A method of controlled stall lift distribution control for an aircraft

Abstract

The application discloses a controlled stall lift distribution control method of an aircraft, and belongs to the field of aircraft flight control technology. The method makes the fuselage angle of attack greater than the wing stall door and the wing lift saturated on the upper limit of the available lift coefficient of the wing at a low speed section with air speed lower than a threshold value, and the continued increasing lift demand is accepted by the fuselage as a central lift body; the aerodynamic lift is calculated according to the air speed and the fuselage angle of attack, the vertical available thrust of the rotor is calculated according to the power budget and the state of the rotor disc, the rotor lift margin is taken as the only envelope protection criterion and the exit criterion of the state, the wing angle of attack stall warning and the downgrade link of the door are removed from the protection, and the target value of the fuselage angle of attack is given by the lower limit of the margin according to the monotonic relationship between the aerodynamic lift and the angle of attack, and the pitch is controlled according to the target value, and the exit is performed when the margin is lower than the lower limit. The application makes the aerodynamic load share rise, the rotor load share drop, and the criterion not be triggered by mistake.

Description

Technical Field

[0001] This invention relates to the field of aircraft flight control technology, specifically to a controlled stall lift distribution control method for an aircraft in the low-speed range where the lift is shared by the wings and the central lifting body, and a control system for implementing this method. Background Technology

[0002] Existing tail-seat compound wing aircraft combine vertical takeoff and landing (VTOL) and fixed-wing cruise capabilities, with the fuselage, wings, and rotor systems each contributing to overall lift. Conventional flight control systems treat the "wing angle of attack must not exceed the stall threshold" as an insurmountable hard protection: once the wing angle of attack or the overall aircraft angle of attack approaches the stall threshold, an alarm is triggered, amplitude is limited, leveling off, or the system is downgraded. This rule holds true in conventional configurations where the wing is the primary lift surface and the rotor only provides propulsion and attitude control. However, in configurations where rotor thrust is traded for payload capacity, it demarcates the entire low-speed, high-load range as a no-go zone.

[0003] The physical quantities selected for the aforementioned hard protection are inappropriate. The consequence of wing stall is that wing lift no longer increases with angle of attack, not that wing lift disappears; the lift already gained is retained after stall. Using angle of attack as a criterion will misjudge the state of "lift saturation but still contributing" as a critical state. What truly determines whether an aircraft can maintain flight is who bears the continued increase in lift demand and how much vertical thrust the rotor still has.

[0004] Correspondingly, the design approach for low-speed applications generally treats rotor thrust as the sole load-bearing source: selecting the upper limit of power for the rotor and motor based on the worst hovering conditions, and using rotor thrust alone to support the entire weight of the aircraft. The structural weight and battery weight increase along with the upper limit of power, the load capacity is squeezed, and the wings and fuselage are excluded from the angle of attack range where they could contribute lift due to protection logic.

[0005] Furthermore, in situations where the load needs to be transferred from the wing to the rotor, the conventional approach is to rely on angle of attack or airspeed schedules to determine the handover point of the lift source; misjudgment of the handover point and the two lift sources competing with each other within the same control phase are both existing failure modes.

[0006] The above practices are underpinned by two technical biases in the field: first, the view that "wing stall equals danger," meaning that wing stall is always considered an abnormal state that must be immediately resolved; and second, the view that "rotor is only responsible for propulsion and attitude control and does not participate in lift distribution," meaning that rotor thrust is excluded from the overall aircraft lift calculation. These two prejudices together lead to the design of the low-speed, high-load segment being solely borne by the rotor. Summary of the Invention

[0007] The technical problem to be solved by this invention is to provide a controlled stall lift distribution control method for an aircraft, so that the low-speed, high-load segment can be set as a stable operating point under the condition that the wing has entered a stall and the fuselage angle of attack is greater than the wing stall gate, and the envelope protection criterion of this operating point falls on the rotor lift margin rather than on the angle of attack, thereby increasing the aerodynamic load share of the low-speed segment without increasing the upper limit of rotor and motor power, and eliminating failures such as the conflict between protection logic and normal operating point.

