Thrust control method and device of combustion chamber, launch vehicle and storage medium

Abstract

The application discloses a thrust control method of a combustion chamber for a small rocket, and the thrust control method comprises the following steps: collecting a total load of propellant; dynamically calculating a thrust adjustment node temperature; dividing a temperature interval; collecting a wall temperature of the combustion chamber in real time; and adjusting the thrust according to the temperature interval. In the early stage of rocket starting, the propellant load is large, the higher thrust adjustment node is automatically matched, the full-thrust working time is prolonged, and the climbing acceleration is improved. In the later stage of flight, the propellant is consumed, and the load is small, the thrust adjustment node temperature is automatically reduced, the thrust is slightly reduced in advance, the wall temperature is prevented from being continuously over-adjusted, the response time of the combustion chamber to be heated to the target temperature is shortened, the invalid consumption of the propellant is reduced, and more effective orbit loading surplus is released. The node temperature is calculated by using a linear fitting function, the calculation amount is small, the requirement of the rocket load controller for the algorithm power is low, the low-cost flight control hardware of the small liquid rocket or solid-liquid mixed rocket can be compatible, and the engineering landing performance is strong.

Description

Technical Field

[0001] This invention relates to the field of thermal control technology for launch vehicle propulsion systems, and in particular to an adaptive control method for rocket engine combustion chamber thrust, supporting devices, launch vehicle, and onboard storage medium. Background Technology

[0002] The liquid main engine of a launch vehicle is the core power component of the rocket. The thermal protection of the high-temperature wall of the combustion chamber directly determines the engine's operating time and overall reliability. Existing technologies usually use cooling films or cooling channels as thermal protection solutions. However, these solutions typically complicate the engine structure, increase the rocket's ineffective load, drive up manufacturing costs, and have long manufacturing cycles.

[0003] Currently, mainstream rocket engines with temperature control protection adopt a fixed threshold thrust reduction strategy to prevent combustion chamber overheating and erosion. Although this fixed threshold temperature control scheme can achieve basic overheat protection, the fixed threshold is a factory-preset fixed constant, which cannot match the dynamic change law of propellant load throughout the rocket flight process: at the initial moment of rocket engine start-up, the total propellant load in the tank reaches its peak value; as the engine continues to work and the propellant is continuously consumed, the longer the working time, the more the remaining total propellant load decreases.

[0004] Therefore, the existing fixed threshold technology has obvious drawbacks: If the fixed threshold trigger temperature is too low, the thrust will lag during the initial stage of engine startup and under conditions of large total propellant load, resulting in a significant decrease in the rocket's climb acceleration efficiency and an inability to meet the rocket's need to increase climb acceleration. If the fixed threshold trigger temperature is too high, in the later stages of flight, when a large amount of propellant is consumed and the total payload is small, the fixed threshold will increase the ineffective consumption of propellant, directly reducing the effective orbital payload that the rocket can carry.

[0005] Ground hot test temperature-time curves demonstrate that, at the same thrust level, a doubling of propellant load increases the time required for the combustion chamber to reach the target temperature by more than 60%. Fixed threshold temperature control strategies cannot adapt to load changes and cannot balance acceleration efficiency with thermal protection safety.

[0006] This invention provides a method, device, launch vehicle, and storage medium for controlling the temperature and thrust of a rocket engine combustion chamber. It addresses the problems of existing fixed threshold temperature control strategies being unable to adapt to changes in propellant load, having low acceleration efficiency, and insufficient thermal protection safety. Based on the real-time total propellant load, it dynamically generates thrust adjustment temperature nodes, adaptively divides the temperature control range, achieves precise thrust control, balances combustion chamber thermal protection safety with rocket climb acceleration efficiency, and reduces propellant loss. Summary of the Invention

[0007] To overcome the shortcomings mentioned above, the present invention aims to provide a technical solution that can solve the above problems.

