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

Abstract

The present application belongs to the field of helicopter overall design, and relates to a rotor moment of inertia design method based on autorotation landing capability. The present application establishes a characteristic quantity "autorotation characteristic time" describing the autorotation landing safety characteristic of a multi-rotor helicopter, and gives a design target of the "autorotation characteristic time", thereby determining the relationship between the autorotation landing safety characteristic of the multi-rotor helicopter and the rotor moment of inertia, giving a theoretical formula and a design method of the rotor moment of inertia design target. The design method of the present application can complete the rotor moment of inertia target design according to the target certification conditions of the helicopter at the helicopter overall design stage, and can judge whether the autorotation landing safety characteristic of the current overall technical scheme can meet the airworthiness requirements under the target certification conditions, and complete the trade-off between the autorotation landing safety characteristic benefit and the rotor system weight.

Description

TECHNICAL FIELD

[0001] The application belongs to the field of helicopter overall design, and particularly relates to a rotor rotational inertia design method based on autorotation landing capability. BACKGROUND

[0002] The autorotation landing capability is an important safety feature of helicopters, which is different from other aircrafts. It means that the helicopter can obtain energy through the "windmill" effect in the descending process after all engine failures, so that the rotor can maintain a certain rotational speed, store energy, and finally complete safe landing by energy conversion to release the rotational kinetic energy of the rotor, reduce the descending speed and reduce the forward flight speed.

[0003] The autorotation landing capability is an important guarantee for the safety of the helicopter and the passengers after engine failure, and is one of the key contents of civil helicopter design and verification.

[0004] The airworthiness regulations have clear requirements for the autorotation landing safety of civil helicopters, and the flight performance and quality of the transition stage, stable autorotation stage and final landing stage after engine failure are all regulated. Among the above three stages, the pilot operates the helicopter to complete the transition into autorotation, and the instantaneous increase in distance from autorotation to landing before landing is the key period of energy conversion, which is the embodiment of the autorotation safety landing capability of the helicopter, that is, it is related to the proficiency of the pilot and the physical characteristics of the helicopter. Some helicopters have low requirements for the driving skills of the pilot, while some helicopters are dangerous even for experienced pilots, so it is necessary to carry out the analysis and design of the autorotation landing safety of the helicopter to ensure that the helicopter can also meet the airworthiness requirements for autorotation landing safety under the normal horizontal pilot operation.

[0005] The transition into autorotation and the instantaneous increase in distance before landing must consider the kinetic energy stored in the rotor and the kinetic energy consumption rate. The greater the rotational kinetic energy stored in the rotor, the slower the kinetic energy decay rate, and the better the autorotation characteristics of the helicopter; however, the rotational kinetic energy stored in the rotor mainly depends on the rotational inertia of the rotor.

[0006] Therefore, how to determine the rotor rotational inertia design target according to the target airworthiness certification height, and then balance the airframe weight cost and the autorotation landing capability benefit to obtain a reasonable and feasible technical scheme, becomes the key to the autorotation landing flight safety design.

[0007] In the prior art, the design evaluation method for the autorotation landing safety and whether the rotor moment of inertia meets the demand mainly has an evaluation method represented by simulation and simulation machine, and the evaluation method has high confidence, but has high cost and long time period, and cannot meet the demand of the rotor moment of inertia design rationality of the helicopter overall design stage rapid evaluation. In the helicopter overall design stage, there is still lack of a perfect rapid evaluation method for the autorotation landing safety of the multi-rotor helicopter, especially for the multi-rotor helicopter needing to meet the plateau environment airworthiness requirement, the prior art cannot rapidly judge whether the autorotation landing safety characteristic of the current overall technical scheme can meet the airworthiness requirement under the target certification condition, so that the rotor moment of inertia design target cannot be determined, and the trade-off between the autorotation landing safety characteristic benefit and the rotor system weight cannot be completed.