[0008] To solve the above-mentioned technical problems, the present invention adopts the following technical solution: a controlled stall lift distribution control method for an aircraft, wherein the aircraft includes a fuselage, wings, and a rotor system, the fuselage constitutes the central lifting body of the aircraft, the difference between the stall angle of attack of the wing and the angle of installation of the wing relative to the fuselage constitutes the stall gate of the wing, and the stall gate is a threshold value in the form of the fuselage angle of attack; the aircraft is in a controlled stall state in the low-speed segment where the airspeed is below a preset threshold, and in the controlled stall state, the angle of attack of the fuselage is greater than the stall gate of the wing and the angle of installation of the rotor system. The lift of the wing is saturated at the upper limit of the available lift coefficient of the wing, and the increased lift demand is borne by the central lifting body; including the following steps: Step 1: Real-time acquisition of the airspeed, fuselage angle of attack and total weight of the aircraft; calculation of the current aerodynamic lift of the aircraft based on the airspeed and fuselage angle of attack; and calculation of the rotor system's vertically available thrust based on the power budget and rotor disk status; Step 2: Calculation of the rotor lift margin, which is the difference between the rotor's vertically available thrust and the rotor's vertically required thrust, and the difference between the rotor's vertically available thrust and the vertically required thrust. Using the thrust ratio, the vertical thrust required by the rotor is determined by the difference between the total weight of the aircraft and the current aerodynamic lift. The rotor lift margin is used as the sole envelope protection criterion and the sole exit criterion for the controlled stall state. The aircraft does not use the fuselage angle of attack or the wing angle of attack as the envelope protection criterion, nor does it set an intervention boundary based on the wing stall gate or stall speed. The criterion chain consisting of wing angle-of-attack stall alarms and the gate-crossing degradation link is removed from the envelope protection of the controlled stall state and replaced by the rotor lift margin. The fuselage angle of attack is large. The wing stall gate thus does not constitute an abnormal event; Step 3: The target value of the fuselage angle of attack is obtained by inversely solving the relationship between the lower limit of the rotor lift margin and the monotonic change of aerodynamic lift with the fuselage angle of attack. The target value of the fuselage angle of attack is given by inversely solving the lower limit of the rotor lift margin, rather than by a preset modal angle of attack or airspeed plan table; Step 4: The pitch of the aircraft is controlled according to the target value of the fuselage angle of attack, so that the fuselage angle of attack tracks the target value of the fuselage angle of attack; and the controlled stall state is exited when the rotor lift margin is lower than the lower limit.

[0009] Further, in step 3, the inverse solution takes the aerodynamic lift corresponding to the lower limit of the rotor lift margin as the inverse solution target. According to the relationship between the aerodynamic lift and the fuselage angle of attack at the current airspeed, the minimum fuselage angle of attack that makes the aerodynamic lift not less than the inverse solution target is obtained, and the minimum fuselage angle of attack is used as the target value of the fuselage angle of attack.

[0010] Further, in step 2, the wing angle-of-attack stall alarm and the gate-crossing degradation link are removed from the envelope protection of the controlled stall state. Specifically, during the controlled stall state, the aircraft sets a modal declaration flag, which prevents the wing angle-of-attack stall alarm, the gate-crossing degradation link, and the intervention boundary calibrated according to the wing stall gate from participating in the calculation of the envelope protection criteria and exit criteria, and prohibits the above items from rearming during the controlled stall state. During the controlled stall state, only one envelope protection chain with the rotor lift margin as input is retained, and the output of the envelope protection chain is the command to exit the controlled stall state. The entry and exit of the controlled stall state are determined according to the lower limit of the rotor lift margin and its hysteresis.

[0011] Furthermore, it also includes a benefit availability criterion: under the controlled stall state, the proportion of the aerodynamic lift after deducting the lift contributed by the central lifting body to the total weight of the aircraft is calculated online to obtain the aerodynamic load share after removing the central lifting body, and the rotor load share required by the rotor system to make up the aerodynamic load share is calculated; when the rotor load share exceeds the preset share, it is determined that the benefit of the controlled stall state is unavailable and the controlled stall state is exited.

[0012] Furthermore, in step 3, the inverse solution also makes the target value of the fuselage angle of attack fall into the following angle of attack range, in which the proportion of the central lifting body in the total lift of the aircraft for each unit increase in fuselage angle of attack is greater than the proportion of the wing.

[0013] Furthermore, in step 4, the conditions for exiting the controlled stall state also include the fuselage angle of attack reaching the upper limit of the angle of attack calibrated according to the effective domain of the aerodynamic characteristics of the central lifting body and the pitch trim capability of the aircraft; exiting the controlled stall state when the airspeed enters the aerodynamic full load zone and the vertical thrust share borne by the rotor system is lower than a preset share; and exiting the controlled stall state when a modal exit command is received from the mission layer or the pilot.