[0008] A thrust control method for a combustion chamber, used in a small rocket, wherein the small rocket integrates a high-temperature sensor and a propellant mass sensor, the thrust control method comprising: The total propellant load is collected by using the propellant mass sensor to detect the total mass of propellant supplied to the engine from the tank, thereby obtaining the total propellant load. The thrust adjustment node temperature is dynamically calculated, and the matching thrust adjustment node temperature is determined according to the total propellant load. Different total propellant loads correspond to different thrust adjustment node temperatures, and all thrust adjustment node temperatures are lower than the combustion chamber target temperature. Divide the temperature ranges, using the thrust adjustment node temperature and the combustion chamber target temperature corresponding to the current rocket flight program as interval dividing points, and divide at least four combustion chamber temperature ranges; The combustion chamber wall temperature is collected in real time using the high-temperature temperature sensor. The thrust is adjusted in stages according to temperature ranges, and the rocket engine thrust level is adjusted in real time according to the temperature range of the combustion chamber wall. Among them, the closer the real-time temperature range of the combustion chamber is to the target temperature, the lower the engine output thrust level; Preferably, the step of calculating the matching thrust adjustment node temperature based on the total propellant load includes: Based on the pre-calibrated fitting function between the thrust adjustment node temperature and the total propellant load, and combined with the real-time collected total propellant load, the corresponding thrust adjustment node temperature is calculated. The fitting function is generated by fitting the relationship curves between different propellant loads and the optimal thrust adjustment node temperature obtained from multiple sets of ground hot test experiments. Preferably, the fitting function is a linear fitting function with the formula: T(G) = A·G + B; Wherein, G is the total propellant load, and A and B are the coefficients of the first-order term and the constant term obtained by linear fitting the experimental relationship curve, respectively; Preferably, the dynamic calculation of the thrust adjustment node temperature and the division of temperature intervals, the determination of the matching thrust adjustment node temperature based on the total propellant load, and the division of at least four combustion chamber temperature intervals using the thrust adjustment node temperature and the combustion chamber target temperature corresponding to the current rocket flight program as interval dividing points, including: Based on the first fitting function T1(G), the second fitting function T2(G), and the total propellant load G, the first thrust adjustment node temperature (T1) and the second thrust adjustment node temperature (T2) are calculated respectively. The first fitting function T1(G) and the second fitting function T2(G) were obtained by fitting the temperature curves of the two-stage optimal thrust reduction nodes corresponding to different propellant loads in multiple sets of ground hot test tests. Using the first thrust adjustment node temperature (T1), the second thrust adjustment node temperature (T2), and the combustion chamber target temperature (T0) as interval dividing points, four combustion chamber temperature intervals are defined, where (T1) <T2<T0); The four combustion chamber temperature ranges are, in order, the first temperature range (-∞, T1), the second temperature range (T1, T2), the third temperature range (T2, T0), and the fourth temperature range (T0, +∞). The real-time adjustment of the rocket engine thrust level based on the real-time temperature range of the combustion chamber wall includes: gradually reducing the rocket engine thrust level as the real-time temperature of the combustion chamber gradually rises from the first temperature range to the fourth temperature range; Preferably, when the real-time temperature of the combustion chamber rises from the first temperature range to the fourth temperature range and the thrust level is gradually reduced, a fixed thrust difference is used for step-down thrust reduction. Preferably, when the real-time temperature of the combustion chamber rises from the first temperature range to the fourth temperature range and the thrust level is reduced step by step, a gradual thrust reduction method is adopted to gradually reduce the thrust difference. Preferably, during the rocket takeoff phase, when the real-time temperature of the combustion chamber is in the first temperature range (-∞, T1), the rocket engine is controlled to output the rated maximum thrust level; when the real-time temperature of the combustion chamber is in the fourth temperature range (T0, +∞), the rocket engine is controlled to switch to the minimum sustaining thrust level to suppress overheating and ablation of the combustion chamber wall.

