to the values ​​shown in FIG. 3. Hereinafter, the dual throat nozzle 10 with the dimensions shown in FIG. 3 may be referred to as "nozzle A." The width (cross-sectional area) of the second throat 16 in nozzle A was set to the same dimension as the width (cross-sectional area) of the first throat 13, similar to conventionally proposed thrust vector control devices using dual throat nozzles (thrust vector control devices using dual throat nozzles in Non-Patent Documents 1 and 2). The calculation started from an initial state in which the entire interior of the dual throat nozzle 10 was vacuum, and continued until it was confirmed that the flow field within the dual throat nozzle 10 was at least in a substantially steady state.

[0012] Prior to this simulation (hereinafter sometimes referred to as the "simulation to confirm the second condition") to confirm whether or not the above-mentioned "second condition" is satisfied, a simulation (hereinafter sometimes referred to as the "simulation to confirm the first condition") was also carried out to confirm whether or not the deflection of the thrust direction due to the "flow separation mechanism" explained in the "Background Art" section actually occurs (whether or not the first condition is satisfied). Specifically, in a two-dimensional nozzle (nozzle A) with the dimensions shown in Figure 3, the main flow F IN1 and secondary jet F IN2 A simulation was performed on the flow inside the dual throat nozzle 10 when the secondary jet F was injected. The analysis method for the first condition verification simulation was the same as that for the second condition verification simulation described above. However, in the first condition verification simulation, as shown in FIG. IN2 is injected. This secondary jet F IN2 The flow rate of the main stream F IN1 and secondary jet F IN2 The secondary jet F was set to 5% of the total flow rate of the IN2The injection direction of the thrust was rotated 150° counterclockwise from the rightward direction (positive x-axis direction) parallel to the nozzle centerline 10a in Figure 1 toward the page. Calculations for this case confirmed that the thrust deflection angle δ in the steady state was approximately -19°. The thrust deflection angle δ calculated using Equation 2 above is positive when the thrust is deflected counterclockwise in Figure 3. Therefore, the thrust deflection angle δ is negative as a result of the thrust being deflected to the upper left by the mechanism shown in Figure 1. The magnitude (absolute value) of the thrust deflection angle δ obtained here is consistent with the magnitude of the thrust deflection angle δ obtained in experiments and numerical simulations in Non-Patent Documents 1 and 2, which were conducted on a dual-throat nozzle with equal cross-sectional areas for the first and second throats, similar to "Nozzle A." The maximum thrust deflection angle δ was approximately 20°, the minimum thrust deflection angle was approximately 16°, and the average thrust deflection angle was approximately 18° when the flow rate ratio was 5%.

[0013] Figure 4 shows the results of the second condition confirmation simulation carried out for nozzle A. This figure is a graph showing the change in thrust deflection angle δ over time. In this figure, the random number sequence used in the simulation was set to 10 different values, and the results of each calculation are displayed overlapping each other. However, all conditions other than the random number sequence are the same. In the second condition confirmation simulation, the series of random numbers are used to set the speed of each inflowing molecule, select molecules that will undergo binary collision, and determine the scattering direction after the collision. Time is also calculated as D1 / (2RT1) 1 / 2 The values ​​are dimensionless, with D1 as the reference value. Here, D1 is the opening width of the nozzle inlet 11, R is a physical quantity equivalent to the gas constant of the simulated molecules used in the calculation, and T1 is the absolute temperature of the inflow from the nozzle inlet 11. As is clear from Figure 4, until around time t = 13, the thrust deflection angle δ is approximately 0° for all random number sequences. However, in the range of t > 13, the thrust deflection angle δ can take positive or negative values ​​depending on the random number sequence, and its absolute value can reach a maximum of approximately 4°. Even though the nozzle shape and inflow conditions are set symmetrically in the vertical direction, the direction of the thrust vector is deflected from the axis of symmetry, and the direction and magnitude of the deflection vary in various ways due to the accidental factor of the random number sequence.

[0014] Figures 5(a) and 5(b) show the streamlines in the nozzle at t=37.5 after sufficient time had passed for random number sequences 4 and 10, which yielded large absolute values ​​for the deflection angle. Figures 5(a) and 5(b) clearly show the occurrence of vertically asymmetric separation regions within the dual throat nozzle 10. Furthermore, for both random number sequences 4 and 10, the entire region within the dual throat nozzle 10 is subsonic or transonic.

[0015] Thus, statistical fluctuations are reflected in the calculation process and results of a simulation. These fluctuations also exist in real flow fields. However, the number of molecules in the simulation is approximately 10^6, which is much smaller than the actual Avogadro's number. Therefore, statistical fluctuations are more amplified in the simulated flow field than in the real flow field. On the other hand, real flow fields also contain macroscopic turbulence that is not considered in the simulation. Because this macroscopic turbulence can have the same effect as the above-mentioned fluctuations, the above-mentioned asymmetry may also occur in real flows. Although the shape is different from that in this case, the phenomenon of asymmetric separation regions occurring when an incompressible flow passes through a sudden expansion section of a two-dimensional pipe symmetrical about the centerline of the flow channel has also been confirmed experimentally (see Non-Patent Documents 3 and 4). Generally, flow separation is likely to occur in a diverging section, and even if the boundary shape is symmetrical about the centerline of the flow channel, there is no constraint imposed that the separation region must also be symmetric. In fact, numerical analysis has shown that solutions with asymmetric separation regions exist if the Reynolds number is not too small (see Non-Patent Documents 3, 4, and 5). In general, the number of solutions increases with the Reynolds number. Which solution an individual flow field will move towards is affected by minute turbulence or fluctuations present in the flow field or slight irregularities on the wall surface.

[0016] Furthermore, in the case of high-speed flows where compressibility cannot be ignored, such as the flow field in a fluidic thrust vector control device, even if a non-separated flow is achieved in the diverging section, if the flow velocity is below the speed of sound, the flow will slow down in the diverging section of the flow channel, creating an adverse pressure gradient (a condition in which pressure increases downstream), and the volume of the fluid will decrease as it moves downstream. In a situation where the flow channel expands and the fluid volume decreases simultaneously, it is difficult to maintain a non-separated flow, and it is natural that the flow will transition to a separated flow at some point. In such a case, if a solution exists in which the separated region has an asymmetric shape, it is not at all unnatural for the flow field to transition in that direction.

[0017] As mentioned above, in a fluidic thrust vector control system, in addition to the first condition, "thrust deflection when it should be," it is important that the second condition, "thrust not deflected when it should not be," is also met. In this regard, the fluctuations in thrust direction when thrust should not be deflected, which appeared in the simulation confirming the second condition (Figure 4), may also appear in a real fluidic thrust vector control system as "thrust deflection when the secondary jet is stopped due to an accidental factor." This could be a serious problem that affects whether a fluidic thrust vector control system using a dual throat nozzle can be put into practical use. Therefore, this problem must be resolved before practical use of a fluidic thrust vector control system using a dual throat nozzle can be realized.

[0018] Incidentally, the present inventor has proposed a technology for a thrust direction control device using a dual throat nozzle that can satisfy not only the condition that "thrust is deflected when it should be deflected" (first condition), but also the condition that "thrust is not deflected when it should not be deflected" (second condition), and has already obtained a patent for this technology (see Patent Document 1).

[0019] The fluid type thrust vector control device of Patent Document 1 is as follows: A fluid type thrust vector control device that changes the direction of a jet flowing from a nozzle outlet by injecting a secondary jet from a secondary jet injection port provided in the middle of the nozzle into a main flow flow that flows from a nozzle inlet to a nozzle outlet, The nozzle is a first throttle portion having an upstream end serving as a nozzle inlet, a downstream end having a cross-sectional area smaller than that of the upstream end, and a downstream end serving as a first throat; a diverging portion connected to a downstream side of the first throttling portion, the diverging portion having a cross-sectional area at a downstream end thereof larger than a cross-sectional area at an upstream end thereof; a second throttle portion connected to the downstream side of the diverging portion, the cross-sectional area of ​​the downstream end of which is smaller than the cross-sectional area of ​​the upstream end of which, and the downstream end of which serves as a second throat; The nozzle is a dual throat nozzle having Cross-sectional area of ​​the second throat (A2 * ") is the cross-sectional area of ​​the first throat ("A1 * "). By setting this value larger than the above, during non-deflection control in which only the main flow flows in without injecting a secondary jet into the nozzle, a shock wave is generated across the nozzle at the diverging section, and the flow in the area upstream of the shock wave becomes supersonic, and by preventing the flow from separating from the nozzle inner wall surface, unintended thrust deflection is prevented. When the Mach number upstream of the shock wave is M1 and the specific heat ratio is γ, Ratio A2 * / A1 * Substituting the value of into the left side of the following equation 3, and solving the following equation 3 and the following equation 4 simultaneously, the cross-sectional area A S In contrast, The secondary jet inlet has a cross-sectional area A S The cross-sectional area is 0.9 to 1.1 times that of the It is characterized by the following.

number

number

[0020] The technology of Patent Document 1 makes it possible to make a thrust vector control device using a dual throat nozzle also satisfy the condition (second condition) that "thrust is not deflected when it should not be deflected." That is, as described above, when only the main flow is allowed to flow into the nozzle without injecting a secondary jet (during non-deflection control), a shock wave becomes stationary at a specific position in the diverging section, and the flow becomes supersonic in the region between the location where the first throat is formed and the location where the shock wave is stationary. This suppresses separation within the nozzle, realizes a flow field that is approximately symmetrical with respect to the nozzle centerline, and makes it possible to suppress fluctuations in the thrust direction (keep the thrust deflection angle approximately constant at approximately 0°).

