Rotor moment-of-inertia design method based on autorotation landing capability

Abstract

A rotor moment-of-inertia design method based on an autorotation landing capability. In the method, a characteristic quantity, i.e., "an autorotation characteristic time", for describing an autorotation-landing safety characteristic of a multi-engine helicopter is established, and a design objective of "the autorotation characteristic time" is provided, such that the relationship between the autorotation-landing safety characteristic of the multi-engine helicopter and a rotor moment of inertia is determined, and a theoretical formula and design method for a design objective of the rotor moment of inertia are provided. On the basis of an objective evidence-collection condition for a helicopter, the method can complete an objective design for a rotor moment of inertia in the overall design stage of the helicopter, and determine whether an autorotation-landing safety characteristic of the current overall technical scheme can meet the airworthiness requirements under the objective evidence-collection condition, thereby balancing the gain of the autorotation-landing safety characteristic and the weight of a rotor system.

Description

A rotor moment of inertia design method based on autorotation landing capability Technical Field

[0001] The invention belongs to the field of overall helicopter design, and in particular relates to a method for designing rotor rotational inertia based on autorotation landing capability. Background Art

[0002] Autorotation landing capability is a key safety feature that distinguishes helicopters from other aircraft. It means that even after all engines fail, a helicopter can still generate energy through the "windmilling" effect during descent, allowing the rotors to maintain a certain rotational speed, storing energy. Ultimately, through energy conversion, the rotor's rotational kinetic energy is released, reducing the rate of descent and forward flight speed, allowing for a safe landing.

[0003] Autorotation landing capability is an important guarantee for the safety of helicopters and crew members after engine failure, and is one of the key contents of civil helicopter design and verification.

[0004] Airworthiness regulations set clear requirements for the autorotation landing safety of civil helicopters, specifying flight performance and quality during the transition phase after engine failure, the stable autorotation phase, and the final landing phase. Among these three phases, the pilot's control of the helicopter to complete the transition to autorotation and the transition from autorotation to instantaneous range gain before touchdown are critical periods of energy conversion and a concentrated reflection of the helicopter's ability to safely autorotate and land. This is related to both the pilot's proficiency and the helicopter's physical properties. Some helicopters require less skilled pilots, while others can be dangerous even for experienced pilots. Therefore, analysis and design of helicopter autorotation landing safety are necessary to ensure that the helicopter meets the airworthiness autorotation landing safety requirements even under normal pilot control.

[0005] The transition into autorotation and the instantaneous distance gain before autorotation touchdown must both consider the kinetic energy stored in the rotor and its rate of energy dissipation. The greater the rotor's stored rotational energy and the slower its decay, the better the helicopter's autorotation characteristics; however, the rotor's stored rotational energy is primarily determined by its moment of inertia.

[0006] Therefore, how to determine the rotor moment of inertia design target based on the target airworthiness certification altitude, and then strike a balance between the empty weight cost and the autorotation landing capability benefit to obtain a reasonable and feasible technical solution, has become the key to the autorotation landing flight safety design.

[0007] In the existing technology, the design evaluation methods for autorotation landing safety and whether the rotor moment of inertia meets the requirements are mainly represented by evaluation methods represented by simulation and simulators. Although this evaluation method has a high confidence level, it is costly and time-consuming, and cannot meet the needs of quickly evaluating the rationality of the rotor moment of inertia design during the overall helicopter design stage. During the overall helicopter design stage, there is currently a lack of a complete rapid evaluation method for the autorotation landing safety of multi-engine helicopters. In particular, for multi-engine helicopters that need to meet the airworthiness requirements of high-altitude environments, the existing technology cannot quickly determine whether the autorotation landing safety characteristics of the current overall technical solution can meet the airworthiness requirements under the target certification conditions. As a result, it is impossible to determine the rotor moment of inertia design target and to complete the trade-off between the benefits of the autorotation landing safety characteristics and the weight of the rotor system.