[0009] A thrust control device for a combustion chamber, comprising a rocket engine integrating a high-temperature temperature sensor and a propellant mass sensor; the control device is used to execute the thrust control method for the combustion chamber described above, the device comprising: The load acquisition module is used to receive the signal collected by the propellant mass sensor, and output the total propellant load after filtering and compensation. The temperature control zone division module has a built-in fitting function for load and thrust adjustment node temperature. It is used to calculate the thrust adjustment node temperature based on the total propellant load and divide at least four combustion chamber temperature zones in combination with the target combustion chamber temperature. The temperature acquisition module is used to receive signals collected by the high-temperature temperature sensor, process them, and output the real-time temperature of the combustion chamber wall. The thrust adjustment module is used to output thrust control commands based on the real-time temperature range of the combustion chamber to adjust the thrust level of the rocket engine; The closer the real-time temperature range of the combustion chamber is to the target temperature of the combustion chamber, the lower the output thrust level.

[0010] A launch vehicle includes: one or more onboard processors, an onboard storage device, a propellant mass sensor, and a combustion chamber high-temperature sensor; The propellant mass sensor is used to detect the total propellant mass supplied to the engine from the tank and output the total propellant load. The combustion chamber high-temperature sensor is used to collect the real-time temperature of the combustion chamber wall. The onboard storage device stores one or more control programs; When one or more control programs are executed by the onboard processor, the onboard processor implements the thrust control method for the combustion chamber described above.

[0011] A storage medium storing an onboard control program, which, when executed by an onboard processor, implements the thrust control method for the combustion chamber described in any of the preceding claims.

[0012] Compared with the prior art, the advantages of the present invention are: This invention collects the total propellant load in real time and dynamically matches the thrust adjustment node temperature as the load continuously decreases during engine operation. In the early stages of rocket startup, when the propellant load is large, a higher thrust adjustment node is automatically matched to extend the full thrust operating time and improve climb acceleration. In the later stages of flight, when propellant is consumed and the load decreases, the thrust adjustment node temperature is automatically reduced to preemptively reduce thrust slightly, avoid continuous wall temperature overshoot, shorten the response time for the combustion chamber to heat up to the target temperature, reduce ineffective propellant consumption, and release more effective orbital payload margin.

[0013] The node temperature calculation of this invention uses a linear fitting function, which requires little computation and has low computational power requirements for the rocket controller. It is compatible with low-cost flight control hardware for small liquid and solid-liquid hybrid rockets and has strong engineering feasibility.

[0014] This invention employs either a step-down thrust reduction with a fixed thrust difference or a progressive thrust reduction with gradually decreasing thrust difference to gradually reduce the thrust level: When using a fixed thrust difference step thrust reduction method, the entire control logic relies on only one set of difference constants to complete the calculation. Only basic subtraction operations are needed to output the target thrust. The calculation logic is extremely simple and occupies very little flight control computing power and storage resources. Ground hot-fire testing only requires calibration of a single thrust difference parameter. The workload of whole-machine debugging and model iteration calibration is low. The thrust solution in a single control cycle consumes only a very small number of CPU clock cycles. It will not cause computational interference to other flight control tasks such as multi-channel thermocouple averaging filtering, propellant load acquisition, attitude calculation, etc. Its operating speed is extremely fast. The flight control system can quickly identify thrust output anomalies. The fault self-checking logic is simple and reliable. It is suitable for rapid thrust control and thermal protection of small rockets. When using a progressive thrust reduction method that gradually decreases the thrust difference, the temperature is rapidly controlled when the combustion chamber wall temperature reaches the safe target temperature of the combustion chamber, and the thrust is slowly reduced in the low load critical range during the later stages of flight without losing power.

[0015] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0016] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0017] Figure 1 This is a schematic diagram illustrating the working principle of the thrust control method of the present invention. Detailed Implementation