[0008] To establish the rotor moment of inertia design method based on the autorotation landing capability, the following three difficulties mainly exist:

[0009] Firstly, the characteristic quantity describing the autorotation landing safety characteristic of the multi-rotor helicopter needs to be established, the relationship between the autorotation landing safety characteristic of the multi-rotor helicopter and the rotor moment of inertia is determined, and the design target of the characteristic quantity is given, so that the rotor moment of inertia target design is obtained;

[0010] Secondly, the influence of the target certification condition, including the target certification height, temperature and landing weight, needs to be considered, and the rotor moment of inertia target design can be carried out for the plateau use performance requirement;

[0011] Thirdly, the airworthiness regulation requirement and the pilot human factor need to be associated to consider the influence of the airworthiness regulation on the rotor moment of inertia target design.

[0012] Therefore, it is urgent to establish a rotor moment of inertia design method based on the autorotation landing capability. SUMMARY

[0013] The purpose of the application is: aiming at the deficiency of the prior art, a rotor moment of inertia design method based on the autorotation landing capability is invented, so that for the multi-rotor helicopter needing to meet the plateau environment airworthiness requirement, the target certification condition requirement is introduced in the overall design stage, the autorotation characteristic time design target is determined through the test flight, the target certification condition characteristic state required power and the rotor unit solidity maximum tension coefficient calculation are introduced, so that the airworthiness requirement of the autorotation landing safety characteristic under the target certification condition, especially the plateau environment, is considered, the trade-off between the autorotation landing safety characteristic benefit and the rotor system weight is completed, the rotor moment of inertia design is completed, and it is ensured that the helicopter overall technical scheme can meet the airworthiness autorotation landing safety characteristic requirement under the target certification height.

[0014] To solve the technical problem, the technical scheme of the application is:

[0015] A rotor moment of inertia design method based on autorotation landing capability is provided, comprising the following steps:

[0016] Step one, determining the design target of autorotation characteristic time:

[0017] The definition of autorotation characteristic time is the longest time that the rotor kinetic energy can maintain the helicopter flight and keep the altitude from decreasing after the engine failure; the maximum value of the pilot's reaction time to identify the last engine failure and the pilot's safe lag time is taken as the design target of autorotation characteristic time; the pilot's reaction time to identify the last engine failure is the average reaction time of the pilot to identify the engine abnormality after the engine failure determined by flight test.

[0018] Step two, calculating the required power Pw of the characteristic state under the target design state: establishing a flight performance calculation model, inputting the target design state, and calculating the required power Pw of the characteristic state under the target design state, which is the required power of the helicopter after the engine failure, representing the decay rate of the rotor kinetic energy storage:

[0019] The target design state refers to the target certification altitude, atmospheric temperature of the helicopter to be designed, and the maximum landing weight Gw under the altitude and temperature conditions;

[0020] Step three, calculating the influence of rotor stall on the rotor unit moment of inertia releasable rotor kinetic energy under the target design state:

[0021] The rotor unit moment of inertia releasable rotor kinetic energy E calculation formula is as follows:

[0022]

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

[0024]

[0025] Wherein C T is the tension coefficient, σ is the rotor solidity, g is the gravity acceleration, ρ is the air density, Ω is the rotor rotation angular velocity, R is the rotor radius, Gw is the take-off weight, C y7max is the maximum lift coefficient of the main rotor blade airfoil at 70% radius;

[0026] Step four, determining C y7max The correction function f(H) changes with the altitude:

[0027] f(H) is the correction function considering the influence of altitude on rotor stall characteristics, which is obtained by aerodynamic calculation or wind tunnel test, in order to consider the influence of different altitude conditions on the aerodynamic characteristics of the blade, wherein H is the altitude;

[0028] Step five, calculate the design target of rotor moment of inertia I based on autorotation landing ability, rotor scheme design needs to ensure that its rotor moment of inertia is greater than the design target:

[0029]

[0030] In the formula, the autorotation characteristic time is the design target of the autorotation characteristic time;

[0031] Step six, calculate the blade weight cost, and make trade-off of the helicopter overall technical scheme:

[0032] According to the position x of the center of gravity of the blade in the span direction of the blade and the rotor moment of inertia I design target, the weight M of each blade in the rotor system with blade number n 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:

[0033]

[0034] x is the distance from the hub center to the center of gravity of the blade, in meters; k is the proportion of the blade moment of inertia in the rotor system moment of inertia.