[0008] There are three main difficulties in establishing a rotor inertia design method based on autorotation landing capability:

[0009] First, it is necessary to establish characteristic quantities that describe the autorotation and landing safety characteristics of multi-engine helicopters, determine the relationship between the autorotation and landing safety characteristics of multi-engine helicopters and the rotor moment of inertia, and give the design target of the characteristic quantities, thereby obtaining the target design of the rotor moment of inertia;

[0010] Second, it is necessary to consider the target certification conditions, including the target certification altitude, temperature, and landing weight, and the impact they bring. The rotor moment of inertia target design can be carried out according to the plateau performance requirements.

[0011] Third, it is necessary to link the requirements of airworthiness regulations and pilot human factors to consider the impact of airworthiness regulations on the target design of rotor moment of inertia.

[0012] Therefore, there is an urgent need to establish a rotor moment of inertia design method based on autorotation landing capability.

[0013] Summary of the Invention

[0014] The purpose of the present invention is to address the shortcomings of the existing technology and invent a rotor rotational inertia design method based on autorotation landing capability, so as to introduce target certification conditions in the overall design stage for multi-engine helicopters that need to meet the airworthiness requirements of plateau environments, determine the autorotation characteristic time design target through test flights, introduce the target certification condition characteristic state required power, and the calculation of the maximum pull coefficient per unit solidity of the rotor, thereby taking into account the target certification conditions, especially the plateau environment, and the airworthiness requirements of the autorotation landing safety characteristics, completing the trade-off between the autorotation landing safety characteristics benefits and the weight of the rotor system, completing the rotor rotational inertia design, and ensuring that the overall technical solution of the helicopter can meet the airworthiness autorotation landing safety characteristics requirements of the target certification altitude.

[0015] To solve this technical problem, the technical solution of the present invention is:

[0016] A method for designing rotor moment of inertia based on autorotation landing capability is provided, comprising the following steps:

[0017] Step 1: Determine the design goal of the rotation characteristic time:

[0018] The autorotation characteristic time is defined as the maximum time after an engine failure during which the energy stored in the rotor rotational kinetic energy can sustain the helicopter's flight and maintain altitude without decreasing. The larger of the pilot's reaction time to identifying the last engine failure and the pilot's operational safety lag time is used as the design target for the autorotation characteristic time. The pilot's reaction time to identifying the last engine failure is the average reaction time for the pilot to identify an engine anomaly after the test flight confirms the engine failure.

[0019] Step 2: Calculate the required power Pw of the characteristic state under the target design state: Establish a flight performance calculation model, input the target design state, and calculate the required power Pw of the characteristic state under the target design state. This refers to the power required for the helicopter to maintain flight after engine failure and represents the decay rate of the rotor's stored rotational kinetic energy:

[0020] The target design state refers to the target certification altitude, atmospheric temperature, and maximum landing weight Gw of the helicopter to be designed under the conditions of altitude and temperature;

[0021] Step 3. Calculate the effect of rotor stall on the rotational kinetic energy released per unit moment of inertia of the rotor under the target design state:

[0022] The formula for calculating the rotational kinetic energy E that can be released per unit moment of inertia of the rotor is as follows:

[0023] C T The calculation formula of / σ is as follows:

[0024] Among them C T is the drag coefficient, σ is the rotor solidity, g is the acceleration of gravity, ρ is the air density, Ω is the rotor rotation angular velocity, R is the rotor radius, Gw is the takeoff weight, C y7max The maximum lift coefficient of the main rotor blade airfoil at 70% radius;

[0025] Step 4: Determine C y7max Correction function f(H) that varies with height:

[0026] f(H) is a correction function that takes into account the effect of altitude on the rotor stall characteristics. It is obtained through aerodynamic calculations or wind tunnel tests to consider the influence of different altitude conditions on the aerodynamic characteristics of the blade, where H is the altitude;

[0027] Step 5: Calculate the design target of the rotor moment of inertia I based on the autorotation landing capability. The rotor design should ensure that its rotor moment of inertia is greater than the design target:

[0028] The rotation characteristic time in the formula is the design target of the rotation characteristic time;

[0029] Step 6: Calculate the blade weight cost and weigh the overall helicopter technical solution:

[0030] Based on the position x of the blade's center of gravity in the blade's span and the rotor's moment of inertia I, the weight M of each blade in a rotor system with n blades is obtained. 桨叶 (unit: kg), and then the influence of the rotor moment of inertia I design target on the blade weight is obtained:

[0031] x is the distance from the center of the hub to the center of gravity of the blade, in meters; k is the proportion of the blade's moment of inertia in the rotor system's moment of inertia.