[0018] The technical solutions in the embodiments of the present invention will be clearly and completely described below. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0019] In the description of this invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0020] Furthermore, in the description of this invention, unless otherwise explicitly specified and limited, the terms "installation," "connection," "linking," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal connection of two components; they can refer to a wireless connection or a wired connection. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0021] Furthermore, the technical features involved in the different embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0022] Please see Figure 1 In this embodiment of the invention, a thrust control method for a combustion chamber temperature range is applied to a small solid-liquid launch vehicle with a takeoff weight of 5 tons. The onboard engine integrates a high-precision propellant mass sensor and a multi-channel high-temperature thermocouple (high-temperature sensor) for the combustion chamber. The onboard embedded processor executes the thrust control method. Ground-based pre-combustion hot-fire tests with multiple propellant loads are performed to calibrate the linear fitting coefficients between the loads and the two-stage thrust nodes. For example, if the preset safe target temperature for the combustion chamber is T0 = 850℃, the thrust control method includes: S11. Collect the total propellant load. A propellant mass sensor is integrated into the rocket engine's propellant input pipeline to collect the total mass of oxidizer and fuel output from the tank in real time. After filtering and temperature compensation, the total propellant load G is obtained, and the load data is uploaded to the onboard processor in real time. Example operating condition: The initial total propellant load G at liftoff is approximately 200 kg. However, as the rocket continues to operate, the oxidizer is continuously consumed, and after 300 seconds of flight, the remaining total propellant load G is approximately 100 kg.

[0023] S12. Dynamically calculate the thrust adjustment node temperature. Different total propellant loads correspond to different thrust adjustment node temperatures, and all thrust adjustment node temperatures are lower than the target temperature within the combustion chamber temperature range. The processor retrieves the pre-stored load-node temperature fitting function. The calculation formulas for the two-stage thrust adjustment nodes are: T1(G)=A1·G+B1, T2(G)=A2·G+B2.

[0024] Substitute the real-time load G to calculate the first-stage cooling node T1 and the second-stage cooling node T2 respectively.

[0025] Operating condition 1: Initial takeoff G=200kg; T1 = 0.12 × 200 + 150 = 390℃; T2 = 0.08 × 200 + 320 = 480℃; Temperature calibration: 390℃ < 480℃ < 850℃.

[0026] Condition 2: After 300 seconds of flight, G = 100 kg; T1 = 0.12 × 100 + 150 = 270℃; T2 = 0.08 × 100 + 320 = 328℃; Temperature calibration: 270℃ < 328℃ < 850℃.

[0027] S13. Divide the temperature range, using the thrust adjustment node temperature and the target combustion chamber temperature corresponding to the current flight program of the rocket as the interval division points, and divide at least four combustion chamber temperature ranges. The ground test run calibration ensures that T1 < T2 < T0, where T0 is the preset combustion chamber safety target temperature for the current flight program. Using T1, T2, and T0 as the division points, four temperature ranges are generated: the first temperature range (-∞, T1), the second temperature range (T1, T2), the third temperature range (T2, T0), and the fourth temperature range (T0, +∞).

[0028] Example, the ground test run calibration coefficients are: A1 = 0.12 °C / kg, B1 = 150 °C / kg, A2 = 0.08 °C / kg, B2 = 320 °C / kg; At the initial takeoff, the total propellant payload G = 200 kg: Using T1 = 390 °C, T2 = 480 °C, and T0 = 850 °C as the division points, four temperature ranges are divided: the first temperature range (-∞, 390 °C), the second temperature range (390 °C, 480 °C), the third temperature range (480 °C, 850 °C), and the fourth temperature range (850 °C, +∞).

[0029] After flying for 300 s, the remaining total propellant payload G = 220 kg: Update the temperature range to the first temperature range (-∞, 270 °C), the second temperature range (270 °C, 328 °C), the third temperature range (328 °C, 850 °C), and the fourth temperature range (850 °C, +∞). The demarcation temperature decreases synchronously with the decrease in the payload, achieving interval adaptive adjustment, and the thrust adjustment node temperature is always lower than the combustion chamber safety target temperature.

[0030] S14. Collect the combustion chamber wall temperature in real time. A number of high-temperature thermocouples are evenly arranged circumferentially on the outer wall of the combustion chamber. The average value of the multi-channel temperature measurement data is used as the real-time wall temperature T of the combustion chamber to eliminate the single-point temperature measurement error.

[0031] S15. Adjust the thrust in grades according to the temperature range. According to the temperature range in which the real-time temperature of the combustion chamber wall in the temperature range is located, the thrust gear of the rocket engine is adjusted in real time; when the wall temperature gradually rises from the low-temperature range to the high-temperature range, the engine thrust gear is gradually lowered. The closer the temperature range is to T0, the smaller the output thrust.

