Inter-stage pneumatic separation method for sounding rocket

Abstract

The invention relates to a sounding rocket interstage pneumatic separation method, which is characterized in that a sounding rocket is designed into a two-stage javelin-shaped structure, the sounding rocket comprises a load stage and a power stage, the outer diameter D2 of the load stage is smaller than the outer diameter D1 of the power stage, a section difference is formed, the load stage and the power stage are in mortise and tenon joint connection, and after the power stage is finished, the power stage is connected with the load stage. And the pneumatic resistance difference of the D1 and the D2 is utilized to form a speed difference, and mortise and tenon friction force is overcome to achieve automatic separation. By elaborately designing a javelin-shaped section difference structure and adopting a tenon-and-mortise connection mode, two-stage stable and efficient separation of the sounding rocket under the high dynamic pressure condition is achieved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of sounding rocket, and particularly relates to a sounding rocket interstage aerodynamic separation method. BACKGROUND

[0002] In the process of launching a rocket, the interstage separation technology refers to the separation mechanism of the lower stage rocket from the upper stage rocket after the lower stage rocket is shut down. This technology mainly includes two types: hot separation and cold separation. The hot separation technology involves starting the upper stage rocket engine before separation, unlocking the connecting device, and making the upper stage rocket accelerate by means of the engine thrust. At the same time, the exhaust flow of the upper stage rocket acts on the lower stage rocket, and the lower stage rocket is slowed down by the pressure and aerodynamic resistance of the exhaust flow, thereby realizing separation. The cold separation technology is to unlock the connecting device when the thrust of the lower stage rocket decays to near zero and the upper stage rocket engine has not been started, and to realize separation by using a separation impulse device (such as a retro-rocket). These two separation methods have their own advantages and disadvantages. For example, the hot separation technology has a faster separation speed and a shorter separation time, thereby reducing the loss of control time of the attitude control system of the upper stage rocket. However, when the hot separation technology is used, the interstage structure must withstand the influence of high-temperature and high-pressure exhaust flow, and the exhaust emission after the engine is started must be considered, which will increase the disturbance in the separation process. In contrast, the separation impact load of the cold separation technology is smaller, the interstage structure is shorter, but a larger separation impulse is needed, the separation time is longer, and the loss of control time of the attitude control system is increased. Therefore, the most suitable separation method needs to be selected according to the specific application environment.

[0003] In the field of sounding rockets, when the uncontrolled rocket separates from the first and second stages, the separation process is often carried out under high dynamic pressure conditions due to the low flight altitude and high speed, which will cause serious aerodynamic interference, making the interstage separation problem extremely complex. In order to ensure flight stability and achieve the predetermined detection height, while simplifying the structure and reducing the cost, a two-stage separation method based on aerodynamics is proposed. SUMMARY

[0004] The present application aims to overcome the many difficulties faced by the existing interstage separation technology of sounding rockets. In view of the problems of hot separation, such as large disturbance, easy damage to the structure of the rocket and the carried equipment, and cold separation, such as complex structure, leading to increased cost and reduced reliability, a separation method based on the unique principle of aerodynamic resistance difference is proposed. By carefully designing the "spear" type cross-section difference structure and using the mortise and tenon connection method, the stable and efficient separation of the two stages of the sounding rocket under high dynamic pressure conditions is realized. At the same time, it is expected that with this innovative design, the structure of the rocket can be greatly simplified, the manufacturing cost can be effectively reduced, and the performance and economic benefits of the sounding rocket in practical application can be improved.

[0005] The present application is realized by the following technical solutions.

[0006] The sounding rocket of the present application adopts a unique two-stage "spear" type structure design. Among them, the load stage presents the characteristics of a slender shape, and its diameter is set as D2. This stage is only used to install the detection instrument, and the internal power device is not equipped, and by skillfully reducing the diameter, the aerodynamic resistance can be significantly reduced, ensuring that the energy loss is as low as possible during flight, and improving the accuracy and stability of detection. The power stage selects a large-diameter solid rocket engine, and its diameter is set as D1, and D1> D2. After the end of the power stage work, the aerodynamic resistance difference generated by the diameter difference of the two-stage rocket is fully utilized to form a speed difference between the two-stage rocket, so as to realize automatic separation, which greatly simplifies the separation process and improves the reliability of separation. The mortise and tenon connection structure is adopted between the two-stage rockets. Specifically, the head of the first-stage rocket (power stage) is provided with a tenon, and the bottom of the second-stage rocket (load stage) is correspondingly provided with a mortise, and a key groove is carefully designed in the mortise. The tenon is provided with a flat key, and through the close cooperation of the key groove and the flat key, the accurate limiting of the connection between the two-stage rockets is realized, ensuring that the connection is stable during flight, and at the same time, it can be smoothly unlocked at the moment of separation.