[0032] Furthermore, step six of the design method of the present invention also includes a step of further optimizing the overall technical solution: when it is necessary to further reduce the blade weight or further expand the target design state, return to step two to adjust the target design state and evaluate again until the overall technical solution of the helicopter reaches the target.

[0033] If the blade weight cost obtained in step 6 is unacceptable and the blade weight must be reduced, or if the blade weight cost is considered acceptable and it is desired to further expand the target design state, return to step 2, adjust the target design state, and evaluate again until the overall technical solution of the helicopter meets the target.

[0034] In step six of this method, the k value is used to correct the weight of the blade, and the k value is 0.95-0.97. Since most of the rotor's moment of inertia is concentrated on the blade, the k value is obtained from engineering practice experience and is generally 95-97%.

[0035] The pilot's reaction time to recognizing the last engine failure in step 1 is determined by conducting flight test simulations on helicopters using the same or similar engines as the helicopter to be designed. Similar engines are engines whose engine torque decreases at the same rate over time after an engine failure.

[0036] The specific test flight simulation method is as follows: during the test flight, the engine is manually shut down to simulate engine failure. After the target engine is shut down, the moment when the engine torque drops by 2% compared to before the failure represents the moment when the target engine fails; the pilot's reaction time to recognize the failure of the last engine is the time interval from the moment of engine failure to the moment the main pilot recognizes the engine abnormality.

[0037] Through flight test modifications, a quick-trigger button switch, named the "status switch," was installed on the pilot's collective pitch lever. When the pilot identifies an engine anomaly, they can press the "status switch." The moment the "status switch" is turned on represents the moment the pilot identifies the engine anomaly. Under normal cruising flight conditions, the aircraft shuts down one engine at a time until the last remaining normally functioning engine is reached. The co-pilot then abruptly shuts down the last engine without notifying the pilot. The time interval from the moment the engine fails to the moment the pilot identifies the engine anomaly and presses the "status switch" is recorded, representing the pilot's reaction time to identifying the last engine failure.

[0038] Determine the pilot's reaction time to recognize the last engine failure by taking the average value from at least five repeated test flights.

[0039] The pilot operation safety lag time in step 1 is set to 1 second. The pilot operation safety lag time is obtained according to airworthiness clauses AC29.79, AC29.75, and AC29.143. Preferably, the value is selected from airworthiness clause AC29.79.

[0040] In step 4, f(H) is obtained by aerodynamic calculation: the maximum lift coefficient of the airfoil under different altitude conditions is obtained by aerodynamic calculation, and then the maximum lift coefficient at 0m altitude is used as a reference, and the expression of f(H) is obtained through polynomial fitting.

[0041] In another embodiment, the step four f(H) is obtained through wind tunnel tests, and the maximum lift coefficient of the airfoil at different altitudes such as 0m, 1000m, ..., 6000m is obtained through wind tunnel tests. Then, the maximum lift coefficient at 0m altitude is used as a benchmark, and the expression of f(H) is obtained through polynomial fitting.

[0042] The present invention also provides a method for designing a multi-engine helicopter rotor. The design target of the rotor's moment of inertia needs to meet the rotor's moment of inertia calculated by the rotor's moment of inertia design method of the present invention, thereby ensuring that the helicopter has the autorotation landing safety characteristics required for airworthiness under the target design state.

[0043] In step 1 of the method of the present invention, the design target of the autorotation characteristic time is first determined, so that the forward design of the autorotation characteristic time can be carried out, and it is first associated with the human factor of the pilot, which solves the problem that the existing research can only rely on statistical equivalent design of existing helicopters and cannot explain the value of its design target.