[0032] Calculate the thrust adjustment node temperature matched according to the total propellant payload, including: according to the fitting function of the pre-calibrated thrust adjustment node temperature and the total propellant payload, combined with the real-time collected total propellant payload, calculate the corresponding thrust adjustment node temperature; the fitting function is generated by fitting the relationship curves of different propellant loads and the optimal thrust adjustment node temperature obtained from multiple groups of ground hot test runs.

[0033] The fitting function is a linear fitting function, and its formula is: T(G)=A·G+B; where G is the total propellant load, and A and B are the coefficients of the first term and the constant term obtained by linear fitting the experimental relationship curve, respectively.

[0034] In this embodiment of the invention, the matching thrust adjustment node temperature is determined based on the total propellant load, and at least four combustion chamber temperature intervals are divided using the thrust adjustment node temperature and the target temperature of the combustion chamber temperature interval corresponding to the current flight program of the rocket as interval dividing points. This includes: calculating the first thrust adjustment node temperature (T1) and the second thrust adjustment node temperature (T2) based on the first fitting function T1(G), the second fitting function T2(G), and the total propellant load G, respectively; the first fitting function T1(G) and the second fitting function T2(G) are respectively obtained by fitting the two-stage optimal thrust reduction node temperature curves corresponding to different propellant loads in multiple sets of ground hot test tests.

[0035] Existing technologies often employ quadratic polynomials, piecewise interpolation, and nonlinear curve fitting to establish the relationship between load and nodal temperature. These calculations involve complex floating-point operations such as squaring, multi-order accumulation, piecewise judgment, and interpolation lookup tables. Quadratic polynomials involve squaring and multiple accumulation operations, while piecewise interpolation requires multiple interval condition judgments and multi-point weighted interpolation calculations. The numerous complex floating-point operations necessitate that the onboard controller be equipped with a hardware floating-point unit, a high-performance DSP, or a multi-core main control chip, significantly increasing the cost of onboard flight control hardware and conflicting with the low-cost design requirements of small rockets. Furthermore, the interpolation lookup table scheme requires real-time reading and caching of multiple sets of discrete load-temperature data points. During the iteration process of the nonlinear fitting algorithm, intermediate calculation variables and error correction parameters also need to be temporarily stored, consuming a large amount of on-chip RAM and processor computing and memory resources, leading to… The flight control system suffers from resource constraints; piecewise interpolation and high-order nonlinear fitting rely on dedicated numerical calculation function libraries, making low-level driver adaptation complex and difficult to quickly port to different models of low-cost open-source flight control boards; thrust adjustment suffers from significant computational delays; combustion chamber thermal safety control lacks real-time performance; algorithm porting and model iteration calibration require a huge amount of work, resulting in high barriers to engineering implementation; complex multi-step floating-point multiplication and division and weighted interpolation operations suffer from multi-order error accumulation; when high and low temperatures change on the rocket and flight vibrations bring small sensor noises, high-order algorithms will amplify the calculation deviations; calculation errors are easily amplified by environmental disturbances; thrust control has poor robustness and is prone to malfunctions.

[0036] Therefore, the existing calculation methods result in the controller having no remaining computing resources to carry redundant protection logic such as multi-channel temperature measurement fault self-check, propellant balance warning, and thrust mechanism failure diagnosis. It can only reduce the on-board safety monitoring module, weaken the fault protection capability throughout the flight cycle, and further reduce the flight safety redundancy of small rockets. The univariate linear fitting scheme adopted in this invention completes the multiplication and addition operation with only two sets of fixed coefficients. It does not require large-capacity storage, high-end computing chips and dedicated algorithm libraries. It has fewer calculation steps, a single source of error, and is easy to port. It comprehensively overcomes the above-mentioned defects of the existing technology from multiple dimensions such as hardware cost, real-time performance, calibration difficulty and control stability. Example 1