[0014] Furthermore, the rotor system includes a forward propulsion rotor assembly and a lift rotor assembly; in steps 1 and 2, the available vertical thrust of the rotor and the required vertical thrust of the rotor are decomposed into vertical components according to the ground thrust direction of each rotor disk of the forward propulsion rotor assembly and the lift rotor assembly, and then synthesized; the forward propulsion rotor assembly and the lift rotor assembly share the required vertical thrust of the rotor according to the thrust ratio that makes the pitch moment of the entire aircraft zero, and the thrust ratio is determined by the ratio of the pitch gain of the rotor disk center of the forward propulsion rotor assembly and the lift rotor assembly relative to the center of gravity of the entire aircraft.

[0015] Furthermore, in the controlled stall state, the roll channel is undertaken by the equal and opposite thrust differential of the lifting rotors symmetrically arranged on the left and right sides of the rotor system, and the aileron deflection authority of the aircraft is less than the roll authority of the equal and opposite thrust differential; the roll channel does not hand over when entering and exiting the controlled stall state.

[0016] Furthermore, in steps 3 and 4, the target value of the fuselage angle of attack increases the weight share borne by the aerodynamic lift to above a preset share, and reduces the vertical thrust share borne by the rotor system to below the share corresponding to the lower limit of the rotor lift margin; in the controlled stall state, the rotor system is in a low thrust trim state, and the vertical thrust of the rotor system only bears the vertical component corresponding to the difference between the total weight of the aircraft and the current aerodynamic lift, as well as the trim and control requirements of the aircraft.

[0017] This invention also provides a controlled stall lift distribution control system for an aircraft, the aircraft comprising a fuselage, wings, and a rotor system, the fuselage constituting the central lifting body of the aircraft, the difference between the stall angle of attack of the wing and the angle of installation of the wing relative to the fuselage constituting the stall gate of the wing, the stall gate being a threshold value in the form of the fuselage angle of attack, comprising: a flight control computer; an attitude inertial measurement unit, an airspeed sensor, and an angle of attack sensor, respectively connected to the signal input terminal of the flight control computer; a thrust readback device for reading back the actual thrust of each rotor in the rotor system, the thrust readback device being at least one of the telemetry readback of the electronic speed controller of the rotor system, a current sensor, a voltage sensor, a speed sensor, and a thrust sensor, the signal output terminal of the thrust readback device being connected to the signal input terminal of the flight control computer; each rotor of the rotor system is driven by an electronic speed controller, the control input terminal of each electronic speed controller being connected to the control output terminal of the flight control computer; the flight control computer is configured to execute claims 1 to 9. The controlled stall lift distribution control method described in any one of the above.

[0018] The beneficial effects of this invention are as follows: First, the criterion replacement prevents the protection logic from conflicting with the normal operating point—in the controlled stall state, the rotor lift margin is used as the sole envelope protection criterion and the sole exit criterion. Wing angle-of-attack stall alarms, gate-crossing degradation links, and intervention boundaries calibrated according to the wing stall gate are removed from this envelope protection. The fuselage angle of attack greater than the wing stall gate no longer constitutes an abnormal event, and the lift retained after wing stall is included in the overall aircraft lift account. Second, the operating point is calculable and retractable—the target value of the fuselage angle of attack is given by inversely solving the relationship between the lower limit of the rotor lift margin and the monotonic change of aerodynamic lift with the fuselage angle of attack, without the need for calculation based on air... The speed plan table or preset modal angle of attack values ​​do not have any misjudgment of the lift source handover point; third, the revenue can be verified online - the load share that the rotor system must make up after removing the central lifting body is calculated online, and this criterion can be turned off, so its conclusion is falsifiable; fourth, no new hardware is added - the central lifting body, wing and rotor system are all existing airframe structures, and the above criteria can be directly calculated from the existing overall weight, airspeed, fuselage angle of attack estimation and rotor thrust readback; fifth, it is compatible with the lateral passage - the roll passage uses the equal reverse thrust differential of the left and right symmetrically arranged lifting rotors, and this state does not require a new lateral passage handover.

[0019] Furthermore, according to design values, the controlled stall state brings about a directional change in load-bearing share at low speeds: by increasing the fuselage angle of attack from the conventional trim to the upper limit of the angle of attack in this state along the same rotor load line, the weight share borne by aerodynamic lift increases from approximately 9.9% to approximately 19.0%, while the share of vertical thrust borne by the rotor system decreases from approximately 90.1% to approximately 81.0%. For example, the vertical thrust required by the rotor can be reduced by approximately 44.4%. N indicates that selecting rotors and motors with lower power limits can cover operating conditions that would otherwise require higher power limits, allowing for more load at the same power. In this state, the rotor lift margin is approximately 0.382, higher than the approximately 0.237 in hover, and monotonically increases with airspeed to approximately 0.999, meaning this state is safer than hovering. After airspeed reaches the aerodynamic full-load region, the rotor returns to a low-thrust trim state, with the power consumption ratio relative to hovering estimated at approximately 0.135~0.194, representing a power saving of approximately 5.2~7.4 times. The above margins and power consumption are design estimates. The rotor's vertical thrust can be calculated using momentum theory with hovering power as the anchor point, not based on actual rotor bench measurements. The above values ​​are only used to illustrate ratios and trends and are not intended to limit the protection range.