[0021] Second throat cross-sectional area A2 * The cross-sectional area of ​​the first throat is A1 * The reason why a shock wave occurs in the diverging section when the cross-sectional area A2 is increased can be explained as follows using the theory of one-dimensional flow approximation. * is the cross-sectional area A1 * In both throat sections (first throat and second throat) of a larger dual throat nozzle, the Mach number is 1, i.e., the flow velocity is equal to the sonic speed, due to the properties of compressible fluid. In a continuous flow tube (a tube without branches or mid-flow injection), the cross-sectional area where the flow becomes sonic is inversely proportional to the total pressure at each position. In other words, when the Mach number is 1 at both throats, the following equation 5 holds (see Non-Patent Document 6). In the following equation 5, "p 01 " and "p 02 " are the total pressures at the first and second throats, respectively.

number

[0022] Furthermore, when the total pressure ratio on the right side of the above equation 5 (which exceeds 1 due to total pressure loss) is caused by one shock wave, the following equation 6 can be derived from the Rankine-Hugoniot equation. In the following equation 6, "M1" is the Mach number upstream of the shock wave, and "γ" is the specific heat ratio.

number

[0023] Ratio of cross-sectional areas of both throats A2 * / A1 * The total pressure ratio p can be calculated by substituting into the above equation 5. 01 / p 02 By substituting into the left side of the above equation 6, the Mach number M1 upstream of the shock wave can be calculated. The cross-sectional area of ​​the nozzle at the point where the Mach number M1 upstream of the shock wave is S ") A shock wave exists at the cross section A S The value of can be calculated by substituting the value of the upstream Mach number M1 of the shock wave obtained from the above equation 6 into the following equation 7 under the approximation that there is an isentropic flow between the first throat and the shock wave. Since the upstream Mach number M1 of the shock wave is greater than 1, the cross-sectional area ratio A S / A1 * is greater than 1. Therefore, the cross-sectional area of ​​the second throat A2 * The cross-sectional area of ​​the first throat is A1 * It can be seen that by making it larger than , a shock wave is generated in the above-mentioned expanding portion.

number

[0024] The reason why separation in the nozzle is suppressed by the flow becoming supersonic in the region between the formation position of the first throat and the stationary position of the shock wave can be explained as follows: In other words, in a supersonic flow, the flow accelerates in the diverging section, creating a favorable pressure gradient (a phenomenon in which pressure decreases downstream), and the fluid expands as it moves downstream (hereinafter, this expansion mechanism may be referred to as the "expansion mechanism of supersonic flow in the diverging section"). Therefore, even if the cross-sectional area of ​​the flow path increases, the fluid expands to fill the flow path, making it easier to achieve a flow without separation. In other words, the second condition, that thrust will not be deflected when it should not be (when a secondary jet is not injected into the nozzle), is more easily satisfied.

[0025] By the way, as mentioned above, the cross-sectional area A2 of the second throat * The cross-sectional area of ​​the first throat is A1 * If the second condition, "thrust deflection when it should not be deflected," is satisfied by making the flow supersonic in the region between the first throat and the shock wave when no secondary jet is injected, a secondary jet is injected through a secondary jet injection port provided on the nozzle wall near the first throat to deflect the thrust, causing flow separation within the nozzle. However, due to the expansion mechanism of the supersonic flow in the diverging section, the main flow expands, and the separation may quickly disappear (the separation may disappear at a location close to the downstream side of the separation). This makes it difficult to satisfy the first condition, "thrust deflection when it should be deflected." To address this issue, the thrust vector control device of Patent Document 1 provides a secondary jet injection port in the diverging section.

[0026] This reduces the cross-sectional area of ​​the second throat A2 * The cross-sectional area of ​​the first throat is A1 *Even if the secondary jet injection rate is set larger than the maximum value (M1) so that the flow becomes supersonic in the region between the first throat and the shock wave when no secondary jet is injected, the Mach number M1 upstream of the shock wave decreases or the shock wave disappears when a secondary jet is injected into the nozzle. When the Mach number M1 upstream of the shock wave decreases, the expansion effect of the supersonic flow in the diverging section weakens, making it difficult for the separated main flow in the nozzle (the main flow pushed away from the secondary jet injection port in the nozzle) to return until it reaches the second throttling section (making separation difficult to resolve). If the injection rate of the secondary jet is sufficiently small, the shock wave traverses the nozzle, just as when no secondary jet is injected. However, the decrease in the Mach number M1 upstream of the shock wave causes the shock wave to move upstream, where the cross-sectional area of ​​the flow path is smaller, expanding the subsonic region, again making it difficult to resolve separation in the main flow. On the other hand, when the shock wave disappears, the flow becomes subsonic throughout the entire diverging section, making it difficult to resolve separation in the main flow. In either case, it is possible to prevent the main flow that has separated from the nozzle from returning (the separation is not resolved) until it reaches the second choke section. This makes it easier to satisfy the first condition that "thrust is deflected when it should be deflected."

[0027] Second throat cross-sectional area A2 * The cross-sectional area of ​​the first throat is A1 * In a nozzle with a cross-sectional area A1 larger than the above, when a secondary jet is injected from the diverging section, the Mach number M1 upstream of the shock wave becomes smaller or the shock wave disappears. This is because the secondary jet injected into the diverging section does not pass through the first throat but is included in the flow that passes through the second throat. Therefore, the cross-sectional area A1 in the above equation (5) * and cross-sectional area A2 * If we replace these with the cross-sectional areas that the main flow can substantially occupy at the first throat and the second throat, respectively, the ratio A2 * / A1 *The value of decreases. As a result, the total pressure loss caused by the shock wave decreases, or the shock wave becomes unnecessary. That is, the Mach number M1 upstream of the shock wave becomes smaller (in other words, the shock wave becomes weaker), or the shock wave disappears. This effect is achieved by injecting the secondary jet into a region downstream of the first throat. If the secondary jet is injected upstream from near the first throat, as in the thrust vector control devices of Non-Patent Documents 1 and 2, such an effect cannot be obtained. As will be described later, in a nozzle in which the cross-sectional area of ​​the second throat is larger than that of the first throat, sufficient thrust vectoring may not be obtained.

[0028] In the thrust direction control device of Patent Document 1, the position of the secondary jet injection port is determined based on the cross-sectional area A S In the vicinity of the point where S The location of the secondary jet injection port is limited to a location where the cross-sectional area is 0.9 to 1.1 times the cross-sectional area of ​​the secondary jet injection port. This is because it was predicted that the optimal location for the injection port would be near the point (P0) where the shock wave front intersects with the nozzle inner wall where the secondary jet injection port is to be installed. This is because, when a secondary jet is injected from an injection port installed in the diverging section, the shock wave moves upstream due to the mechanism described above, and the shock wave front either becomes located upstream of P0 or the shock wave itself disappears. In either case, if the secondary jet injection port is installed at P0 or downstream of P0, the flow becomes subsonic near the secondary jet injection port. On the other hand, as a general trend, the separation region becomes smaller as the secondary jet injection port is installed downstream, and the thrust vectoring effect weakens. Therefore, it is preferable to install the injection port as far upstream as possible. Based on these factors, it was predicted that the optimal location for the injection port would be at P0. The position of P0 is included in the shock wave front when no secondary jet is injected, so the cross-sectional area A can be calculated by simultaneously using the above equations 5, 6, and 7. S This coincides with the position where the nozzle cross-sectional area is equal to the nozzle diameter.

[0029] However, since the above formulas 5, 6, and 7 are derived based on approximations such as one-dimensional flow and isentropic flow, there is a possibility that they contain some errors when compared with the actual flow inside the nozzle. Therefore, taking into consideration the possibility of such errors, Patent Document 1 provides a width at the position of the injection port to calculate the cross-sectional area A S The cross-sectional area was set to 0.9 to 1.1 times the area of ​​the

[0030] As described above, the thrust vector control device of Patent Document 1 satisfies not only the first condition that "thrust is deflected when it should be deflected" which was already confirmed in Non-Patent Document 1 and Non-Patent Document 2, which represent the previous technologies, but also the second condition that "thrust is not deflected when it should not be deflected" which was not confirmed in Non-Patent Document 1 or Non-Patent Document 2. On the other hand, when the thrust vectoring performance, which is a requirement for the first condition to be satisfied, is quantitatively evaluated, the deflection performance of the thrust vector control device proposed in Patent Document 1 is somewhat inferior to the performance reported in Non-Patent Document 1 and Non-Patent Document 2. For example, in the case of the main stream F IN1 and secondary jet F IN2 Secondary jet F for the sum of the flow rates of IN2When the flow rate ratio (hereinafter sometimes referred to as the "flow rate ratio") is 5%, the magnitude of the thrust vectoring angle (referred to as "δ") calculated by the experiments and numerical simulations reported in Non-Patent Document 1 and Non-Patent Document 2 is approximately 20° at maximum, 16° at minimum, and 18° on average, whereas the magnitude of δ obtained by the numerical simulation performed by the inventor for the thrust vector control device of Patent Document 1 is approximately 17° at maximum, 7° at minimum, and 13° on average. This performance degradation is more pronounced when the flow rate ratio is small; when the flow rate ratio is 3%, the magnitude of the thrust vectoring angle δ reported in Non-Patent Document 1 and Non-Patent Document 2 is approximately 17° at maximum, 12° at minimum, and 14° on average, whereas the magnitude of the thrust vectoring angle δ obtained by the thrust vector control device of Patent Document 1 is approximately 10°. When the flow rate ratio was 2%, the magnitude of the thrust vector angle δ reported in Non-Patent Document 1 and Non-Patent Document 2 was approximately 14° at maximum, 10° at minimum, and 12° on average, whereas the magnitude of the thrust vector angle δ obtained with the thrust vector control device of Patent Document 1 was approximately 8°. Furthermore, when the flow rate ratio was 1%, the magnitude of the thrust vector angle δ reported in Non-Patent Document 1 and Non-Patent Document 2 was approximately 11° at maximum, 7° at minimum, and 9° on average, whereas the magnitude of δ obtained with the thrust vector control device of Patent Document 1 was approximately 5°. To summarize the above results, it is estimated that the magnitude of the thrust deflection angle δ by the thrust vector control device of Patent Document 1 decreases to approximately 74% of the magnitude of the thrust deflection angle δ in Non-Patent Documents 1 and 2 on average when the flow rate ratio is 5%, but this ratio decreases to approximately 71% when the flow rate ratio is 3%, approximately 66% when the flow rate ratio is 2%, and approximately 54% when the flow rate ratio is 1%, and the rate of decrease in the magnitude of the thrust deflection angle δ becomes more significant as the flow rate ratio decreases.