[0037] When the real-time temperature of the combustion chamber rises from the first temperature range to the fourth temperature range, and the thrust is gradually reduced, a fixed thrust difference is used for step-by-step thrust reduction. Numerical example: the thrust of the first range is 10kN; the thrust of the second range is 8kN; the thrust of the third range is 6kN; and the thrust of the fourth range is 4kN. The difference between adjacent ranges is a fixed thrust difference of 2kN. This embodiment does not require real-time calculation of the variable thrust reduction based on load, temperature, and environmental parameters. It only relies on preset constants for basic subtraction calculations, combined with linear fitting to determine the node temperature and simple numerical comparison to determine the range. The entire temperature-controlled thrust control logic has extremely low computational load and can be stably operated by a common low-cost microcontroller without the need for a high-end floating-point processor. Furthermore, the thrust reduction range between adjacent ranges is completely consistent, avoiding alternating small and sudden thrust reductions. The propellant supply flow rate changes smoothly, effectively suppressing combustion oscillations, violent fluctuations in combustion chamber pressure, and instantaneous flameout, thus meeting the stable combustion requirements of small solid-liquid rockets. Example 2

[0038] As the real-time temperature of the combustion chamber rises from the first temperature range to the fourth temperature range, and the thrust is gradually reduced at each level, a gradual thrust reduction is adopted, with the thrust difference decreasing step by step. Numerical example: 10kN thrust in the first range; 7kN thrust in the second range; 5kN thrust in the third range; 4kN thrust in the fourth range. Its thrust curve is a gently sloping downward curve, with a steep initial drop followed by a gradual flattening and decreasing slope. Therefore, it significantly reduces thrust immediately upon triggering the first-level warning (T1) to quickly suppress rapid temperature rise. Once the temperature approaches the safe limit (T0), the thrust reduction slows down, preserving basic power. This achieves rapid temperature control when the combustion chamber wall temperature reaches the safe target temperature, and a gradual thrust reduction without loss of power in the low-load critical range during the later stages of flight, balancing ballistic stability and thermal protection. Example 3

[0039] During rocket launch, when the real-time temperature of the combustion chamber is in the first temperature range (-∞, T1), the combustion chamber wall temperature is determined to be in a safe low-temperature state, with no risk of high-temperature ablation or thermal fatigue. The rocket engine is controlled to continuously output the rated maximum thrust, fully releasing the engine's power performance to meet the thrust requirements for rocket launch acceleration, trajectory climb, and orbital maneuvering, prioritizing flight payload and acceleration indicators. When the real-time temperature of the combustion chamber is in the fourth temperature range (T0, +∞), there is a risk of liner ablation and high-temperature shell failure. The rocket engine is immediately switched to the minimum sustaining thrust level. The minimum sustaining thrust only supplies the basic propellant flow to maintain engine stability and prevent flameout, significantly reducing the heat release power of combustion. This curbs the continued rise in wall temperature from the heat source end, quickly suppresses overheating ablation of the combustion chamber wall, and forms the last stage of thermal safety protection for the entire system.

[0040] This embodiment achieves full power output in the cryogenic range during rocket takeoff, maximizing power utilization; and in the overheated range, it forces heat limitation under the premise of minimum thrust maintenance, providing a safety net for the combustion chamber structure.

[0041] A thrust control device for a combustion chamber temperature range, comprising a rocket engine integrating a high-temperature sensor and a propellant mass sensor; the control device is used to execute a thrust control method for the combustion chamber temperature range described above, and the device includes: The load acquisition module is used to receive the signal collected by the propellant mass sensor, and output the total propellant load after filtering and compensation. The temperature control zone division module has a built-in fitting function for load and thrust adjustment node temperature. It is used to calculate the thrust adjustment node temperature based on the total propellant load and divide at least four combustion chamber temperature zones in combination with the target temperature of the combustion chamber temperature zone. The temperature acquisition module is used to receive signals collected by the high-temperature temperature sensor, process them, and output the real-time temperature of the combustion chamber wall within the temperature range. The thrust adjustment module is used to output thrust control commands based on the real-time temperature range of the combustion chamber and adjust the thrust level of the rocket engine. The closer the real-time temperature range of the combustion chamber is to the target temperature range, the lower the output thrust level.