[0020] It should be noted that the claims of this invention and the applicant's related applications are as follows: The claims of making the fuselage angle of attack exceed the wing stall angle of attack and putting the wing in a stall state, and of pulling the wing angle of attack past its stall angle of attack during hovering mode, takeoff hovering mode, and the transition from cruise to hover, with the power unit or lift rotor bearing the overall lift, have already been claimed in prior applications. This invention does not reiterate the state of "permissible stall" or "controlled stall" itself, nor does it reiterate the action of pulling the angle of attack past the stall angle of attack itself. This invention claims that this state is set as a stable operating point in the low-speed range, and that the rotor lift margin is used as the criterion for this operating point; after the wing stalls, the fuselage lift body continues to provide lift, and the overall load is maintained by the fuselage and rotor together. As described in a prior application of the same family, this invention only references the existing aerodynamic characteristics of this configuration and does not reiterate the aerodynamic characteristics themselves. This invention proposes using the marginal allocation of the lift increment after wing lift saturation as the design basis for the operating point of this mode, and based on this, using the margin to solve for the target value of the fuselage angle of attack; the resultant force of the rotor's vertical thrust components compensates for the difference between the overall weight and the lift shared by the wing and horizontal stabilizer. This has been claimed in the thrust distribution and trim equations of a prior application of the same family. This invention does not reiterate this trim relationship, but only uses this difference as the defined quantity of the vertical thrust required by the rotor, and uses the ratio of this difference to the available vertical thrust of the rotor as the unique envelope protection criterion and exit criterion for this mode; lift rotor. The stall fallback power protection and its warning and intervention boundaries calibrated by stall angle of attack or stall speed, the switching of the power load state after stall, the stick control handling aimed at reducing the angle of attack to a safe angle of attack, and the transition trigger-prevention logic during modal transition that suspends stall criteria based on airspeed and rate of climb / drop while retaining stall criteria based on angle of attack and temporarily relaxes its threshold are all claimed by prior applications in the same family. The distinction between this invention and those prior applications in the same family is that this invention does not set intervention boundaries calibrated by stall angle of attack or stall speed, does not set a stick control chain to reduce the angle of attack to a safe angle of attack, and does not use wing angle of attack or fuselage angle of attack as envelope protection criteria. Wing angle of attack exceeding the threshold does not constitute an abnormal event or trigger a stall alarm in this invention. It does not trigger a door-crossing degradation; the power load-bearing state of the aforementioned prior application is that the entire weight of the aircraft is borne by the lift rotor system and the propulsion system, while in this invention, the weight share borne by aerodynamic lift increases and the vertical thrust share borne by the rotor decreases in this mode, with the two changing in opposite directions in terms of load share; the vertical thrust component of the rotor is used to make up for the lift gap and thrust distribution is performed accordingly, and the fuselage angle of attack is used to distinguish between several modes and to perform mode conversion, and the arrival criterion and abnormal retreat defined by the transition corridor are all claimed by the prior application of the same family. This invention does not repeat the claims of mode set, transition corridor, conversion feasibility judgment and abnormal retreat, nor does it repeat the claim of giving pitch target with preset mode angle of attack;The absence of ailerons on the trailing edge of the wing, the absence of elevators on the trailing edge of the horizontal stabilizer, and the differential thrust between symmetrically arranged thrust units in the roll channel are all claimed in prior applications of the same family. This invention does not repeat the claims of these higher-level contents. The lateral channel in this mode is a direct adoption of the channel claimed in the prior application of the same family. This mode does not involve a new lateral channel handover. The tail lift rotor assembly remains operational in cruise mode and undertakes lift sharing and pitch trim, as well as generating pitch moment by synchronously increasing or decreasing the thrust of the left and right lift rotors. These are all claimed in prior applications of the same family. This invention does not repeat the claims of pitch channel and cruise lift sharing allocation. The movable load that translates along the longitudinal axis of the fuselage, its load position adjustment device and linear motor drive mechanism, the adjustment of the longitudinal projection of the aircraft's center of gravity according to the torque balance relationship, the setting of the aircraft's center of gravity position according to the static stability margin target, and the driving of the center of gravity to move towards the dynamic load balance position when stall intervention or failure handling are all claimed by prior applications in the same family. This invention does not repeat the claim of the center of gravity adjustment device itself, the setting of the static stability margin target, or the movement of the center of gravity towards the dynamic load balance position; the tail rotor tilting mechanism and its tilting angle inversely proportional to the aircraft's angle of attack, the tail oblique cut vector attitude propeller and control distribution, and the tail rotor The dynamic horizontal stabilizer layout and the horizontal stabilizer-mounted lift rotor layout with the upper surface of the horizontal stabilizer as the mounting reference surface for the lift rotor drive motor are both claimed in prior applications of the same family. This invention does not involve tilting mechanisms, tilt angle linkages, oblique-cut vector attitude propellers, or the tail dynamic horizontal stabilizer layout, nor does it reiterate the horizontal stabilizer-mounted lift rotor layout itself. The boundary layer suction, equipment cooling, and integrated propulsion and exhaust internal flow system, the detection, latching, and graded handling method for single rotor failure in a tail-seat coaxial twin-rotor aircraft, the retractable landing gear retraction structure, and the mounting and deployment structure for the belly-embedded mission pod are all claimed in prior applications of the same family. This invention does not reiterate the claims mentioned above. The internal flow theme, rotor failure detection and handling methods, landing gear structure and mission pod structure only refer to their parking and high angle of attack landing conditions as environmental characteristics; the two appearance design applications for the landing gear and the whole aircraft only submit views and brief descriptions, without disclosing the structure and principle, and do not constitute a technical disclosure of this invention; the difference between this invention and another scheme in the same family that uses rotor differential to replace ailerons to undertake roll is that the latter scheme solves the lateral channel of the aileron-less configuration, while this invention solves the longitudinal and lift distribution channel. The two are complementary divisions of labor for the same aircraft rather than the same path. This invention does not advocate the distribution law of the roll channel, nor does it advocate the differential application method and gain ratio;This invention proposes that, in the low-speed range where airspeed is below a preset threshold, the target value of the fuselage angle of attack is set to be greater than the wing stall threshold, intentionally saturating the wing lift to its upper limit of available lift coefficient, and ensuring that any further increase in lift demand is borne by the central lifting body. The ratio of the difference between the rotor's vertically available thrust and the rotor's required vertical thrust to the available vertical thrust is used as the unique envelope protection criterion and exit criterion for this mode. The required vertical thrust is determined by the difference between the overall weight and the current aerodynamic lift. The target value of the fuselage angle of attack is inversely solved from the lower limit of this ratio or a given rotor load share according to the monotonic relationship between aerodynamic lift and angle of attack. This criterion ensures that exceeding the wing angle of attack threshold does not constitute an abnormal event. Attached Figure Description