[0031] On the other hand, both experimental and numerical simulation results show that when the flow rate ratio exceeds 5%, the rate of increase in the magnitude of the thrust vector angle δ relative to the flow rate ratio significantly decreases, or the magnitude of the thrust vector angle δ hardly increases at all. Therefore, to prevent an increase in the power required to drive the thrust vector control device and to prevent the device from becoming too large and heavy, it is desirable to operate the device within a range where the flow rate ratio is as small as possible. Operation in a range where the flow rate ratio exceeds 5% is particularly impractical. Furthermore, a tendency can be discerned from the experimental results and numerical simulation results reported in Non-Patent Documents 1 and 2, as well as the numerical simulation results for the thrust vector control device in Patent Document 1, that the value obtained by dividing the magnitude of the thrust vector angle δ by the flow rate ratio (hereinafter sometimes referred to as "thrust vector efficiency") increases as the magnitude of the thrust vector angle δ decreases. For this reason, when designing various flying vehicles and aircraft incorporating thrust vector control devices, it is considered appropriate to primarily assume operation at flow rate ratios of 3% or less. Therefore, deflection performance in a range where the flow rate ratio is small is extremely important.

[0032] However, as mentioned above, when comparing the thrust vectoring performance of the thrust vector control device of Patent Document 1 with the performance reported in Non-Patent Documents 1 and 2, the rate of performance degradation becomes more pronounced as the flow rate ratio decreases. As a result, the thrust vectoring efficiency in that range is significantly reduced compared to when the thrust vector control devices of Non-Patent Documents 1 and 2 are used, which may lead to an increase in the power required to inject the secondary jet and the weight of the device in the thrust vector control device of Patent Document 1. Therefore, improving the vectoring performance in the range of small flow rate ratios (for example, a flow rate ratio of 3% or less) is a very important issue in order to improve the overall system performance (maneuverability, fuel efficiency, etc.) of flying vehicles and aircraft incorporating thrust vector control devices.

[0033] The present invention has been made to solve the above-mentioned problems, and provides a thrust direction control device that not only satisfies both the first condition that "thrust is deflected when it should be deflected" and the second condition that "thrust is not deflected when it should not be deflected," but also satisfies the first condition even with a secondary jet with a small flow rate. [Means for solving the problem]

[0034] The above issues are: A fluid type thrust vector control device that changes the direction of a jet flowing from a nozzle outlet by injecting a secondary jet from a secondary jet injection port provided in the middle of the nozzle into a main flow flow that flows from a nozzle inlet to a nozzle outlet, The nozzle is a first throttle portion having an upstream end serving as a nozzle inlet, a downstream end having a cross-sectional area smaller than that of the upstream end, and a downstream end serving as a first throat; a diverging portion connected to a downstream side of the first throttling portion, the diverging portion having a cross-sectional area at a downstream end thereof larger than a cross-sectional area at an upstream end thereof; a second throttle portion connected to the downstream side of the diverging portion, the cross-sectional area of ​​the downstream end of which is smaller than the cross-sectional area of ​​the upstream end of which, and the downstream end of which serves as a second throat; A dual throat nozzle having Cross-sectional area of ​​the first throat (A1 * ") relative to the cross-sectional area of ​​the second throat ("A2 * ") Ratio A2 * / A1 * By setting the value to 1.2 to 1.8, during non-deflection control in which only the main flow flows in without injecting a secondary jet into the nozzle, a shock wave is generated in the diverging section, and the flow becomes supersonic upstream of that, preventing the flow from separating from the nozzle inner wall surface, thereby preventing unintended thrust deflection. When the Mach number upstream of the shock wave is M1 and the specific heat ratio is γ, Ratio A2 * / A1 * Substituting the value of into the left side of the above equation 3, and solving the above equation 3 and the above equation 4 simultaneously, the cross-sectional area A S In contrast, The secondary jet inlet has a cross-sectional area A S The cross-sectional area is 0.85 times or less, and the cross-sectional area A obtained from the following formula 8 min Installed in a location with a cross-sectional area of ​​more than A fluid type thrust vector control device characterized by This is solved by providing

number

[0035] Here, the cross-sectional area A in the above formula 8 min is the cross-sectional area of ​​the nozzle divergence where the secondary jet inlet is provided ("A i ") means the lower limit of A in the above formula 8. S is obtained by simultaneously solving the above equation 3 and the above equation 4, and the cross-sectional area A min The value of the cross-sectional area A1 * and cross-sectional area A S The value of is divided internally by 3 to 7 on the number line, so the cross-sectional area is A min The points where the cross-sectional area of ​​the first throat and the diverging section is A S It is located between the points where

[0036] In Patent Document 1, the position of the secondary jet injection port is set to the cross-sectional area A S The reason for determining the location where the cross-sectional area is 0.9 to 1.1 times that of the secondary jet inlet has already been mentioned, but to reiterate, it is as follows: In other words, as a sufficient condition for the flow to become subsonic near the secondary jet inlet when the secondary jet is injected, the position of the secondary jet inlet is limited to "P0 or downstream of P0", and as a general trend, locating the inlet as far upstream as possible strengthens the thrust vectoring effect, so based on these two conditions, the position of the inlet is set to "the position of P0", and the cross-sectional area is set to A S Considering the influence of the position error between the point where SHowever, the logical structure of this basis does not exclude the possibility that a better thrust vectoring effect can be obtained when the secondary jet inlet is located at a position outside the above range. Based on this speculation, the inventors performed simulation calculations by setting the secondary jet inlet at various positions over a wider range in the diverging portion of the nozzle. They found that when the secondary jet inlet is located at a position upstream of the above range set in Patent Document 1 (a position where the nozzle cross-sectional area is smaller), better thrust vectoring performance can be obtained than when the secondary jet inlet is located at a position within the above range set in Patent Document 1, especially in cases where the flow rate ratio is small (for example, when the flow rate ratio is 3% or less). As already mentioned, improving thrust vectoring performance in the range where the flow rate ratio is small can be said to be a very important issue in terms of improving the overall system performance of flying vehicles and aircraft incorporating thrust vector control devices. As a means of solving this issue, in the present invention, the secondary jet inlet is located at a position upstream of the range set in Patent Document 1 (a position where the nozzle cross-sectional area is smaller than the above range). S The cross-sectional area is 0.85 times or less of the cross-sectional area A min The cross-sectional area of ​​the location where the secondary jet injection port is to be installed ("A i ") upper limit (0.85A S ) and the lower limit (A min ) is a cross-sectional area A in a case where the improvement of thrust vectoring performance was clearly confirmed in comparison with the thrust vector control device of Patent Document 1, especially in the region where the flow rate ratio is 3% or less. i The value of A min ≦A i ≦0.85×A S ) were selected to be included in the

[0037] Here, the secondary jet inlet is defined as the cross-sectional area A i is the cross-sectional area A S 0.85 times (=0.85×A S) or less, the thrust vectoring performance is improved compared to the thrust vector control device of Patent Document 1. This can be presumed to be because, as a general trend, the further upstream the secondary jet injection port is located, the wider the range in which the main flow separates from the nozzle inner wall, resulting in greater deflection. i min When the secondary jet inlet is located in a position where the secondary jet inlet is too close to the first throat, a phenomenon was observed in which the thrust vectoring performance was reduced, and the reason for this can be inferred as follows. That is, when the secondary jet inlet is located too close to the first throat, it was found that during the deflection control in which the secondary jet is injected, a place appears where the flow pipe (a virtual pipe as a collection of streamlines) surrounding the main flow is pushed by the secondary jet and appears to have a smaller cross-sectional area than the first throat (see Figure 34). i >A min Position R C Figure 36 shows the streamline diagram when an inlet is provided at R. Therefore, for the main flow, that location becomes the effective first throat, and the cross-sectional area ratio of the flow tube to the second throat (effective cross-sectional area of ​​the second throat / effective cross-sectional area of ​​the first throat) increases. As a result, as shown in Figure 35, C 37), the shock wave became stronger (i.e., the Mach number upstream of the shock wave became larger), which is thought to have suppressed the separation of the main flow and the thrust deflection.