[0042] A launch vehicle includes: one or more onboard processors, onboard storage devices, a propellant mass sensor, and a high-temperature sensor for the combustion chamber temperature range; The propellant mass sensor is installed in the engine propellant input pipeline to detect the total mass of fuel and oxidizer delivered from the tank to the engine in real time and output the original propellant load signal; The high-temperature sensor in the combustion chamber is a multi-channel high-temperature thermocouple arranged circumferentially to collect the wall temperature at multiple points on the outer wall of the combustion chamber in real time and output multiple temperature acquisition signals. The onboard storage device is an on-chip Flash memory chip that stores one or more onboard control programs. When the onboard processor retrieves and executes the control program in the onboard storage device, it drives the hardware sensors to collect data and fully implements the thrust control method for the combustion chamber temperature range described in any of the previous embodiments. It automatically completes the entire process logic, including propellant load filtering compensation, two-stage node temperature linear calculation, four-segment temperature range adaptive division, wall temperature average processing, graded step thrust reduction control, and over-temperature fallback protection.

[0043] The rocket relies on its existing low-cost onboard processor and storage devices to run the control program, without the need to upgrade to high-performance DSPs and large-capacity storage chips, which fits the low-cost R&D and mass production positioning of small rockets. It significantly suppresses the temperature in advance during the full load stage of takeoff, and smoothly reduces thrust in the low load critical range during the middle and late stages of flight to retain power, suppressing the overheating and erosion of the combustion chamber throughout the entire process, extending the service life of the combustion chamber, and improving the reliability of rocket flight.

[0044] A storage medium stores an onboard control program, which, when executed by an onboard processor, implements a thrust control method for the combustion chamber temperature range described above. The program algorithm for this thrust control method occupies minimal storage space, achieving lightweight design, and can utilize low-cost, small-capacity Flash or EEPROM storage media.

[0045] It will be apparent to those skilled in the art that the present invention is not limited to the details of the exemplary embodiments described above, and that the invention can be implemented in other specific forms without departing from the spirit or essential characteristics of the invention. Therefore, the embodiments should be considered in all respects as exemplary and non-limiting, and the scope of the invention is defined by the appended claims rather than the foregoing description. Thus, it is intended that all variations falling within the meaning and scope of equivalents of the claims be included within the present invention.

Claims

1. A thrust control method for a combustion chamber, used in a small rocket, characterized in that, The small rocket integrates a high-temperature sensor and a propellant mass sensor, and the thrust control method includes: The total propellant load is collected by using the propellant mass sensor to detect the total mass of propellant supplied to the engine from the tank, thereby obtaining the total propellant load. The thrust adjustment node temperature is dynamically calculated, and the matching thrust adjustment node temperature is determined according to the total propellant load. Different total propellant loads correspond to different thrust adjustment node temperatures, and all thrust adjustment node temperatures are lower than the combustion chamber target temperature. Divide the temperature ranges, using the thrust adjustment node temperature and the combustion chamber target temperature corresponding to the current rocket flight program as interval dividing points, and divide at least four combustion chamber temperature ranges; The combustion chamber wall temperature is collected in real time using the high-temperature temperature sensor. The thrust is adjusted in stages according to temperature ranges, and the rocket engine thrust level is adjusted in real time according to the temperature range of the combustion chamber wall. Among them, the closer the real-time temperature range of the combustion chamber is to the target temperature, the lower the engine output thrust level.

2. The thrust control method for the combustion chamber according to claim 1, characterized in that, The calculation of the matching thrust adjustment node temperature based on the total propellant load includes: Based on the pre-calibrated fitting function between the thrust adjustment node temperature and the total propellant load, and combined with the real-time collected total propellant load, the corresponding thrust adjustment node temperature is calculated. The fitting function is generated by fitting the relationship curves between different propellant loads and the optimal thrust adjustment node temperature obtained from multiple sets of ground hot test experiments.

3. The thrust control method for the combustion chamber according to claim 2, characterized in that, The fitting function is a linear fitting function of the first order, and its formula is: T(G)=A·G+B; Where G is the total propellant load, and A and B are the coefficients of the first-order term and the constant term obtained by linear fitting the experimental relationship curve, respectively.