[0044] According to helicopter aerodynamics theory, the required power increases as the speed decreases during takeoff and landing, and is the largest when hovering out of ground effect, representing the harsh state of the helicopter's autorotation transition. Therefore, in step 2, the required power for hovering out of ground effect under the target design state is selected as the characteristic state required power under the target design state.

[0045] As the rotational kinetic energy stored in the rotor system is released, the rotor speed decreases, and the thrust provided by the rotor at the same collective pitch decreases accordingly. The collective pitch must be increased to maintain rotor thrust and, in turn, to keep the helicopter flying. However, increasing the rotor collective pitch is not without limit. When the rotor collective pitch increases to the point where the rotor stalls, the rotor thrust no longer increases with the increase in rotor collective pitch, resulting in the rotor being unable to provide sufficient thrust to maintain helicopter flight. Therefore, a correction for rotor stall is required when calculating the releasable rotational kinetic energy per unit moment of inertia, E. In step three of this method, the rotor stall boundary is equivalently considered using the maximum lift coefficient of the main rotor blade airfoil at 70% radius in the calculation formula for the releasable rotational kinetic energy per unit moment of inertia, E.

[0046] The beneficial effects of the present invention are:

[0047] During the overall helicopter design phase, the helicopter's autorotation and landing safety characteristics must be designed based on the target design state to meet airworthiness requirements. Theoretical analysis and simulation studies have shown that the rotational kinetic energy stored in the helicopter rotor has a significant impact on the helicopter's autorotation and landing safety characteristics. This rotational kinetic energy is primarily determined by the rotor speed and the rotor system's moment of inertia. The rotor speed, which has a significant impact on the overall helicopter design, is subject to comprehensive constraints such as performance, quality, and payload, and cannot be arbitrarily changed. Therefore, the rotor's moment of inertia becomes a key design parameter for helicopter autorotation and landing safety. During the research process, we further analyzed the components that influence the rotor's moment of inertia. The rotor's moment of inertia is primarily contributed by the main rotor blades, which are primarily influenced by the blade's weight and center of gravity position. The heavier the main rotor blade and the further outboard the center of gravity is located, the greater the main rotor blade's moment of inertia. However, blade mass distribution is a critical parameter in blade dynamics design. Constrained by blade dynamics and loads, it has design limits within current material systems and dynamic design constraints, with most models operating at 50% to 60% of the radius. Therefore, increasing the rotor's moment of inertia often comes at the cost of increasing the main rotor blade weight, which in turn increases the aircraft's empty weight and reduces its performance in high-altitude missions.

[0048] The design method of the present invention can complete the target rotor moment of inertia design during the overall helicopter design phase based on the helicopter's target certification conditions, primarily the target certification altitude, temperature, and landing weight. It also determines whether the autorotation landing safety characteristics of the current overall technical solution meet the airworthiness requirements under the target certification conditions, and achieves a trade-off between the benefits of the autorotation landing safety characteristics and the weight of the rotor system. Specifically, it has the following technical effects:

[0049] 1. In step 1, the method of the present invention establishes the characteristic quantity "autorotation characteristic time" that describes the autorotation landing safety characteristics of a multi-engine helicopter and provides a design target for the "autorotation characteristic time." In step 5, the relationship between the autorotation landing safety characteristics of a multi-engine helicopter and the rotor moment of inertia is determined, and a theoretical formula and design method for the rotor moment of inertia design target are provided.

[0050] 2. In step 1, the method of the present invention proposes a method for determining the design target of "autorotation characteristic time", introducing the airworthiness requirement for the pilot's reaction time after engine failure, thereby ensuring that the autorotation landing safety characteristic analysis results of the method of the present invention meet the relevant airworthiness requirements.