[0038] As already mentioned, in the thrust direction control device of Patent Document 1, the cross-sectional area A1 of the first throat * than the cross-sectional area of ​​the second throat A2 * Increase the ratio A2 * / A1 * is set to be greater than 1) to satisfy the second condition that "the thrust is not deflected when it should not be deflected." In the thrust direction control device of the present invention, the ratio A2 * / A1 * ​is set to 1.2 to 1.8, which is greater than 1. As a result, during non-deflection control when no secondary jet is injected, a shock wave is generated in the diverging section downstream of the first throat, and the flow becomes supersonic upstream of that, thereby suppressing separation within the nozzle and preventing unintended deflection of thrust.

[0039] For reference, Fig. 38 shows the technical scope of the present invention (the area indicated by the diagonal hatching that slopes upward to the right) and the technical scope of the invention according to claim 1 of Patent Document 1 (the area indicated by the diagonal hatching that slopes downward to the right) in terms of the ratio A2 * / A1 * is taken on the horizontal axis, and the ratio A i / A S The figure shows the relationship between the specific heat ratio γ and the specific heat ratio γ on the vertical axis. In this figure, Equation 8 is expressed by two curves: the upper curve represents the case where the specific heat ratio γ is 5 / 3, and the lower curve represents the case where the specific heat ratio γ is 1.3. A specific heat ratio γ of 5 / 3 corresponds to the case where the main flow and secondary jet flowing through the nozzle are composed of monoatomic molecules, while a specific heat ratio γ of 1.3 corresponds to a typical case where the main flow and secondary jet flow are composed of a mixture of diatomic and triatomic molecules, or where the influence of internal molecular vibrations cannot be ignored due to high temperatures. The properties of gases ejected from the combustion chambers of actually operating rocket and jet engines can be accurately represented by setting the value of γ between the two values ​​mentioned above. As can be seen from Figure 38, there is no overlap between the technical scope of the present invention and the technical scope of the invention according to claim 1 of Patent Document 1.

[0040] In the fluid-type thrust-vector control device of the present invention, the nozzle is not particularly limited as long as it has the following sections provided in this order from the nozzle inlet to the nozzle outlet: a first throttle section, a first throat, a diverging section, a second throttle section, and a second throat. Another diverging section (second diverging section) may be provided further downstream of the second throat. Examples of nozzles include those with the same cross-sectional shape parallel to a plane including their centerline (so-called two-dimensional nozzles) and those shaped like a body of revolution with the centerline as its axis. [Effects of the Invention]

[0041] As described above, the present invention makes it possible to provide a thrust-vector control device that not only satisfies both the first condition that "thrust is deflected when it should be" and the second condition that "thrust is not deflected when it should not be deflected," but also satisfies the first condition even with a small flow rate of secondary jet. This not only makes it possible to reduce the size and weight of the thrust-vector control device, but also reduces the power required to drive the thrust-vector control device. [Brief explanation of the drawings]

[0042] [Figure 1] This is a conceptual diagram of a conventionally proposed dual throat nozzle (Nozzle A). [Figure 2] FIG. 1 is a diagram showing a two-dimensional nozzle used in the simulation. [Figure 3] FIG. 1 is a diagram illustrating simulated dimensions of a conventionally proposed dual throat nozzle (nozzle A). [Figure 4] This graph shows the results of a simulation to confirm the second condition conducted for a conventionally proposed dual throat nozzle (Nozzle A), and shows the change in thrust deflection angle over time for each random number sequence when only the main flow flows into the nozzle without injecting a secondary jet (non-deflection control). [Figure 5] This figure shows the results of a simulation to confirm the second condition conducted on a conventionally proposed dual throat nozzle (nozzle A). The figure shows the streamlines in the nozzle when only the main flow flows in without injecting a secondary jet into the nozzle (non-deflection control), for (a) the case of random number sequence 4 in Figure 4 and (b) the case of random number sequence 10 in Figure 4. [Figure 6] FIG. 1 is a conceptual diagram of a dual throat nozzle used in a fluidic thrust vector control device. [Figure 7] FIG. 1 is a perspective view showing an example of a three-dimensional shape of a dual throat nozzle used in a fluid-type thrust vector control device. [Figure 8] FIG. 1 is a diagram illustrating simulated dimensions of a dual throat nozzle (nozzle B) used in a fluid-type thrust vector control device. [Figure 9] This graph shows the results of a second condition confirmation simulation conducted on a dual throat nozzle (nozzle B) used in a fluidic thrust vector control device, and shows the change in thrust deflection angle over time for each random number sequence when only the main flow flows into the nozzle without injecting a secondary jet (non-deflection control). [Figure 10] This figure shows the results of a second condition confirmation simulation conducted on a dual throat nozzle (nozzle B) used in a fluidic thrust vector control device, and shows the streamlines inside the nozzle when only the main flow flows in without injecting a secondary jet into the nozzle (non-deflection control). [Figure 11] This figure shows the results of a second condition confirmation simulation conducted on a dual throat nozzle (nozzle B) used in a fluidic thrust vector control device, and shows the Mach number distribution inside the nozzle when only the main flow flows in without injecting a secondary jet into the nozzle (non-vectoring control). [Figure 12] FIG. 1 is a diagram illustrating simulated dimensions of a dual throat nozzle (nozzle C) used in a fluid type thrust vector control device. [Figure 13] This graph shows the results of a second condition confirmation simulation conducted on a dual throat nozzle (nozzle C) used in a fluidic thrust vector control device, and shows the change in thrust deflection angle over time for each random number sequence when only the main flow flows into the nozzle without injecting a secondary jet (non-deflection control). [Figure 14] This figure shows the results of a second condition confirmation simulation conducted on a dual throat nozzle (nozzle C) used in a fluidic thrust vector control device, and shows the streamlines inside the nozzle when only the main flow flows in without injecting a secondary jet into the nozzle (non-deflection control). [Figure 15]This figure shows the results of a second condition confirmation simulation conducted on a dual throat nozzle (nozzle C) used in a fluidic thrust vector control device, and shows the Mach number distribution inside the nozzle when only the main flow flows in without injecting a secondary jet into the nozzle (non-vectoring control). [Figure 16] FIG. 1 is a diagram illustrating simulated dimensions of a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device. [Figure 17] FIG. 1 is a diagram illustrating the setting position of the secondary jet injection port used in the first condition confirmation simulation (part 1: injection from near the first throat) performed on a dual throat nozzle (nozzle B) used in a fluid-type thrust vector control device. [Figure 18] FIG. 1 is a diagram explaining the setting position of the secondary jet injection port used in the first condition confirmation simulation (Part 1: injection from near the first throat) performed on a dual throat nozzle (Nozzle C) used in a fluid-type thrust vector control device. [Figure 19] FIG. 1 is a diagram illustrating the setting position of the secondary jet injection port used in the first condition confirmation simulation (part 1: injection from near the first throat) performed on a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device. [Figure 20] This is a graph showing the results of a first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle B) used in a fluid-type thrust direction control device, and shows the change over time in the thrust deflection angle for each random number sequence when a secondary jet is injected from position P1 in Figure 17 (during deflection control). [Figure 21] This figure shows the results of the first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle B) used in a fluid-type thrust vector control device, and shows the streamlines inside the nozzle when a secondary jet is injected from position P1 in Figure 17 (during deflection control). [Figure 22]This figure shows the results of the first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle B) used in a fluid-type thrust vector control device, and shows the Mach number distribution inside the nozzle when a secondary jet is injected from position P1 in Figure 17 (during deflection control). [Figure 23] This is a graph showing the results of a first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle C) used in a fluid-type thrust direction control device, and shows the change over time in the thrust deflection angle for each random number sequence when a secondary jet is injected from position Q1 in Figure 18 (during deflection control). [Figure 24] This figure shows the results of a first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle C) used in a fluid-type thrust vector control device, and shows the streamlines inside the nozzle when a secondary jet is injected from position Q1 in Figure 18 (during deflection control). [Figure 25] This figure shows the results of a first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle C) used in a fluid-type thrust vector control device, and shows the Mach number distribution inside the nozzle when a secondary jet is injected from position Q1 in Figure 18 (during deflection control). [Figure 26] This is a graph showing the results of a first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle D) used in a fluid-type thrust direction control device, and shows the change over time in the thrust deflection angle for each random number sequence when a secondary jet is injected from position R1 in Figure 19 (during deflection control). [Figure 27]This figure shows the results of a first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device, and shows the streamlines inside the nozzle and outside the nozzle outlet when a secondary jet is injected from position R1 in Figure 19 (during deflection control). [Figure 28] This figure shows the results of a first condition confirmation simulation (part 1: injection from near the first throat) conducted on a dual throat nozzle (nozzle D) used in a fluidic thrust vector control device, and shows the Mach number distribution inside the nozzle and outside the nozzle exit when a secondary jet is injected from position R1 in Figure 19 (during deflection control). [Figure 29] FIG. 1 is a diagram illustrating the setting position of the secondary jet injection port used in the first condition confirmation simulation (part 2: injection into the nozzle divergence section) performed on a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device. [Figure 30] FIG. 1 shows the results of a first condition confirmation simulation (part 2: injection into the nozzle divergence section) performed on a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device. [Figure 31] This figure shows the results of the first condition confirmation simulation (part 2: injection into the nozzle divergence section) conducted for a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device, and shows the Mach number distribution and streamlines inside the nozzle and outside the nozzle exit when a secondary jet is injected at a flow rate ratio of 1% from injection position RB in the nozzle divergence section. [Figure 32] This figure shows the results of the first condition confirmation simulation (part 2: injection into the nozzle divergence section) conducted for a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device, and shows the Mach number distribution and streamlines inside the nozzle and outside the nozzle exit when a secondary jet is injected at a flow rate ratio of 2% from injection position RB in the nozzle divergence section. [Figure 33]This figure shows the results of the first condition confirmation simulation (part 2: injection into the nozzle divergence section) conducted for a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device, and shows the Mach number distribution and streamlines inside the nozzle and outside the nozzle exit when a secondary jet is injected at a flow rate ratio of 4% from injection position RB in the nozzle divergence section. [Figure 34] This figure shows the results of a first condition confirmation simulation (part 2: injection into the nozzle diverging section) conducted on a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device, and shows the streamlines inside the nozzle and outside the nozzle outlet when a secondary jet is injected from injection position RA within the nozzle diverging section. [Figure 35] This figure shows the results of the first condition confirmation simulation (part 2: injection into the nozzle divergence section) conducted on a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device, and shows the Mach number distribution inside the nozzle and outside the nozzle exit when a secondary jet is injected from injection position RA within the nozzle divergence section. [Figure 36] This figure shows the results of the first condition confirmation simulation (part 2: injection into the nozzle diverging section) conducted on a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device, and shows the streamlines inside the nozzle and outside the nozzle outlet when a secondary jet is injected from injection position RC within the nozzle diverging section. [Figure 37] This figure shows the results of the first condition confirmation simulation (part 2: injection into the nozzle divergence section) conducted on a dual throat nozzle (nozzle D) used in a fluid-type thrust vector control device, and shows the Mach number distribution inside the nozzle and outside the nozzle exit when a secondary jet is injected from injection position RC within the nozzle divergence section. [Figure 38] 1 is a diagram showing the technical scope of the present invention and the technical scope of the invention according to claim 1 of Patent Document 1, with the ratio A2* / A1* on the horizontal axis and the ratio Ai / AS on the vertical axis. DETAILED DESCRIPTION OF THE INVENTION