[0035] Further, the method of the present application further includes the step of continuing to optimize the overall technical scheme in step six: when it is necessary to continue to 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 helicopter overall technical scheme reaches the target.

[0036] If the blade weight cost obtained in step six is unacceptable, the blade weight must be reduced, or the blade weight cost is acceptable and it is desired to further expand the target design state, then return to step two to adjust the target design state and evaluate again until the helicopter overall technical scheme reaches the target.

[0037] In step six of the method, the value of k is used to correct the weight of the blade, and k is 0.95-0.97; since most of the rotor moment of inertia is concentrated in the blade, the value of k is obtained from engineering practice experience, and is generally 95-97%.

[0038] The reaction time of the pilot to identify the last engine failure in step one is determined by flight test simulation on a helicopter using the same or similar engine as the helicopter to be designed.

[0039] The specific flight test simulation method is that: in the flight test, the engine is artificially closed to simulate engine failure, and the moment when the engine torque drops by 2% compared to before the failure represents the moment when the target engine fails; the reaction time of the pilot to identify the last engine failure is the time interval from the moment of engine failure to the moment when the pilot identifies the engine abnormality.

[0040] Through the flight test modification, a quick trigger button switch is arranged on the total distance lever of the main pilot, which is named as a "state switch", and the pilot can press the "state switch" when identifying the engine abnormality, and the moment when the "state switch" is connected represents the moment when the pilot identifies the engine abnormality. In the normal cruising flight state, one engine is closed each time until the last normally working engine is left, and the last engine is suddenly closed by the copilot without informing the main pilot, and the time interval from the moment of engine failure to the moment when the main pilot identifies the engine abnormality and presses the "state switch" represents the reaction time of the pilot to identify the last engine failure.

[0041] The reaction time of the pilot to identify the last engine failure is determined by averaging at least 5 repeated flight tests.

[0042] The pilot operation safety lag time in step one is 1 second. The pilot operation safety lag time is obtained according to the airworthiness clauses AC29.79, AC29.75 and AC29.143. Preferably, the value is selected from the airworthiness clause AC29.79.

[0043] The step four f(H) is obtained by aerodynamic calculation: the maximum lift coefficients of the airfoil under different height conditions are obtained by aerodynamic calculation, then the maximum lift coefficient at 0m height is taken as a reference, and a polynomial fitting is performed to obtain the expression of f(H).

[0044] In another mode, the step four f(H) is obtained by wind tunnel test: the maximum lift coefficients of the airfoil under different height conditions such as 0m, 1000m, …, 6000m are obtained by wind tunnel test, then the maximum lift coefficient at 0m height is taken as a reference, and a polynomial fitting is performed to obtain the expression of f(H).

[0045] The application also provides a design method of a multi-launch helicopter rotor, and the rotor moment of inertia design target needs to meet the rotor moment of inertia calculated by the rotor moment of inertia design method of the application, so as to ensure that the helicopter has the autorotation landing safety characteristics required by airworthiness under the target design state.

[0046] In step one of the method of the application, the design target of the rotation characteristic time is determined first, so that the forward design of the rotation characteristic time can be carried out, and the rotation characteristic time is associated with the pilot's human factors, which solves the problem that the existing research can only rely on the statistics of the existing helicopters for equivalent design and cannot explain the design target value.