[0021] Figure 1 This is a schematic diagram of lift distribution in an embodiment of the present invention, showing the three-source division of labor: wing lift saturation, central lifting body bearing the load, and rotor supplementing the lift margin according to the rotor's lift margin.

[0022] Figure 2 This is a grouped lift-angle of attack curve diagram of an embodiment of the present invention, showing that after the wing stall gate, the lift saturates, the lift of the central lifting body increases with the fuselage angle of attack, and the overtake point to which the margin of the lift increment belongs and the vertical line of the wing stall gate are marked.

[0023] Figure 3 The load-airspeed curve diagram of this invention shows the changes of aerodynamic load share and rotor load share with airspeed, and marks the upper limit line of rotor load share, aerodynamic full load point, and comparison curve of rotor bearing load alone.

[0024] Figure 4 This is a block diagram of the envelope protection criterion in an embodiment of the present invention. It shows an envelope protection chain with rotor lift margin as the only input, the wing angle-of-attack stall alarm and gate-crossing degradation link that is removed from the envelope protection chain, and the exit conditions.

[0025] Figure 5 This is a comparison diagram of lifting body removal in an embodiment of the present invention. It shows the aerodynamic load share of the aircraft in three cases: full model, removal of the central lifting body, and removal of the horizontal tail. The load share that the rotor system must make up after removal is also marked.

[0026] The attached diagrams in this case are line graphs and block diagrams. The diagrams do not have any attached figure labels, so there is no attached figure label for comparison. Detailed Implementation

[0027] The present invention will be further described below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only for explaining the present invention and are not intended to limit the present invention. The overall weight and geometric parameters in the embodiments are design values ​​at the prototype level, and the grouped aerodynamic data, load share, and power consumption are design values ​​and design estimates, used to illustrate the grouped composition of lift at each angle of attack, the direction of change of load share, and the relative relationship of each criterion, rather than limiting the scope of protection of the present invention.