[0043] 1. Overview of the fluid thrust vector control device of the present invention A preferred embodiment of the fluidic thrust-vector control device of the present invention will be described in more detail with reference to the drawings. Figure 6 is a conceptual diagram of a dual throat nozzle 10 used in the fluidic thrust-vector control device of the present invention. In the figure, reference numeral "10" denotes the nozzle, reference numeral "10a" denotes the centerline of the nozzle, reference numeral "11" denotes the nozzle inlet, reference numeral "12" denotes the first throttle portion, reference numeral "13" denotes the first throat, reference numeral "14" denotes the diverging portion, reference numeral "15" denotes the second throttle portion, reference numeral "16" denotes the second throat, reference numeral "18" denotes the nozzle outlet, reference numeral "19" denotes the secondary jet inlet, reference numeral "D1" denotes the nozzle inlet opening width, and reference numeral "F IN1 " represents the main flow from the engine combustion chamber to the nozzle inlet, and "F IN2 " indicates the secondary jet injected from the secondary jet inlet, and is indicated by the symbol "F OUT " indicates the jet flow coming out of the nozzle outlet.

[0044] FIG. 6 above depicts a cross-sectional view of the nozzle 10 cut along a plane (parallel to the xz plane) including the centerline 10a. FIG. 7 is a perspective view showing an example of the three-dimensional shape of the dual throat nozzle 10 used in the fluidic thrust-vector control device of the present invention. Examples of the three-dimensional shape of the nozzle 10 include a nozzle 10 having the same shape in any cross section perpendicular to the y-axis direction (a so-called two-dimensional nozzle), as shown in FIG. 7(a), and a nozzle 10 having the shape of a body of revolution with the centerline 10a as its axis (a so-called axisymmetric nozzle), as shown in FIG. 7(b). A nozzle such as this axisymmetric nozzle, in which the shape of any cross section perpendicular to the y-axis direction is not the same, is sometimes called a "three-dimensional nozzle" in contrast to the two-dimensional nozzle described above.

[0045] As shown in FIG. 6, the fluid type thrust vector control device of the present invention is configured by a main flow F flowing from a nozzle inlet 11 to a nozzle outlet 18 in a nozzle 10. IN1 On the other hand, secondary jet F IN2 By injecting IN1 This causes a change in the flow of the jet F OUTIt is possible to change the direction of the

[0046] In the nozzle 10, a first throttle section 12 is formed downstream of the nozzle inlet 11. This first throttle section 12 is a portion formed by being narrowed from the upstream side (negative side in the x-axis direction) toward the downstream side (positive side in the x-axis direction), and the cross-sectional area (opening area) of the downstream end of the first throttle section 12 is smaller than the cross-sectional area (opening area) of the upstream end of the first throttle section 12. The downstream end of the first throttle section 12 forms a first throat 13.

[0047] Furthermore, a diverging section 14 is formed downstream of the first throttling section 12 in the nozzle 10. This diverging section 14 is formed by widening from the upstream side toward the downstream side, and the cross-sectional area (opening area) of the downstream end of the diverging section 14 is larger than the cross-sectional area (opening area) of the upstream end of the diverging section 14.

[0048] Furthermore, a second throttle section 15 is formed downstream of the diverging section 14 in the nozzle 10. Similar to the first throttle section 12, this second throttle section 15 is formed by being narrowed from the upstream side to the downstream side, and the cross-sectional area (opening area) of the downstream end of the second throttle section 15 is narrower than the cross-sectional area (opening area) of the upstream end of the second throttle section 15. The downstream end of the second throttle section 15 forms a second throat 16. In the nozzle 10, this second throat 16 portion forms a nozzle outlet 18.

[0049] In the dual throat nozzles proposed in the past, the cross-sectional area of ​​the second throat 16 is set equal to the cross-sectional area of ​​the first throat 13. In contrast, in the dual throat nozzle 10 used in the fluidic thrust vector control device of the present invention, as shown in FIG. * The cross-sectional area A1 of the first throat 13 * Specifically, the cross-sectional area A1 of the first throat 13 is set to be larger than * Cross-sectional area A2 of the second throat 16 * Ratio A2 * / A1 * Therefore, the secondary jet F IN2 Mainstream F without injection IN1 During non-vectoring control, in which only thrust is introduced, a shock wave is generated in the diverging section 14 across the nozzle 10, and the flow between the shock wave and the first throat 13 becomes supersonic, so the above-mentioned "expansion mechanism of the supersonic flow in the diverging section 14" makes it possible to prevent the flow from separating from the inner wall surface of the nozzle 10. Therefore, the second condition that the thrust is not deflected when it should not be deflected (during non-vectoring control) is satisfied.

[0050] In addition, a secondary jet F is formed in the peripheral wall of the nozzle 10. IN2 The secondary jet injection port 19 is provided at a location in the diverging portion 14 away from the first throat 13. Specifically, the ratio A2 * / A1 * Substituting the value of into the left side of the above equation 3, and solving the above equation 3 and the above equation 4 simultaneously, the cross-sectional area A S The cross-sectional area is 0.85 times or less, and the cross-sectional area A obtained from the above formula 8 min A secondary jet injection port 19 is provided at a location where the cross-sectional area is equal to or greater than this.

[0051] The secondary jet flow inlet 19 is connected to a gas transfer means (not shown), and a gas transfer path connecting the gas transfer means and the secondary jet flow inlet 19 is provided with an on-off valve for opening and closing the gas transfer path. IN1 is supplied from the engine combustion chamber, the gas transfer path from this engine combustion chamber is the main stream F IN1 and a secondary jet F IN2 It may be branched into a secondary jet transport for transporting the above.

[0052] The number of secondary jet injection ports 19 to be provided is not particularly limited. However, if only one secondary jet injection port 19 is provided, the thrust can only be deflected to the same side. In contrast, if multiple secondary jet injection ports 19 are provided, the secondary jet F IN2 By switching the secondary jet injection port 19 into which the secondary jet F is injected, it is possible to deflect the thrust to a different direction. For example, in the case of the two-dimensional nozzle shown in FIG. 7(a), as shown in FIG. 6, a lower secondary jet injection port 19a and an upper secondary jet injection port 19b are provided as the secondary jet injection ports 19, and the secondary jet F is injected from either one of the secondary jet injection ports 19 depending on the direction in which the thrust is to be deflected. IN2 This makes it possible to change the direction of the thrust not only to the positive z-axis direction but also to the negative z-axis direction. The angle at which the thrust is deflected can be adjusted by adjusting the opening of the on-off valve in the gas transfer path to change the secondary jet F IN2 This can be adjusted by adjusting the flow rate of the

[0053] The principle of thrust deflection by the fluidic thrust vector control device equipped with the nozzle 10 is as follows: In other words, in the fluidic thrust vector control device of the present invention, combustion gas delivered from an engine combustion chamber (such as a combustion chamber of a rocket engine or a jet engine) (not shown) is deflected by the main flow F. IN1 The jet F flows into the nozzle 10 from the nozzle inlet 11. OUT The jet F is ejected from the nozzle outlet 18. OUT The reaction of the jet F OUT The engine receives a force in the opposite direction to the total momentum of the rocket, generating thrust.

[0054] Here, when the thrust is not deflected (non-deflection control), a secondary jet F IN2 Mainstream F without injection IN1 In the nozzle 10 used in the fluid type thrust vector control device of the present invention, when non-deflection control is performed, only the jet flow F indicated by the dashed arrow in FIG. OUT As shown, the jet F OUTThe engine is designed so that the momentum of the entire engine faces one side (positive x-axis direction) parallel to the center line 10a of the nozzle 10. As a result, the engine generates thrust in the opposite direction (negative x-axis direction).