[0051] 3. The theoretical formula and design method of the rotor moment of inertia design target of the method of the present invention introduces the required power of the characteristic state of the target certification condition through step 2, and introduces the influence of rotor stall at different altitudes on the rotational kinetic energy that can be released per unit moment of inertia of the rotor through steps 3 and 4, thereby taking into account the influence of the target certification conditions, including the target certification altitude, temperature and landing weight, and carrying out the target design of the rotor moment of inertia according to the performance requirements of plateau use. DETAILED DESCRIPTION

[0052] To make the purpose, technical solutions, and advantages of the embodiments of the present invention more clear, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0053] The following is an example of a certain type of helicopter to illustrate the specific process of the rotor moment of inertia design method based on the autorotation landing capability of the present invention. The basic data of the certain type of helicopter is as follows:

[0054] Rotor radius R = 9.45m;

[0055] Rotor speed Nr = 212 rpm;

[0056] Blade chord length b = 0.58m;

[0057] Number of blades n = 6;

[0058] The center of gravity of the main rotor blade is located at x = 5.00 m in the span direction.

[0059] The specific design steps are as follows:

[0060] Step 1: Determine the design target of the rotation characteristic time.

[0061] The pilot's reaction time to the last engine failure is determined by flight testing a helicopter with an engine similar to the target helicopter. The rate of torque reduction over time after engine failure for this similar engine is the same as that of the target engine for the target helicopter.

[0062] First, the test flight helicopter was modified for test flight. By adding a test system to the helicopter, the changes in the engine parameters on the helicopter over time can be recorded; during the test flight, the engine was manually shut down to simulate engine failure. After the target engine was shut down, the moment when the engine torque dropped by more than 2% compared to before the failure represented the moment when the target engine failed; a quick-trigger button switch was set on the main pilot's collective pitch lever, and the switch was named "status switch". When the pilot recognized that there was an abnormality with the engine, he could press the "status switch". The moment the "status switch" was turned on represented the moment when the pilot recognized the abnormality with the engine; the pilot's reaction time for recognizing the failure of the last engine was the time interval between the moment the last engine failed and the moment the main pilot pressed the "status switch".

[0063] A test flight test is then conducted. The helicopter is first in a normal cruising state with all engines operating normally. The co-pilot, without notifying the pilot, suddenly shuts down one of the engines. Upon recognizing the engine anomaly, the pilot immediately presses the "status switch" and begins emergency maneuvers to restore the helicopter to stable flight. This step is repeated until only the last engine is left operating normally. After the helicopter has restored stable flight with only the last engine operating normally, the co-pilot, without notifying the pilot, suddenly shuts down the last engine. Upon recognizing the last engine anomaly, the pilot immediately presses the "status switch" and begins emergency maneuvers to restore the helicopter to a stable autorotation state. After achieving stable autorotation, all engines are restored to normal operation, and the pilot maneuvers the helicopter out of autorotation and back to normal cruising.

[0064] Finally, the test results were statistically analyzed. Through 5 tests, 5 sets of valid data were obtained. The time results of the 5 valid test flights were 0.41 seconds, 0.45 seconds, 0.52 seconds, 0.48 seconds, and 0.44 seconds respectively. The reaction time for the pilot to recognize the failure of the last engine was 0.46 seconds.

[0065] Determine the pilot's safe operation lag time. Consider the airworthiness clauses AC29.79, AC29.75, and AC29.143, which require that there must be at least a 1-second lag between the failure of the last engine and the start of pilot control in multi-engine helicopters. The pilot's safe operation lag time is determined to be 1 second.

[0066] The pilot's reaction time of 0.46 seconds to recognize the failure of the last engine is compared with the pilot's operational safety lag time of 1 second, whichever is greater, and 1 second is selected as the design target of the autorotation characteristic time.

[0067] Step 2: Calculate the required power Pw of the characteristic state under the target design state

[0068] The target design state is selected as the altitude Hp = 4500m and the atmospheric temperature T = 10.75℃ (the air density under the altitude and temperature conditions can be obtained as ρ = 0.708kg / m 3 ), with a maximum landing weight Gw = 11,000 kg, and hovering out of ground effect as the characteristic state, the required power is calculated. Based on the flight performance calculation model established by aerodynamic theory, the hovering out of ground effect power Pw is 3,090 kW.