[0047] According to the aerodynamics theory of helicopters, the required power increases with the decrease of the speed in the take-off and landing stages, and the required power is the largest in the ground effectless hovering, which represents the severe state of the rotation transition of the helicopter, so the required power in the ground effectless hovering in the target design state is selected as the characteristic state required power in the target design state in step two.

[0048] With the release of the rotational kinetic energy stored in the rotor system, the rotor speed decreases, and the pull provided by the rotor at the same total pitch also decreases, so the total pitch must be increased to maintain the flight of the helicopter, but the increase of the total pitch is not unlimited, when the total pitch of the rotor increases to the rotor stall, the pull of the rotor will no longer increase with the increase of the total pitch, which leads to the rotor being unable to provide enough pull to maintain the flight of the helicopter, so the rotor stall correction must be considered when calculating the rotational kinetic energy E that can be released by the unit moment of inertia of the rotor. In step three of the method, the rotor stall boundary is equivalent to the maximum lift coefficient of the airfoil at 70% radius of the main rotor blade in the calculation formula of the rotational kinetic energy E that can be released by the unit moment of inertia of the rotor.

[0049] The beneficial effects of the application are:

[0050] In the helicopter overall design stage, it is necessary to design the helicopter autorotation landing safety characteristics for the design target state to meet the airworthiness requirements. Through theoretical analysis and simulation research, the rotational kinetic energy stored in the rotor of the helicopter has an important influence on the autorotation landing safety characteristics of the helicopter, and the rotational kinetic energy stored in the rotor of the helicopter mainly depends on the rotational speed of the rotor and the rotational inertia of the rotor system. Among them, the rotational speed of the rotor has an important influence on the overall design of the helicopter, and needs to be comprehensively constrained by performance, quality and load, etc., and cannot be changed at will, so the rotational inertia of the rotor becomes a key design parameter of the autorotation landing safety design of the helicopter. In the research process, the composition affecting the rotational inertia of the rotor is further disassembled, the rotational inertia of the rotor is mainly provided by the main rotor blade, and the main rotor blade weight and the position of the center of gravity mainly affect the rotational inertia of the main rotor blade. The greater the weight of the main rotor blade, the farther the position of the center of gravity, and the greater the rotational inertia of the main rotor blade. However, the mass distribution of the blade is an important parameter of the blade dynamics design, which is constrained by the blade dynamics and load, and there is a design limit under the current material system and dynamics design constraint, and most models are at 50% to 60% radius. Therefore, in order to increase the rotational inertia of the rotor, the weight of the main rotor blade often needs to be increased, which leads to the increase of the weight of the whole machine and the decrease of the performance of the highland task.

[0051] The design method of the application can complete the rotor rotational inertia target design according to the target certification conditions of the helicopter, mainly the target certification height, temperature and landing weight, at the helicopter overall design stage, and judge whether the autorotation landing safety characteristics of the current overall technical scheme can meet the airworthiness requirements under the target certification conditions, and complete the trade-off between the autorotation landing safety characteristics and the weight of the rotor system. There are the following technical effects:

[0052] 1. The method of the application establishes the characteristic quantity "autorotation characteristic time" describing the autorotation landing safety characteristics of the multi-rotor helicopter in step one, and gives the design target of "autorotation characteristic time", so as to determine the relationship between the autorotation landing safety characteristics of the multi-rotor helicopter and the rotational inertia of the rotor in step five, and give the theoretical formula and design method of the rotor rotational inertia design target.

[0053] 2. The method of the application proposes a method for determining the design target of "autorotation characteristic time" in step one, introduces the lag requirement of airworthiness for the pilot reaction time after engine failure, so as to ensure that the autorotation landing safety characteristic analysis result of the method of the application meets the relevant requirements of airworthiness.