[0055] On the other hand, when it is desired to tilt the direction of the thrust to the positive side in the z-axis direction (when it is desired to deflect the thrust in the upper left direction on the paper surface of FIG. 6), a secondary jet F is injected into the nozzle 10 from the secondary jet injection port 19a on the lower side in the same figure. IN2 Then, the mainstream F IN1 1, is pushed upward (positive side in the z-axis direction) within the nozzle 10, but is bent downward (negative side in the z-axis direction) by the second throttle portion 15 so as to face in a downward and rightward direction. OUT The total momentum is the jet F indicated by the black arrow in Figure 6. OUT As a result, the thrust of the engine is deflected in the upper left direction on the paper in Figure 6.

[0056] On the other hand, when it is desired to tilt the direction of the thrust to the negative side in the z-axis direction (when it is desired to deflect the thrust to the lower left direction on the paper surface of FIG. 6), a secondary jet F is injected into the nozzle 10 from the secondary jet injection port 19b on the upper side in the same figure. IN2 As a result, the secondary jet F flows from the lower secondary jet injection port 19a. IN2 By the same principle as when a thrust force is injected, the thrust force is deflected in the downward left direction on the paper surface of FIG.

[0057] The above description has been given for the case where the nozzle 10 is the two-dimensional nozzle of Fig. 7(a), but the same applies to the case where the nozzle 10 is the axisymmetric nozzle of Fig. 7(b). However, when the nozzle 10 is an axisymmetric nozzle, it is preferable to provide secondary jet injection ports 19 at at least three locations within a range exceeding 180° around the center line 10a of the nozzle 10. This allows the secondary jets F injected from each secondary jet injection port 19 to be uniformly distributed. IN2 By adjusting the flow rate, it is possible to change the direction of the thrust to the positive y-axis direction or the negative y-axis direction.

[0058] 2. Second condition confirmation simulation In order to confirm whether the dual throat nozzle 10 related to the fluidic thrust vector control device satisfies the second condition of "not deflecting when it should not be deflected," a simulation (second condition confirmation simulation) was performed. Specifically, in the nozzle 10 related to the fluidic thrust vector control device (FIG. 6), the secondary jet F IN2 Mainstream F without injection IN1 The flow field and thrust deflection angle in the dual throat nozzle 10 when only the air flows in (non-deflection control) were calculated by numerical simulation.

[0059] The analysis method and conditions for the second condition confirmation simulation conducted for the dual throat nozzle 10 of the fluidic thrust-vector control device are the same as those for the conventionally proposed dual throat nozzle (Nozzle A) described in the "Problems to be Solved by the Invention" section, except for the dimensions of the nozzle 10. Specifically, for the conventionally proposed dual throat nozzle 10, the second condition confirmation simulation was conducted by setting the nozzle 10 to the dimensions shown in FIG. 3 (the cross-sectional areas of the first throat 13 and the second throat 16 were set equal). In contrast, for the dual throat nozzle 10 of the fluidic thrust-vector control device of the present invention, the second condition confirmation simulation was conducted by setting the nozzle 10 to two different dimensions, as shown in FIGS. 8 and 12. Hereinafter, the dual throat nozzle 10 with the dimensions shown in FIG. 8 may be referred to as "Nozzle B," and the dual throat nozzle 10 with the dimensions shown in FIG. 12 may be referred to as "Nozzle C." In both "Nozzle B" and "Nozzle C" of the fluidic thrust-vector control device of the present invention, the cross-sectional area of ​​the second throat 16 is set to be larger than the cross-sectional area of ​​the first throat 13. In nozzle B, the cross-sectional area A1 of the first throat 13 * Cross-sectional area A2 of the second throat 16 * Ratio A2 * / A1 * is about 1.7, and for nozzle C, the ratio A2 * / A1 *is approximately 1.3.

[0060] 2.1 In the case of "Nozzle B" Figures 9, 10, and 11 show the results of the simulation to confirm the second condition conducted for "Nozzle B" (see Figure 8). The graph in Figure 9 shows the time variation of the thrust vectoring angle δ during non-vectoring control for each random number sequence. Figure 10 shows the streamlines inside the nozzle during non-vectoring control. Figure 11 shows the Mach number distribution inside the nozzle during non-vectoring control.

[0061] Ratio A2 * / A1 * In "Nozzle A" with a ratio of 1, as shown in Figure 4, the thrust deflection angle δ varied greatly on the positive or negative side depending on the random number sequence, and its absolute value reached a maximum of approximately 4°. * / A1 * For "Nozzle B," where R is approximately 1.7, as shown in Figure 9, there is no significant variation in the thrust deflection angle δ due to the random number sequence, and the thrust deflection angle δ remains steady at a value close to 0° for all random number sequences.

[0062] Also, the ratio A2 * / A1 * In "Nozzle A" with a value of 1, as shown in Figure 5, the streamlines in the nozzle (streamlines in the nozzle at t = 37.5) in random number sequence 4 and random number sequence 10 (random number sequence that is significantly deflected in Figure 4) are formed asymmetrically in the vertical direction, and the secondary jet F IN2 Even though the non-deflection control was performed without injecting oxygen, a large separation region was formed. * / A1 * In "Nozzle B" where t is approximately 1.7, as shown in Figure 10, the streamlines in the nozzle (streamlines in the nozzle at t = 37.5) are formed almost symmetrically in the vertical direction, and no separation region is observed.

[0063] Furthermore, the ratio A2 * / A1 *In "Nozzle A" with a ratio of 1, the flow was subsonic or transonic throughout the entire area of ​​the dual throat nozzle 10, regardless of whether the random number sequence was 4 or 10. In contrast, the ratio A2 * / A1 * In "Nozzle B" with a σ of approximately 1.7, as shown in Figure 11, a shock wave was formed downstream of the center of the diverging section 14 (at a location close to the downstream end of the diverging section 14), and the flow rate was supersonic over a wide range in the diverging section 14 (a wide range from the first throat 13 to the shock wave).

[0064] From the above results, the ratio A2 * / A1 * It was found that the second condition, "no deflection when it should not be," was met for "Nozzle B," where σ is approximately 1.7.

[0065] 2.2 In the case of "Nozzle C" In Figures 13, 14 and 15, the ratio A2 * / A1 * This figure shows the results of a simulation to confirm the second condition, conducted for "Nozzle C" (see Figure 12) with a Mach number of approximately 1.3. The graph in Figure 13 shows the time variation of the thrust vectoring angle δ during non-vectoring control for each random number sequence. Figure 14 shows the streamlines inside the nozzle during non-vectoring control. Figure 15 shows the Mach number distribution inside the nozzle during non-vectoring control.

[0066] As already mentioned, the ratio A2 * / A1 * In "Nozzle C" where the ratio A2 is about 1.3, * / A1 * The ratio A2 is about 1.7 compared to "Nozzle B". * / A1 * Although the ratio is smaller, the ratio of "Nozzle A" (ratio A2 * / A1 *is larger than 1). Looking at Figure 13, in "Nozzle C," as in "Nozzle B," the thrust deflection angle δ reaches a steady state at a value close to 0° for all random number sequences. However, in "Nozzle C," the range of variation in random number sequences 1 to 10 is slightly larger than in "Nozzle B."

[0067] 14, it can be seen that in "Nozzle C," as in "Nozzle B," the streamlines within the nozzle (streamlines within the nozzle at t=37.5) are formed approximately symmetrically in the vertical direction, and no separation region is observed. Furthermore, in "Nozzle C," as in "Nozzle B," a shock wave is formed in the diverging section 14, and a region where the flow becomes supersonic is formed in the diverging section 14. However, in "Nozzle B," a shock wave is formed downstream of the center of the diverging section 14, whereas in "Nozzle C," a shock wave is formed slightly upstream of the center of the diverging section 14.

[0068] From the above results, it was found that "Nozzle C" according to the fluid type thrust vector control device of the present invention also satisfies the second condition of "not deflecting when it should not be deflected." However, "Nozzle C" is somewhat disadvantageous in terms of satisfying the second condition, for example, the variation in the thrust deflection angle δ is slightly larger than that of "Nozzle B." For this reason, the ratio A2 * / A1 * If is set too small (too close to 1), it is expected that the second condition will be difficult to satisfy. * / A1 * It was found that it is preferable to set the value to 1.2 or more.

[0069] 2.3 Summary of the second condition confirmation simulation From the above simulation to confirm the second condition, in order for the dual throat nozzle to satisfy the second condition that "thrust is not deflected when it should not be deflected," the cross-sectional area A2 of the second throat 16 must be * The cross-sectional area A1 of the first throat 13 * It was confirmed that it is effective to make the cross-sectional area A1* Cross-sectional area A2 * Ratio A2 * / A1 * Since the second condition is met when the ratio is set to about 1.3 and about 1.7, the ratio A2 * / A1 * It was also found that the second condition can be satisfied if is around 1.2 to 1.8.