[0028] As shown in Figure 1, the aircraft in this embodiment has a tail-seat compound wing layout, including a fuselage, wings, and a rotor system. The fuselage constitutes the central lifting body of the aircraft. The aircraft in this embodiment also includes a flight control computer, and an attitude inertial measurement unit, airspeed sensor, angle of attack sensor, and thrust readback device connected to the signal input terminal of the flight control computer. Each rotor of the rotor system is driven by an electronic speed controller, and the control input terminal of each electronic speed controller is connected to the control output terminal of the flight control computer. The attitude inertial measurement unit provides the aircraft's attitude and attitude angular velocity, the airspeed sensor provides the aircraft's airspeed, the angle of attack sensor provides the fuselage angle of attack, and the thrust readback device reads the actual thrust of each rotor in the rotor system. Design parameters: takeoff weight of 490.3 N; wing reference area of ​​0.6350 m², mean aerodynamic chord of 0.3203 m, wingspan of 2.1934 m, length of 1.0962 m; wing angle of installation relative to fuselage of 8°. The rotor system includes a lifting rotor assembly and a forward propulsion rotor assembly. The lifting rotor assembly includes two sets of coaxial twin rotors, each set consisting of two rotors with coincident rotation axes and opposite rotation directions. The rotor diameter is approximately 609.6 mm according to the design. The forward propulsion rotor assembly includes two propulsion rotors located at the left and right wingtips of the aircraft, with a rotor diameter of approximately 711.2 mm according to the design. The lower fuselage of the aircraft is equipped with a boomerang-shaped landing gear to support the aircraft during high angle-of-attack landings under controlled stall conditions. Figure 1 illustrates the three-source division of lift in this embodiment: under controlled stall conditions, the lift of the wing is saturated at its upper limit of available lift coefficient and no longer increases with the angle of attack of the fuselage. The continued increase in lift demand is borne by the central lifting body, and the vertical thrust of the rotor system supplements the remaining vertical gap according to the rotor lift margin.

[0029] As shown in Figure 2, the wing stall gate in this embodiment is determined as follows: the wing stall angle of attack is approximately 17° based on the wing diameter, and the wing's angle of attack relative to the fuselage is 8°. The difference of 9° constitutes the wing stall gate, which is a threshold value in the form of the fuselage angle of attack. The wing chord angle of attack is equal to the sum of the fuselage angle of attack and the wing's angle of attack; therefore, the wing enters a stall state when the fuselage angle of attack is greater than 9°. Checking the fuselage angles of attack at 8° and 12°: at 8°, the corresponding wing chord angle of attack is 16°, and the wing has not yet crossed the threshold; at 12°, the corresponding wing chord angle of attack is 20°, and the wing has crossed the threshold, consistent with this threshold. In this embodiment, the upper limit of the usable lift coefficient on the wing side is set to 1.10~1.15 according to the design value, and the overall lift coefficient estimated according to this caliber is approximately 1.35~1.45. It should be noted that this value is based on the wing-only stall criterion and is not directly comparable to the overall lift coefficient given by the grouped lift account below, as the latter also includes the contributions of the central lifting body and the horizontal stabilizer. The preset threshold for the low-speed range is calibrated based on the airspeed at which the overall aerodynamic lift just bears the weight of the aircraft at the operating angle of attack in the controlled stall state. According to the design value, this airspeed is approximately 27.3 m / s. In the low-speed range where the airspeed is below this threshold, the flight control is set to a controlled stall state. At this time, the fuselage angle of attack is greater than the wing stall threshold, the wing lift is saturated at its upper limit of usable lift coefficient, and the increased lift demand is borne by the central lifting body. In this embodiment, the fuselage angle of attack operating point in this state is 12°~18°.