4. The thrust control method for the combustion chamber according to claim 2, characterized in that, The dynamic calculation of thrust adjustment node temperature and the division of temperature intervals, the determination of matching thrust adjustment node temperature based on the total propellant load, and the division of at least four combustion chamber temperature intervals using the thrust adjustment node temperature and the combustion chamber target temperature corresponding to the current rocket flight program as interval dividing points, including: Based on the first fitting function T1(G), the second fitting function T2(G), and the total propellant load G, the first thrust adjustment node temperature (T1) and the second thrust adjustment node temperature (T2) are calculated respectively. The first fitting function T1(G) and the second fitting function T2(G) were obtained by fitting the temperature curves of the two-stage optimal thrust reduction nodes corresponding to different propellant loads in multiple sets of ground hot test tests. Using the first thrust adjustment node temperature (T1), the second thrust adjustment node temperature (T2), and the combustion chamber target temperature (T0) as interval dividing points, four combustion chamber temperature intervals are defined, where (T1) <T2<T0); The four combustion chamber temperature ranges are, in order, the first temperature range (-∞, T1), the second temperature range (T1, T2), the third temperature range (T2, T0), and the fourth temperature range (T0, +∞). The method of adjusting the rocket engine thrust level in real time according to the temperature range of the combustion chamber wall includes: gradually reducing the rocket engine thrust level as the real-time temperature of the combustion chamber gradually rises from the first temperature range to the fourth temperature range.

5. The thrust control method for the combustion chamber according to claim 4, characterized in that, When the real-time temperature of the combustion chamber rises from the first temperature range to the fourth temperature range and the thrust level is gradually reduced, a fixed thrust difference is used for step-by-step thrust reduction.

6. The thrust control method for a combustion chamber according to claim 4, characterized in that, When the real-time temperature of the combustion chamber rises from the first temperature range to the fourth temperature range and the thrust level is reduced step by step, a gradual thrust reduction method is adopted to gradually reduce the thrust difference.

7. The thrust control method for the combustion chamber according to claim 4, characterized in that, During the rocket's takeoff phase, when the real-time temperature of the combustion chamber is in the first temperature range (-∞, T1), the rocket engine is controlled to output the rated maximum thrust. When the real-time temperature of the combustion chamber is in the fourth temperature range (T0, +∞), the rocket engine is controlled to switch to the minimum sustaining thrust to suppress overheating and ablation of the combustion chamber wall.

8. A thrust control device for a combustion chamber, characterized in that, The rocket engine integrates a high-temperature sensor and a propellant mass sensor; the control device is used to execute the thrust control method of the combustion chamber according to any one of claims 1-6, and the device includes: The load acquisition module is used to receive the signal collected by the propellant mass sensor, and output the total propellant load after filtering and compensation. The temperature control zone division module has a built-in fitting function for load and thrust adjustment node temperature. It is used to calculate the thrust adjustment node temperature based on the total propellant load and divide at least four combustion chamber temperature zones in combination with the target combustion chamber temperature. The temperature acquisition module is used to receive signals collected by the high-temperature temperature sensor, process them, and output the real-time temperature of the combustion chamber wall. The thrust adjustment module is used to output thrust control commands based on the real-time temperature range of the combustion chamber to adjust the thrust level of the rocket engine; The closer the real-time temperature range of the combustion chamber is to the target temperature of the combustion chamber, the lower the output thrust level.

9. A launch vehicle, characterized in that, include: One or more onboard processors, onboard storage devices, propellant mass sensors, and combustion chamber high-temperature sensors; The propellant mass sensor is used to detect the total propellant mass supplied to the engine from the tank and output the total propellant load. The combustion chamber high-temperature sensor is used to collect the real-time temperature of the combustion chamber wall. The onboard storage device stores one or more control programs; When one or more control programs are executed by the onboard processor, the onboard processor implements the thrust control method for the combustion chamber as described in any one of claims 1-6.

10. A storage medium storing an onboard control program, characterized in that, When the onboard control program is executed by the onboard processor, it implements the thrust control method for the combustion chamber as described in any one of claims 1-6.

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