[0069] Step 3. Calculate the effect of rotor stall on the rotational kinetic energy released per unit moment of inertia of the rotor under the target design state:

[0070] Rotor solidity

[0071] Rotor angular velocity

[0072] Unit solidity tensile coefficient

[0073] The maximum lift coefficient C of the main rotor blade at 70% radius at sea level y7max is 1.2.

[0074] Therefore, the rotor's unit moment of inertia can release the following rotational kinetic energy E:

[0075] Step 4: Determine C y7max Correction function f(H) that varies with height.

[0076] Wind tunnel testing revealed that the maximum lift coefficient of the main rotor blade airfoil at 70% radius at sea level is 1.2, 1.164 at 1000 m, 1.128 at 2000 m, 1.092 at 3000 m, 1.056 at 4000 m, 1.02 at 5000 m, and 0.984 at 6000 m. A linear fit was performed on the test data, using the maximum lift coefficient at 0 m altitude as a benchmark, to obtain a correction function f(H) = -0.00003 × H + 1.

[0077] Step 5. Calculate the rotor moment of inertia I design target based on the autorotation landing capability.

[0078] Step 6: Calculate the rotor weight cost and weigh the overall helicopter technical solution

[0079] Through the above analysis of the present invention, the design target of the rotor moment of inertia I based on the autorotation landing capability is determined to be 31933.8 kg.m 2 , and the corresponding blade weight is 206.5kg.

[0080] Based on the blade weight, the overall technical solution of the helicopter is weighed. If it is considered that the blade weight cost is unacceptable and the blade weight must be reduced, return to step 2, lower the target certification altitude, atmospheric temperature, and the maximum landing weight Gw under the altitude and temperature conditions of the helicopter to be designed, and re-evaluate to reduce the blade weight. If it is considered that the blade weight cost is acceptable and it is desired to further expand the target design state, return to step 2, increase the target certification altitude, atmospheric temperature, and the maximum landing weight Gw under the altitude and temperature conditions of the helicopter to be designed, and re-evaluate to increase the blade weight until the overall technical solution of the helicopter meets the target.

[0081] The design method of the present invention considers the target certification conditions, especially the plateau environment and the airworthiness requirements of the autorotation and landing safety characteristics, during the overall helicopter design phase. It balances the benefits of the autorotation and landing safety characteristics with the weight of the rotor system, and completes the rotor moment of inertia design, ensuring that the overall technical solution of the helicopter can meet the airworthiness autorotation and landing safety characteristics requirements at the target certification altitude. Thus, a design method for high-altitude civilian helicopters is established for the unique plateau environment in China, considering the airworthiness requirements of the plateau autorotation and landing safety characteristics during the overall design phase, laying a solid foundation for the helicopter to ultimately obtain airworthiness certification for the plateau environment.

[0082] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, but the scope of protection of the present invention is not limited thereto. Any technician familiar with the field can easily think of various equivalent modifications or replacements within the technical scope disclosed by the present invention, and these modifications or replacements should all be covered by the scope of protection of the present invention.