[0030] As shown in Figure 2, the grouped lift coefficients given in this embodiment according to the design values ​​are shown in the table below. The angle of attack of the wing chord is equal to the sum of the fuselage angle of attack and the wing installation angle (8°). Each item represents the proportion of the lift coefficient in that group to the total aircraft lift coefficient. The wing slope ratio is the ratio of the slope of the wing lift coefficient with respect to the fuselage angle of attack to its slope in the non-stall segment. 0° 8° 0.4003 0.0899 −0.0271 0.4631 86.4% 19.4% 1.000 Did not cross the door 8° 16° 0.7803 0.3009 0.0230 1.1042 70.7% 27.3% 0.801 Did not cross the door 12° 20° 0.9112 0.3967 0.0470 1.3549 67.3% 29.3% 0.227 Crossing the gate 18° 26° 1.0799 0.5352 0.0734 1.6885 64.0% 31.7% 0.227 Crossing the gate As shown in the table, within the range of fuselage angle of attack increasing from 8° to 12°, the wing slope ratio decreased from 0.801 to 0.227, indicating that the wing had entered a stall. However, the wing's lift coefficient remained around 1.08, approximately on the order of the upper limit of usable lift coefficient on the wing side, and did not return to zero. The lift already gained was retained after the stall, and the overall lift coefficient still increased from 1.3549 to 1.6885. The lift coefficient of the central lifting body increased linearly with the fuselage angle of attack, without any inflection point, and its proportion increased from 19.4% to 31.7%. Therefore, wing stall does not equate to lift loss: after wing stall, if the fuselage angle of attack continues to increase, the overall lift continues to grow, but the main source of this growth has shifted from the wing to the central lifting body. Furthermore, the data in the table above ends at a fuselage angle of attack of 18°. The upper limit of the angle of attack design in this embodiment is based on this fuselage angle of attack of 18°. Lift data for fuselage angles of attack above 18° are pending. This specification does not cite any lift coefficient values ​​for fuselage angles of attack above 18°. In addition, the above-mentioned grouped lift and load-bearing ratios are given according to the steady-state solution method. After verification by unsteady-state solution under the same angle of attack condition, the lift coefficient given by the steady-state solution in the stall region is too high, and the aerodynamic load-bearing ratio calculated from it is too optimistic. The relative trend in this state and the aforementioned criterion relationship are not affected by this deviation.

[0031] As shown in Figure 2, the marginal allocation of the lift increase after overtaking the gate is shown in the table below, that is, the proportion of each item in the total lift increase of the aircraft for each 1° increase in fuselage angle of attack: 0°~4° 61.1% 32.0% 6.9% 8°~10° 53.8% 36.6% 9.6% 12°~14° 43.1% 47.3% 9.6% 16°~18° 31.8% 59.0% 9.2% As shown in the table, starting from a fuselage angle of attack of 12°, the primary source of lift increase changes from the wing to the central lifting body: in the 12°~14° range, the marginal share of the central lifting body is 47.3%, and that of the wing is 43.1%; in the 16°~18° range, the central lifting body is 59.0%, and that of the wing is 31.8%. Based on this, this embodiment sets the target value for the fuselage angle of attack within the following angle-of-attack range: within this range, the share of the increased overall lift from each unit increase in fuselage angle of attack is greater for the central lifting body than for the wing, i.e., fuselage angle of attack 12°~18°; this overshoot point is marked in Figure 2.

[0032] As shown in Figure 3, the load capacity calculation based on the design values ​​in this embodiment is shown in the following table: 11.9 m / s 5° (standard balancing) 48.6 N 9.9% 90.1% Approximately 0.313 11.9 m / s 18° (Operating point in this state) 93.0 N 19.0% 81.0% Approximately 0.382 18.0 m / s 6° 120.5 N 24.6% 75.4% Approximately 0.424 18.0 m / s 18° 212.8 N 43.4% 56.6% Approximately 0.568 24.3 m / s 18° 387.8 N 79.1% 20.9% Approximately 0.840 27.3 m / s 18° 489.4 N 99.8% 0.2% Approximately 0.999 The table shows the direction of the load-bearing share change brought about by this state: at the load-bearing line where the rotor's load-bearing share is 90%, i.e., at an airspeed of 11.9 m / s and a fuselage angle of attack of 5° for conventional trim, the weight share borne by aerodynamic lift is approximately 9.9%. After increasing the fuselage angle of attack to the operating point of 18° in this state, the weight share borne by aerodynamic lift rises to approximately 19.0%, while the vertical thrust share borne by the rotor system decreases from approximately 90.1% to approximately 81.0%. For example, the vertical thrust required by the rotor can be reduced by approximately 44.4 N (design value). That is, in this state, the weight share borne by aerodynamic lift increases, while the vertical thrust share borne by the rotor system decreases, in the opposite direction to the conventional route of treating rotor thrust as the sole load-bearing source. This benefit peaks at the fuselage angle of attack of 18°, which is the upper limit of the angle of attack design in this embodiment; the benefits at higher fuselage angles of attack need to be supplemented. When the airspeed increases to 27.3 m / s, the aerodynamic lift accounts for approximately 99.8% of the weight. The aerodynamic lift is sufficient to support the entire aircraft in terms of weight. The rotor is reduced to a low thrust trim state, and its vertical thrust only needs to support drag, trim, and control requirements.