[0070] 3. Simulation to confirm the first condition (Part 1: Injection from near the first throat) Next, in order to confirm whether the dual throat nozzle 10 related to the fluid type thrust vector control device satisfies the first condition that "deflects when it should deflect," in addition to the above "Nozzle B" and "Nozzle C," a new nozzle A2 * / A1 * A simulation (first condition confirmation simulation) was carried out for "Nozzle D" in which the value of is set to 1.5, which is the intermediate value between "Nozzle B" and "Nozzle C." The dimensions of "Nozzle D" are shown in FIG. 16. Specifically, in the nozzle 10 (FIG. 6) related to the fluidic thrust vector control device of the present invention, the main flow F IN1 and secondary jet F IN2 The flow field and thrust deflection angle in the dual throat nozzle 10 when the secondary jet F was injected (during deflection control) were calculated by numerical simulation. IN2 The flow rate of the main stream F IN1 and secondary jet F IN2 The secondary jet F was set to 5% of the total flow rate of the IN2 The injection direction was a direction rotated 150° counterclockwise from the rightward direction (positive side in the x-axis direction) parallel to the nozzle center line 10a.

[0071] In this first condition verification simulation (part 1), the secondary jet F IN2 The position where the secondary jet F was injected (the position of the secondary jet injection port 19) was set near the first throat 13 (see the secondary jet injection port 19a in FIG. 3), similar to the first condition confirmation simulation using "Nozzle A" (see the description in the "Problem to be Solved by the Invention" section above). Specifically, for Nozzle B, the secondary jet FIN2 The position where the secondary jet F is injected is the position P1 in FIG. 17, the position Q1 in FIG. 18 for nozzle C, and the position R1 in FIG. 19 for nozzle D. In FIG. 17, FIG. 18, and FIG. 19, the secondary jet F is injected into each nozzle. IN2 Mainstream F without injection IN1 The figure shows the Mach number distribution when only the flow of air is allowed in. Other than that, the analysis method and conditions of the simulation are the same as those of the second condition confirmation simulation above.

[0072] 3.1 Secondary jet F from P1 at "Nozzle B" IN2 When injected 20, 21, and 22 show the secondary jet F from P1 in FIG. 17 in "Nozzle B" (see FIG. 8). IN2 This figure shows the results of a simulation to confirm the first condition when injecting. The graph in Figure 20 shows the time variation of the thrust vector angle δ during vector control for each random number sequence. Figure 21 shows the streamlines inside the nozzle during vector control. Figure 22 shows the Mach number distribution inside the nozzle during vector control.

[0073] Looking at Figure 20, the secondary jet F IN2 Even though the thrust deflection control is being performed while injecting the secondary jet F, the thrust deflection angle δ is close to 0° in all random number sequences, and it can be seen that the thrust direction is hardly deflected. From this, it can be seen that in "Nozzle B", the secondary jet F IN2 It was found that the first condition is difficult to satisfy when a configuration in which

[0074] The reason for this can be understood from Figures 21 and 22. In other words, looking at Figure 21, the secondary jet F IN2 By injecting the secondary jet F, the flow in the nozzle is separated from the nozzle wall. IN2 The separation is eliminated near the injection point of the secondary jet F in "Nozzle B". IN2When no secondary jet F is injected, as shown in FIG. 11, a shock wave is formed near the downstream end of the diverging section 14, and most of the diverging section 14 is in a supersonic region, creating an environment where flow separation is easily eliminated. IN2 As mentioned in the above "Problems to be Solved by the Invention," even when the secondary jet F is injected, the shock wave does not weaken or disappear, so the supersonic region does not shrink (rather, it expands). In fact, looking at Figure 22, it is clear that the secondary jet F IN2 It can be seen that most of the diverging portion 14 is a supersonic region, even though the gas is injected.

[0075] 3.2 Secondary jet F from Q1 at "Nozzle C" IN2 When injected 23, 24, and 25 show the secondary jet F from Q1 in FIG. 18 in "Nozzle C" (see FIG. 12). IN2 This figure shows the results of a simulation to confirm the first condition when injecting. The graph in Figure 23 shows the time change in thrust vector angle δ during vector control for each random number sequence. Figure 24 shows the streamlines inside the nozzle during vector control. Figure 25 shows the Mach number distribution inside the nozzle during vector control.

[0076] Looking at Figure 24, the secondary jet F IN2 It can be seen that the flow separation that occurred near the injection position of the secondary jet F is maintained up to the nozzle exit 18. Furthermore, looking at Figure 25, it can be seen that the secondary jet F IN2 It can also be seen that the flow becomes subsonic near the injection point. As a result, the thrust vectoring angle δ is in a steady state at a value close to -12° for all random number sequences, as shown in Figure 23. This indicates that the cross-sectional area A2 of the second throat 16 is * The cross-sectional area A1 of the first throat 13 * When it is larger than (ratio A2 * / A1 * is greater than 1), and the secondary jet F IN2Even when the fuel is injected from the vicinity of the first throat 13, the ratio A2 * / A1 * It was confirmed that the first condition is satisfied when the flow rate ratio is set to about 1.3. On the other hand, the magnitude (absolute value) of the thrust vectoring angle δ of the thrust vector control device reported in Non-Patent Documents 1 and 2 is approximately 20° at maximum, 16° at minimum, and 18° on average when the flow rate ratio is 5% (however, these reports use nozzles in which the cross-sectional areas of the first and second throats are equal, so it is presumed that the second condition is not satisfied). Therefore, the magnitude of the thrust vectoring angle δ in the steady state when a secondary jet with a flow rate ratio of 5% is injected from Q1 in "Nozzle C" is only about 67% of that in the cases reported in Non-Patent Documents 1 and 2.

[0077] 3.3 Secondary jet F from R1 at "Nozzle D" IN2 When injected 26, 27, and 28 show the secondary jet F from R1 in FIG. 19 in "Nozzle D" (see FIG. 16). IN2 This figure shows the results of a simulation to confirm the first condition when an external computational domain is injected. The graph in Figure 26 shows the time variation of the thrust deflection angle δ during deflection control for each random number sequence. Figure 27 shows the streamlines inside the nozzle during deflection control. Figure 28 shows the Mach number distribution inside the nozzle during deflection control. Note that in the simulation calculation for "Nozzle D," a square computational domain was also set outside the nozzle exit, and Figures 27 and 28 also show a portion of the flow field outside the nozzle. The reason for setting up a new computational domain is to accurately reflect the effect of backflow molecules (molecules moving in the opposite direction at a thermal velocity greater than the macroscopic flow velocity) entering from the nozzle exit in the simulation calculation. However, when the simulation results with and without an external computational domain were compared for multiple cases, almost no quantitative difference was observed.

[0078] As can be seen from Figure 26, the thrust deflection angle δ in the steady state was approximately -5° for all random number sequences, which was only approximately 28% of the cases reported in Non-Patent Documents 1 and 2. From this, it can be seen that in "Nozzle D", the secondary jet F IN2 It was found that the first condition is difficult to satisfy when a configuration in which

[0079] Looking at Figure 27, it can be inferred that the main flow is pushed by the secondary jet, but then expands again near the center of the diverging section, preventing the separation region from expanding significantly, which is why thrust vectoring remains slight. Looking at Figure 28, it can be inferred that the flow is supersonic in about half of the diverging section, which is why the expansion of the separation region is suppressed.

[0080] 3.4 Summary of the first condition confirmation simulation (Part 1: injection from near the first throat) By the above simulation to confirm the first condition, in the dual throat nozzle, the cross-sectional area A2 of the second throat 16 * The cross-sectional area A1 of the first throat 13 * When the second condition is satisfied, the secondary jet F IN2 It has been found that the first condition is difficult to satisfy when the cross-sectional area A1 * Cross-sectional area A2 * Ratio A2 * / A1 * When is equal to or greater than 1.5, the magnitude (absolute value) of the thrust vectoring angle δ becomes approximately 0%, down from approximately 28% in the cases reported in Non-Patent Document 1 and Non-Patent Document 2 (which use a nozzle in which the second condition is presumed not to be satisfied).

[0081] 4. First condition confirmation simulation (Part 2: Injection into the nozzle diverging part) Next, the secondary jet inlet 19 was set at various positions in the nozzle divergence section, and the flow rate ratio (main flow F IN1 The flow rate r1 and secondary jet F IN2 The secondary jet F for the sum of the flow rate r and r IN2 The ratio of the flow rate r2 (r2 / (r1+r2)) was also set to various values, and the simulation to confirm the first condition was performed. Here, the nozzle shape is the cross-sectional area ratio A2 * / A1 * The nozzle D was used, which has a value of 1.5 (the median value of the technical range of 1.2 to 1.8 of the present invention). Specifically, the position (injection position) of the secondary jet injection port 19 was set to R A (A i / A S = 0.646) as the injection position, and B (A i / A S = 0.738) as the injection position, and C (A i / A S = 0.830) as the injection position, and D (A i / A S = 0.923) is used as the injection position, and R E (A i / A S =1.107) is used as the injection position, and the flow rate ratio r2 / (r1+r2)(=r s The deflection angle δ was calculated by numerical simulation when the injection position R D is the secondary jet F in the simulation calculation when "Nozzle D" is used. IN2 This coincides with the position where the shock wave front intersects with the nozzle inner wall when no gas is injected ("P0" in the explanation of "Problem to be Solved by the Invention"). According to theories using one-dimensional flow approximation, P0 (i.e., R D ) is the nozzle cross-sectional area in the diverging section. S However, as already mentioned, there is some error in the actual flow field, so R D Nozzle cross-sectional area A i is AS FIG. 29 is a diagram for explaining the setting position of the secondary jet inlet 19 used in the first condition confirmation simulation (part 2). The Mach number distribution shown is the secondary jet F IN2 This is when no R is injected. A Cross section A at i is R D Cross section A at i It is 0.7 times the R B Cross section A at i is R D Cross section A at i It is 0.8 times the R C Cross section A at i is R D Cross section A at i It is 0.9 times the R E Cross section A at i is R D Cross section A at i This is 1.2 times the previous figure.