[0054] 3、The method of the application introduces the required power of the target evidence condition characteristic state through step two, and introduces the influence of the rotor stall at different altitudes on the released rotational kinetic energy of the unit rotational inertia of the rotor through step three and step four, so as to consider the influence of the target evidence conditions, including the target evidence altitude, temperature and landing weight, and to carry out the rotor rotational inertia target design according to the highland use performance requirements. DETAILED DESCRIPTION

[0055] To make the objectives, technical solutions and advantages of the embodiments of the application clearer, the technical solutions in the embodiments of the application will be described clearly and completely below in combination with the embodiments of the application. Obviously, the described embodiments are part of the embodiments of the application, rather than all the embodiments of the application. Based on the embodiments in the application, all other embodiments obtained by those skilled in the art without creative work fall within the protection scope of the application.

[0056] The specific process of the rotor rotational inertia design method based on the autorotation landing capability of the application will be demonstrated below by taking a certain type of helicopter as an example. The basic data of the certain type of helicopter are as follows:

[0057] The rotor radius R = 9.45 m;

[0058] The rotor rotational speed Nr = 212 rpm;

[0059] The blade chord length b = 0.58 m;

[0060] The number of blades n = 6;

[0061] The position of the center of gravity of the main rotor blade in the blade span direction x = 5.00 m;

[0062] The specific design steps are as follows:

[0063] Step one, determining the design target of autorotation characteristic time.

[0064] Through the test flight of the helicopter with the similar engine used by the certain type of helicopter to be designed, the reaction time of the pilot to identify the last engine failure is determined. The torque of the similar engine after the engine failure decreases at the same speed as the target engine used by the certain type of helicopter to be designed.

[0065] Firstly, the test flight helicopter is modified for test flight. By installing test system on the helicopter, engine parameters on the helicopter can be recorded over time. During the test flight, the engine is manually turned off to simulate engine failure. After the target engine is turned off, the moment when the engine torque drops more than 2% compared to before failure represents the moment of target engine failure. A quick trigger button switch is installed on the total distance lever of the main pilot. This switch is named "state switch". When the pilot identifies the engine anomaly, he can press the "state switch". The moment when the "state switch" is turned on represents the moment when the pilot identifies the engine anomaly. The reaction time of the pilot to identify the last engine failure is the time interval between the moment when the last engine fails and the moment when the main pilot presses the "state switch".

[0066] Then the test flight test is carried out. The helicopter is first in a normal cruising state with all engines working normally. The co-pilot suddenly turns off one of the engines without informing the main pilot. When the main pilot identifies the engine anomaly, he immediately presses the "state switch" and starts emergency handling to make the helicopter resume stable flight. Repeat this step until only the last engine is working normally. After the helicopter resumes stable flight with only the last engine working normally, the co-pilot suddenly turns off the last engine without informing the main pilot. When the main pilot identifies the last engine anomaly, he immediately presses the "state switch" and starts emergency handling to make the helicopter enter a stable autorotation state. After completing the entry into the stable autorotation state, all engines are restored to normal work, and the pilot controls the helicopter to exit the autorotation state and resume normal cruising.

[0067] Finally, the test results are statistically analyzed. Through 5 tests, 5 sets of effective data are obtained. The time results of 5 effective test flights are 0.41 seconds, 0.45 seconds, 0.52 seconds, 0.48 seconds and 0.44 seconds. The reaction time of the pilot to identify the last engine failure is 0.46 seconds.

[0068] The pilot's safe lag time is determined. Considering the requirements in airworthiness clauses AC29.79, AC29.75 and AC29.143 that there is at least a 1-second lag between the last engine failure and the pilot's start of operation for multi-engine helicopters, the pilot's safe lag time is determined to be 1 second.

[0069] The reaction time of the pilot to identify the last engine failure is 0.46 seconds, which is larger than the pilot's safe lag time of 1 second. Therefore, 1 second is selected as the design target of the autorotation characteristic time.

[0070] Step two, calculate the required power Pw of the characteristic state under the target design state

[0071] The selected target design state is the height Hp = 4500m, atmospheric temperature T = 10.75℃ (the air density under this height and temperature condition can be obtained as p = 0.708kg / m 3 ), the maximum landing weight Gw = 11000kg, and the hover without ground effect is selected as the characteristic state, and the required power is calculated. According to the flight performance calculation model established based on aerodynamics theory, the hover power Pw without ground effect is 3090kW.