Claims

1. A method for designing rotor moment of inertia based on autorotation landing capability, characterized in that: The following steps are involved: Step 1: Determine the design goal of the rotation characteristic time: The definition of autorotation characteristic time is the longest time that the energy stored in the rotor rotation kinetic energy can maintain the helicopter flight and keep the altitude without decreasing after the engine fails; the larger value of the pilot's reaction time to recognize the last engine failure and the pilot's operation safety lag time is taken as the design target of the autorotation characteristic time; the pilot's reaction time to recognize the last engine failure is the average reaction time of the pilot to recognize the abnormality of the engine after the test flight confirms the engine failure; Step 2: Calculate the required power Pw of the characteristic state under the target design state: Establish a flight performance calculation model, input the target design state, and calculate the required power Pw of the characteristic state under the target design state, which refers to the power required for the helicopter to maintain flight after the engine fails, and represents the decay rate of the rotational kinetic energy stored in the rotor: The target design state refers to the target certification altitude, atmospheric temperature, and maximum landing weight Gw under the altitude and temperature conditions of the helicopter to be designed; Step 3: Calculate the effect of rotor stall on the rotational kinetic energy released per unit moment of inertia of the rotor under the target design state: The formula for calculating the rotational kinetic energy E that can be released per unit moment of inertia of the rotor is as follows: C T The calculation formula of / σ is as follows: Among them C T is the pull coefficient, σ is the solidity of the rotor, g is the acceleration of gravity, ρ is the air density, Ω is the rotor Wing rotation angular velocity, R is the rotor radius, Gw is the takeoff weight, C y7max The maximum lift coefficient of the main rotor blade airfoil at 70% radius; Step 4: Determine C y7max Correction function f(H) that varies with height: f(H) is a correction function after considering the influence of altitude on the rotor stall characteristics, obtained through aerodynamic calculation or wind tunnel test to consider the influence of different altitude conditions on the aerodynamic characteristics of the blade, where H is the altitude; Step 5: Calculate the design target of the rotor moment of inertia I based on the autorotation landing capability. The rotor design needs to ensure that its rotor moment of inertia is greater than the design target: The rotation characteristic time in the formula is the design target of the rotation characteristic time; Step 6: Calculate the blade weight cost and weigh the overall helicopter technical solution: According to the position x of the blade's center of gravity in the blade's span direction and the rotor's moment of inertia I, the weight M of each blade in a rotor system with n blades is obtained. 桨叶 , and then the influence of the rotor moment of inertia I design target on the blade weight is obtained: x is the distance from the center of the hub to the center of gravity of the blade, in meters; the weight of the blade is corrected using the k value, where k is the proportion of the blade's moment of inertia in the rotor system's moment of inertia.

2. The method for designing the rotor moment of inertia according to claim 1, characterized in that: Step six also includes the step of further optimizing the overall technical solution: when it is necessary to further reduce the blade weight or further expand the target design state, return to step two to adjust the target design state and evaluate again until the overall technical solution of the helicopter reaches the target.

3. The method for designing the rotor moment of inertia according to claim 1, characterized in that: In step six, the value of k is 0.95-0.

97.

4. The method for designing the rotor moment of inertia according to claim 1, characterized in that: The pilot's reaction time for recognizing the last engine failure in step one is determined by conducting a test flight simulation on a helicopter using the same or similar engine as the helicopter to be designed; a similar engine is an engine for which the engine torque decreases at the same rate over time after an engine failure.

5. The method for designing the rotor moment of inertia according to claim 4, characterized in that: The specific method of test flight simulation is as follows: during the test flight, the engine is artificially shut down to simulate engine failure. After the target engine is shut down, the moment when the engine torque drops by 2% compared to before the failure represents the moment when the target engine fails; the pilot's reaction time to recognize the failure of the last engine is the time interval from the moment of engine failure to the moment when the main pilot recognizes the engine abnormality.

6. The method for designing the rotor moment of inertia according to claim 1, characterized in that: The pilot operation safety delay time is set to 1 second.

7. The method according to claim 3, characterized in that: The pilot operation safety lag time is obtained according to airworthiness clauses AC29.79, AC29.75 and AC29.

143.

8. The method for designing rotor moment of inertia according to claim 1, characterized in that: In step 4, f(H) is obtained by aerodynamic calculation: the maximum lift coefficient of the airfoil under different altitude conditions is obtained by aerodynamic calculation, and then the maximum lift coefficient at 0m altitude is used as a reference, and the expression of f(H) is obtained through polynomial fitting.

9. The method for designing rotor moment of inertia according to claim 1, characterized in that: The step four f(H) is obtained through wind tunnel tests. The maximum lift coefficient of the airfoil under different height conditions is obtained through wind tunnel tests. Then, the maximum lift coefficient at 0m height is used as a benchmark, and the expression of f(H) is obtained through polynomial fitting.

10. A method for designing a multi-engine helicopter rotor, characterized in that: The design target of the rotor's moment of inertia needs to meet the rotor's moment of inertia calculated by the design method described in claim 1, thereby ensuring that the helicopter has the autorotation landing safety characteristics required for airworthiness under the target design state.

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