[0007] According to the in-depth separation principle research and a large amount of experimental data verification, the diameters of the power stage and the load stage need to meet a specific mathematical relationship:

[0008]

[0009] In the formula, m1 and m2 are the masses of the power stage and the load stage at the end of the power stage, λ is the ratio of the outer diameters of the load stage and the power stage D1 / D2, and β is the ratio of the friction force to the aerodynamic section resistance of the load stage. Its calculation formula is:

[0010]

[0011] In the formula, f represents the interstage friction, ρ is the atmospheric density, v is the rocket flight speed, C d is the resistance coefficient, and S2 is the aerodynamic section area of the load stage. By accurately determining these parameters, the aerodynamic resistance difference between the two stages during flight can meet the separation requirements and realize reliable separation.

[0012] When the work of the power stage ends, the speed of the two-stage rocket remains the same. Since the diameter D1 of the power stage is larger than the diameter D2 of the load stage, according to the aerodynamic resistance formula:

[0013]

[0014] In the formula, F is the aerodynamic resistance, ρ is the atmospheric density, v is the rocket flight speed, C dis the drag coefficient, and S is the windward area. Therefore, it can be known that the aerodynamic drag suffered by the power stage is greater than the aerodynamic drag suffered by the two-stage rocket, which makes the deceleration speed of the power stage rocket faster, so that a speed difference is rapidly formed between the two stages. With the continuous increase of the speed difference, the axial force gradually overcomes the friction between the mortise and tenon connection structure. When the axial force reaches a certain degree, the flat key can be smoothly slid out of the key groove, and thus the two-stage rocket successfully realizes separation and continues to fly according to the predetermined orbit.

[0015] Compared with the prior art, the advantages of the present application are:

[0016] 1. The separation structure is greatly simplified.

[0017] 2. The cost is effectively reduced. BRIEF DESCRIPTION OF DRAWINGS

[0018] Figure 1 is a schematic diagram of the mortise and tenon connection and separation structure of the sounding rocket according to the present application;

[0019] Figure 2 is a schematic diagram of the three-dimensional structure according to the present application;

[0020] Figure 3 is a schematic diagram of the front view structure according to the present application;

[0021] In the figure: 1, load stage, 2, mortise, 3, tenon, 4, key groove, 5, flat key, 6, interstage connection section, 7, power stage, D1 is the outer diameter of the power stage, and D2 is the outer diameter of the load stage. DETAILED DESCRIPTION

[0022] The present application will be further described below in combination with the drawings and specific embodiments, but is not limited by the present application.

[0023] As shown in Figure 1 and Figure 2 , a sounding rocket interstage aerodynamic separation method is provided, in which the sounding rocket is designed as a two-stage “spear” type structure, the sounding rocket comprises a load stage 1 and a power stage 7, the outer diameter D2 of the load stage 1 is smaller than the outer diameter D1 of the power stage 7, forming a cross-sectional difference, the load stage 1 and the power stage 7 are connected by a mortise and tenon connection, the head of the power stage 7 is a conical interstage connection section 6, the mortise 2 at the bottom of the load stage 1 is provided with a key groove 4 inside, and the flat key 5 of the tenon 3 is matched and limited in position;

[0024] After the power stage ends, a speed difference is formed by the aerodynamic drag difference between D1 and D2, the mortise and tenon friction is overcome to realize automatic separation, and the outer diameter D2 of the load stage 1 and the outer diameter D1 of the power stage 7 satisfy:

[0025]

[0026] Wherein m1, m2 are the mass of the power stage 7 and the load stage 1 at the end of the power stage, β is the ratio of the friction force to the aerodynamic cross-sectional resistance of the load stage, and λ is the ratio of the outer diameter of the load stage to the power stage D1 / D2.

[0027] Further, the load stage 1 is a non-powered slender structure, only used for mounting the detection instrument, and the outer diameter D2 is 86 mm.