[0072] Step three, calculate the influence of rotor stall on the rotor unit moment of inertia releasable rotational kinetic energy under the target design state:

[0073] Rotor solidity

[0074] Rotor rotation angular velocity

[0075] Unit solidity tension coefficient

[0076]

[0077] The sea level maximum lift coefficient C y7max of the main rotor blade airfoil at 70% radius is 1.2.

[0078] Therefore, the rotor unit moment of inertia releasable rotational kinetic energy E is as follows:

[0079]

[0080] Step four, determine the correction function f(H) of C y7max with altitude.

[0081] Through wind tunnel test, the sea level maximum lift coefficient of the main rotor blade airfoil at 70% radius is 1.2, the maximum lift coefficient at 1000m is 1.164, the maximum lift coefficient at 2000m is 1.128, the maximum lift coefficient at 3000m is 1.092, the maximum lift coefficient at 4000m is 1.056, the maximum lift coefficient at 5000m is 1.02, and the maximum lift coefficient at 6000m is 0.984. Linear fitting is performed on the test data, and the maximum lift coefficient at 0m height is taken as the reference to obtain the correction function f(H) = -0.00003xH + 1.

[0082] Step five, calculate the rotor moment of inertia I design target based on autorotation landing ability.

[0083]

[0084] Step six, calculate the rotor weight cost and perform helicopter overall technical scheme trade-off

[0085]

[0086] Through the above step analysis of the present application, the rotor moment of inertia I design target based on autorotation landing ability is determined as 31933.8 kg·m 2 , and the corresponding blade weight is 206.5 kg.

[0087] Based on the blade weight, the helicopter overall technical scheme is weighed, if the blade weight cost is considered unacceptable, the blade weight must be reduced, then return to step two, reduce the target certification height, atmospheric temperature of the helicopter to be designed, and the maximum landing weight Gw under the height and temperature conditions, and perform evaluation again, the blade weight can be reduced; if the blade weight cost is considered acceptable, and the target design state is desired to be further expanded, then return to step two, increase the target certification height, atmospheric temperature of the helicopter to be designed, and the maximum landing weight Gw under the height and temperature conditions, and perform evaluation again, the blade weight can be increased, until the helicopter overall technical scheme reaches the target.

[0088] The design method of the present application considers the target certification conditions, especially the plateau environment, the airworthiness requirements of autorotation landing safety characteristics, completes the trade-off between the autorotation landing safety characteristic benefits and the rotor system weight, completes the rotor moment of inertia design, and ensures that the helicopter overall technical scheme can meet the airworthiness autorotation landing safety characteristic requirements of the target certification height. Thus, the design method of the plateau civil helicopter is established for the domestic unique plateau environment, the airworthiness requirements of the plateau autorotation landing safety characteristics are considered in the overall design stage, and a solid foundation is laid for the helicopter to finally obtain the plateau environment airworthiness certificate.

[0089] Finally, it should be noted that: the above examples are only used to illustrate the technical solutions of the present application, but the protection scope of the present application is not limited thereto, any skilled person in the art can easily think of various equivalent modifications or replacements within the technical range disclosed by the present application, and these modifications or replacements should be covered in the protection scope of the present application.