[0082] Secondary jet F into the nozzle divergence section IN2 The injection direction of the thrust vector δ differs from that of injection from the vicinity of the first throat 13. Instead, it is rotated 120° counterclockwise from the rightward direction (positive x-axis direction) parallel to the nozzle centerline 10a in Figure 1 toward the page. The rotation angle was changed from 150° to 120° because the lower wall surface of the nozzle divergence section in Figure 1 is inclined downward to the right, and it is considered unrealistic to have an angle between the wall surface and the injection direction that is too small when considering actual operation. Preliminary calculations confirmed that the magnitude (absolute value) of the thrust vector angle δ decreases as the rotation angle decreases. However, the decrease is relatively gradual within the range of rotation angles above 120°. Therefore, the rotation angle was set to 120°, the lower limit of that range.

[0083] Injection position R A ,R B ,R C ,R D ,R E In the figure of Figure 38 (ratio A2 * / A1* is taken on the horizontal axis, and the ratio A i / A S As can be seen from Figure 38, the injection position R B ,R C The injection position R is within the scope of the present invention. D However, this falls within the technical scope of the invention according to claim 1 of Patent Document 1.

[0084] Figure 30 shows the results of the first condition confirmation simulation (part 2: injection into the nozzle diverging part). A ,R B ,R C ,R D ,R E In either case, the flow rate ratio r s It can be seen that as the injection position R increases, the absolute value of the deflection angle δ |δ| also tends to increase. A ,R B ,R C ,R D ,R E In both cases, the deflection angle δ appears to saturate at a maximum value of 14 to 16 degrees. However, the flow rate ratio r s In the region where R is 3% or less, A ,R B ,R C ,R D ,R E The absolute value of the deflection angle δ |δ| varies greatly depending on the injection position R B ,R C In this case, the flow rate ratio r s When the absolute value |δ| is 2%, it is 10.5 to 12.7°, and the flow rate ratio r s The absolute value |δ| when the angle is 3% is 14.0 to 15.0°. A ,R D ,R E Then, the flow rate ratio r s When |δ| is 4.1 to 7.8°, the flow rate ratio r s When |δ| is 3%, the absolute value is 5.9 to 10.3°.B ,R C The absolute value |δ| at injection position R B So, the flow rate ratio r is only 2%. s As mentioned in "Problems to be Solved by the Invention," in the design of various flying objects and aircraft incorporating thrust vector control devices, it is considered appropriate to assume that they will be operated mainly in a region where the flow rate ratio is relatively small (for example, 3% or less). B Or R C When the flow rate ratio is set to 2% and 3%, the improvement in deflection performance may make an important contribution to improving the performance (maneuverability, fuel efficiency, etc.) of the entire flying object or aircraft system.

[0085] In addition, the deflection performance reported in Non-Patent Documents 1 and 2 for a dual throat nozzle in which the cross-sectional areas of both throats are the same is s When the flow rate ratio r is 2%, the absolute value of the deflection angle |δ| is approximately 14° at maximum, 10° at minimum, and 12° on average. s When |δ| was 3%, the absolute value was approximately 17° at maximum, 12° at minimum, and 14° on average (however, it is presumed that the second condition is not satisfied in these nozzles). Therefore, in the dual throat nozzles within the technical scope of the present invention, the second condition (thrust is not deflected when it should not be) is satisfied, and it can be said that in the cases of these flow rate ratios, performance almost equivalent to the deflection performance reported in Non-Patent Document 1 and Non-Patent Document 2 is obtained.

[0086] In the results of the first condition confirmation simulation (part 2), the secondary jet injection port is B The flow rate ratio r s The Mach number distribution and streamlines in the nozzle when the flow rate ratio r is set to 1%, 2%, and 4% are shown in Figures 31, 32, and 33. s It can be seen that as the Mach number increases, the shock wave becomes weaker (the Mach number upstream of the shock wave becomes smaller), and the separation region becomes larger accordingly.

[0087] As is clear from FIG. 30, the flow rate ratio r s In the region where R is 3% or less, the secondary jet injection position is B ,R C ,R D ,R E When comparing the cases where the injection is placed further upstream, the deflection efficiency |δ| / r s However, the position R set most upstream in the diverging section tends to be higher. A Only when injected from B As mentioned in the "Problems to be Solved by the Invention" section, these trends are due to the B ,R C ,R D ,R E In the case of injection from R, this can be explained by the fact that the further upstream the secondary jet injection port 19 is located, the wider the range in which the main flow separates from the nozzle inner wall. A When injected from the mainstream F IN1 The flow tube surrounding the secondary jet F IN2 A place appears where the cross-sectional area is smaller than the first throat 13 due to the pressure of the nozzle, and this place becomes the actual first throat. As a result, the shock wave in the nozzle becomes stronger (i.e., the Mach number upstream of the shock wave becomes larger), and the main flow F IN1 This can be explained by the fact that the separation of the thrust and the deflection of the thrust are suppressed. A and R C This coincides with the flow field conditions shown in the figure (comparison of streamlines and Mach number distribution inside the nozzle when installed in a nozzle with a nozzle hole).

[0088] The absolute value |δ| is the flow rate ratio r2 / (r1+r2)(=r s ) divided by |δ| / r s is the deflection efficiency, the injection position R B ,R C In this case, the flow rate ratio r s Deflection efficiency |δ| / r when is 2% s is 5.25 to 6.35° / %, and the flow rate ratio r s Deflection efficiency |δ| / r when is 3%s In this way, when the configuration of the present invention is adopted, the flow rate ratio r s High deflection efficiency of 4° / % or more in a small range of |δ| / r (4% or less, 3% or less, or 2% or less) s It is possible to achieve even higher deflection efficiencies |δ| / r of 5° / % or more and 6° / % or more. s It is also possible to achieve a deflection efficiency of |δ| / r s Although the upper limit of is unknown, in this simulation it could not exceed 10° / %. For reference, as a very simple comparison example, s When a secondary jet with a deflection efficiency of 2% is ejected from the nozzle exit in a direction perpendicular to the nozzle central axis with the same flow velocity, density, and temperature as the main flow at that point, the absolute value of |δ| is calculated to be only about 1.2°, and the deflection efficiency |δ| / r s This also shows that the injection position R B ,R C The results obtained with such high deflection efficiency |δ| / r are surprising. s By realizing this, the thrust vectoring angle δ required for the operation of various flying bodies and aircraft incorporating a thrust vector control device can be realized with lighter weight and less power, which can be said to lead to improvements in the performance (maneuverability, fuel efficiency, etc.) of the entire system of these flying bodies and aircraft. [Explanation of symbols]

[0089] 10 Dual throat nozzle 10a Nozzle centerline 11 Nozzle inlet 12 First drawing section 13 First Throat 14 Spreading section 15 Second constriction section 16 Second Throat 18 Nozzle outlet 19 Secondary jet inlet A1 * Cross-sectional area of ​​the first throat A2* Cross-sectional area of ​​the second throat D1 Nozzle inlet opening width F a Typical streamlines F b Typical streamlines F c Typical streamlines F d Typical streamlines F IN1 Main flow from the engine combustion chamber into the nozzle inlet F IN2 Secondary jet F OUT Jet ejected from the nozzle outlet α1 peeling area

Claims

1. A fluid type thrust vector control device that changes the direction of a jet flowing from a nozzle outlet by injecting a secondary jet from a secondary jet injection port provided in the middle of the nozzle into a main flow flow that flows from a nozzle inlet to a nozzle outlet, The nozzle is a first throttle portion having an upstream end serving as a nozzle inlet, a downstream end having a cross-sectional area smaller than that of the upstream end, and a downstream end serving as a first throat; a diverging portion connected to a downstream side of the first throttling portion, the diverging portion having a cross-sectional area at a downstream end thereof larger than a cross-sectional area at an upstream end thereof; a second throttle portion connected to the downstream side of the diverging portion, the cross-sectional area of ​​the downstream end of which is smaller than the cross-sectional area of ​​the upstream end of which, and the downstream end of which serves as a second throat; A dual throat nozzle having The cross-sectional area of ​​the first throat ("A １ ＊ ") relative to the cross-sectional area of ​​the second throat ("A ２ ＊ ") Ratio A ２ ＊ / A １ ＊ By setting this to 1.2 to 1.8, during non-deflection control in which only the main flow flows in without injecting a secondary jet into the nozzle, a shock wave is generated in the diverging section, and the flow becomes supersonic upstream of this, preventing the flow from separating from the nozzle inner wall surface, thereby preventing unintended thrust deflection. The Mach number upstream of the shock wave is M １ and the specific heat ratio is γ, Ratio A ２ ＊ / A １ ＊ Substituting the value of into the left side of the following equation 3, and solving the following equation 3 and the following equation 4 simultaneously, the cross-sectional area A Ｓ In contrast, The secondary jet inlet has a cross-sectional area A Ｓ The cross-sectional area is 0.85 times or less, and the cross-sectional area A obtained from the following formula 8 ｍｉｎ Installed in a location with a cross-sectional area of ​​more than A fluid type thrust vector control device characterized by: [Equation 9] [Equation 10] [0011]

2. 2. A fluid thrust vector control device according to claim 1, wherein the nozzles have the same cross-sectional shape parallel to a plane including the centerline of the nozzle.

3. 2. A fluid type thrust vector control device according to claim 1, wherein the nozzle is in the form of a body of revolution with its center line as its axis.

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