[0028] Further, the key groove 4 has a depth of 8.5 mm, and the flat key 5 is made of the same material as the tenon 3, i.e., an aluminum alloy.

[0029] Further, β is the ratio of the friction force to the aerodynamic cross-sectional resistance of the load stage, and the calculation formula is as follows:

[0030]

[0031] In the formula, f represents the inter-stage friction, ρ is the atmospheric density, v is the rocket flight speed, C d is the resistance coefficient, and S2 is the aerodynamic cross-sectional area of the load stage. By accurately determining these parameters, the aerodynamic resistance difference between the two stages during the flight of the rocket can be ensured to meet the separation requirements, and reliable separation can be achieved.

[0032] When the work of the power stage ends, the speed of the two-stage rocket remains the same. Since the diameter D1 of the power stage is larger than the diameter D2 of the load stage, according to the aerodynamic resistance formula:

[0033]

[0034] In the formula, F is the aerodynamic resistance, ρ is the atmospheric density, v is the rocket flight speed, C d is the resistance coefficient, and S is the windward area. Therefore, it can be known that the aerodynamic resistance of the power stage is greater than that of the two-stage rocket, which makes the deceleration speed of the power stage rocket faster, so that a speed difference is rapidly formed between the two stages. As the speed difference continuously increases, the axial force gradually overcomes the friction between the mortise and tenon connection structure. When the axial force reaches a certain degree, the flat key can smoothly slide out of the key groove, and thus the two-stage rocket successfully realizes separation and continues to fly according to the predetermined orbit.

[0035] Embodiment

[0036] The load stage is determined according to the size of the detection equipment carried, and the outer diameter D2 of the load stage is 86 mm, the length is 1100 mm, and the estimated mass m2 is 15.9 kg (including the detection equipment).

[0037] The power stage selects a high-performance solid rocket engine, and the outer diameter is designed as D1=255 mm. After the work of the power stage ends, the mass of the power stage m1 is 59.2 kg.

[0038] According to the experiment, the ratio of the interstage friction force and the load stage aerodynamic section resistance β=0.7784, the ratio of the docking section and the load stage section λ≈0.22. The above data are substituted into the automatic separation condition expression:

[0039]

[0040] D1=237.4mm is obtained, because the actual outer diameter of the power stage is 255mm, so the outer diameter difference between the two stages can meet the automatic separation condition.

[0041] The above only describes the preferred embodiments of the present application, and does not limit the implementation and protection scope of the present application. For those skilled in the art, it should be realized that the schemes obtained by equivalent replacement and obvious changes of the present application description and drawings should be included in the protection scope of the present application.

Claims

1. A method of aerodynamic staging of a sounding rocket, characterized in that The sounding rocket is designed as a two-stage "lance" type structure, and the sounding rocket comprises a load stage (1) and a power stage (7), the outer diameter D2 of the load stage (1) is smaller than the outer diameter D1 of the power stage (7), a section difference is formed, the load stage (1) and the power stage (7) are connected through a mortise and tenon joint, the head of the power stage (7) is a tapered interstage connecting section (6), the front end of the interstage connecting section (6) is provided with a tenon (3), the bottom of the load stage (1) is provided with a mortise (2), and a key groove (4) in the mortise (2) is matched with a flat key (5) of the tenon (3) to limit the position; After the power section ends, a speed difference is formed by using the aerodynamic resistance difference between D1 and D2, the automatic separation is realized by overcoming the mortise and tenon friction force, the outer diameter D2 of the load stage (1) and the outer diameter D1 of the power stage (7) satisfy the following formula: Wherein m1 and m2 are the masses of the power stage (7) and the load stage (1) at the end of the power section, β is the friction force to load stage aerodynamic section resistance ratio, and λ is the load stage to power stage outer diameter ratio D1 / D2.

2. A method of aerodynamic staging of a sonde rocket according to claim 1, characterized in that, The load stage (1) is a non-powered slender structure, and the outer diameter D2 is 86mm.

3. The method of aerodynamic staging of a sonde rocket according to claim 1, wherein, The depth of the key groove (4) is 8.5mm.

4. The method of aerodynamic staging of a sonde rocket according to claim 1, wherein, β is the friction force to load stage aerodynamic section resistance ratio, and the calculation formula is as follows: where f represents the interstage drag force, p is the atmospheric density, v is the rocket flight velocity, C d is the drag coefficient, and S2 is the aerodynamic cross-sectional area of the payload stage.

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