Claims

1. A method for designing rotor rotational inertia based on autorotation landing capability, characterized in that: Includes the following steps: Step 1: Determine the design objective for the rotation characteristic time: The autorotation characteristic time is defined as the longest time that the energy stored from the rotor rotation kinetic energy can sustain helicopter flight and maintain altitude without decreasing after engine failure. The larger value between the pilot's reaction time to identify the failure of the last engine and the pilot's safe operation lag time is taken as the design target for the autorotation characteristic time. The pilot's reaction time to identify the failure of the last engine is the average reaction time after the test flight determines that the engine has failed and the pilot has identified the abnormality of the engine. Step 2: Calculate the required power of the characteristic states under the target design state. Establish a flight performance calculation model, input the target design state, and calculate the required power of characteristic states under the target design state. This refers to the power required for a helicopter to maintain flight after engine failure, and represents the rate at which the rotor's stored rotational kinetic energy decays. 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 altitude and temperature conditions. Step 3: Calculate the impact of rotor stall on the rotational kinetic energy that can be released per unit moment of inertia under the target design conditions: The formula for calculating the rotational kinetic energy E released per unit moment of inertia of a rotor is as follows: The calculation formula is as follows: in The tensile coefficient, For rotor realism, It is the acceleration due to gravity. air density, The rotor's angular velocity, The rotor radius is... Takeoff weight The maximum lift coefficient of the airfoil at 70% radius for the main rotor blade; Step 4: Confirm Correction function varying with altitude : The correction function, which takes into account the influence of altitude on rotor stall characteristics, is obtained through aerodynamic calculations or wind tunnel tests to account for the impact of different altitude conditions on the blade aerodynamic characteristics. H For height; Step 5: Calculate the design target for the rotor rotational inertia I based on autorotation landing capability. The rotor design must ensure that its rotor rotational inertia is greater than this design target. In the formula, the rotation characteristic time is the design target for the rotation characteristic time; Step Six: Calculate the cost of rotor blade weight and weigh the overall technical options for the helicopter. Based on the position of the blade's center of gravity along the blade's span. x The number of blades was obtained by calculating the rotor rotational inertia I as the design objective. n The weight of each blade in the rotor system This leads to the influence of the rotor rotational inertia I design target on the blade weight: x The distance from the center of the propeller hub to the center of gravity of the propeller blades, in meters; using k The value is corrected for the weight of the blades. k This represents the proportion of the blade's moment of inertia to the rotor system's moment of inertia. Step six also includes steps to further optimize 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 conduct another evaluation until the overall technical solution of the helicopter achieves the target.

2. The rotor rotational inertia design method according to claim 1, characterized in that: In step six k The value ranges from 0.95 to 0.

97.

3. The rotor moment of inertia design method according to claim 1, characterized in that: In step one, the pilot's reaction time to identify the failure of the last engine is determined by conducting flight simulations on helicopters that use the same or similar engines as the helicopter to be designed; similar engines are those whose engine torque decreases at the same rate over time after engine failure.

4. The rotor moment of inertia design method according to claim 3, characterized in that: The specific method of flight test simulation is as follows: During the flight test, the engine is deliberately shut down to simulate engine failure. The moment when the engine torque drops by 2% compared to before the failure after the target engine is shut down represents the instant when the target engine fails. The reaction time for the pilot to identify the failure of the last engine is the time interval from the moment the engine fails to the moment the pilot identifies the engine abnormality.

5. The rotor rotational inertia design method according to claim 1, characterized in that: The pilot's safe operation delay time is set to 1 second.

6. The method according to claim 2, characterized in that: The pilot operation safety delay time is obtained according to airworthiness provisions AC29.79, AC29.75, and AC29.

143.

7. The rotor moment of inertia design method according to claim 1, characterized in that: Step four Aerodynamic calculations were performed to obtain the maximum lift coefficient of the airfoil under different altitude conditions. Then, using the maximum lift coefficient at 0m altitude as a benchmark, polynomial fitting was used to obtain... The expression.

8. The rotor moment of inertia design method according to claim 1, characterized in that: Step four The maximum lift coefficient of the airfoil was obtained through wind tunnel tests at different altitudes. Then, using the maximum lift coefficient at 0m as a benchmark, polynomial fitting was performed to obtain... The expression.

9. A design method for a multi-engine helicopter rotor, characterized in that: The design target for the rotor's moment of inertia needs to meet the rotor 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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