[0033] As shown in Figure 4, the available vertical thrust of the rotor is calculated based on the power budget of the rotor system and the state of the rotor disks. In this embodiment, the following parameters are used based on design: The available vertical thrust of all rotor disks is calculated using momentum theory and hover power as the anchor point, under a given electric power budget. When the electric power budget is 1.0, 1.5, and 2.0 times the hover power, the available vertical thrust of all rotor disks is estimated to be 1.00, 1.31, and 1.59 times the total weight, respectively. In this embodiment, 1.5 times the hover power budget is used as the available parameter. This parameter is a power-limited parameter and not based on actual rotor bench measurements; therefore, this specification does not cite absolute thrust and power values ​​for the rotor or motor. The vertically available thrust and the required vertical thrust of the rotor are synthesized by decomposing them into vertical components according to the ground thrust direction of each rotor disk of the forward propulsion rotor group and the lift rotor group. In this embodiment, the vertical component coefficient of the forward propulsion rotor group is taken as 0.8925 and the vertical component coefficient of the lift rotor group is taken as 0.9063, both of which are determined by the ground thrust direction of each rotor disk. The actual thrust of each rotor can be read back from at least one of the following: the telemetry readback of the rotor system's electronic speed controller, current sensor, voltage sensor, speed sensor, and thrust sensor, and reconciled with the vertically available thrust of the aforementioned power-limited caliber.

[0034] As shown in Figure 4, the rotor lift margin is the ratio of the difference between the rotor's available vertical thrust and the rotor's required vertical thrust to the rotor's available vertical thrust. The required vertical thrust is determined by the difference between the total weight of the aircraft and the current aerodynamic lift. Based on the available caliber of the above 1.5 times hovering power budget and the load calculation in the table above, the relative levels of the rotor lift margin in this embodiment are: approximately 0.237 in hovering state, approximately 0.382 at the low-speed operating point of controlled stall (airspeed 11.9 m / s, fuselage angle of attack 18°), approximately 0.568 at airspeed 18.0 m / s, and approximately 0.999 at the aerodynamic full load point of airspeed 27.3 m / s. That is, the rotor lift margin in controlled stall state is higher than that in hovering state and increases monotonically with increasing airspeed. The lower limit of the rotor lift margin is determined based on the upper limit of the rotor load share: in this embodiment, the upper limit of the rotor load share is taken as 90%, and the corresponding lower limit of the rotor lift margin is approximately 0.313; the entry and exit of the controlled stall state are determined according to this lower limit and its hysteresis. The above margin values ​​are taken from the vertical available thrust of the power-limited caliber, and are design estimates rather than actual bench measurements.

[0035] As shown in Figure 4, in the controlled stall state, the rotor lift margin serves as the sole envelope protection criterion and the sole exit criterion for this state. During this state, the aircraft sets a modal declaration flag, which prevents wing angle-of-attack stall alarms, gate-crossing degradation links, and intervention boundaries calibrated according to the wing stall gate from participating in the calculation of the envelope protection and exit criteria, and prohibits rearmapping of the above items during this state. During this state, only one envelope protection chain with rotor lift margin as input is retained, and the output of this envelope protection chain is the command to exit the controlled stall state. The aircraft does not use the fuselage angle of attack or the wing angle of attack as the envelope protection criterion, and does not set intervention boundaries calibrated according to the wing stall gate or stall speed. If the fuselage angle of attack is greater than the wing stall gate, it does not constitute an abnormal event, and neither stall alarms nor gate-crossing degradation are triggered. In Figure 4, the wing angle-of-attack stall warning and gate-crossing degradation link are drawn outside the envelope protection chain with a dashed box, indicating that they have been removed from the envelope protection. The conditions for exiting the controlled stall state include: the rotor lift margin is lower than its lower limit; the fuselage angle of attack reaches the upper limit of the following angle of attack; the airspeed enters the aerodynamic full load zone; the vertical thrust share borne by the rotor system is lower than a preset share. In this embodiment, the preset share is 1%, corresponding to an airspeed of approximately 27.3 m / s and an aerodynamic load share of approximately 99.8%; and receiving a modal exit command from the mission layer or the pilot.

[0036] As shown in Figure 4, the inverse solution for obtaining the target value of the fuselage angle of attack from the lower limit of the rotor lift margin is performed as follows: Taking the aerodynamic lift corresponding to the lower limit of the rotor lift margin as the inverse solution target, and based on the monotonic relationship between aerodynamic lift and fuselage angle of attack at the current airspeed, the minimum fuselage angle of attack that ensures the aerodynamic lift is not less than the inverse solution target is obtained, and this minimum fuselage angle of attack is taken as the target value of the fuselage angle of attack. This relationship is given by the aforementioned grouped lift accounts and the overall aircraft polar curve: Below the wing stall gate, the aerodynamic lift increases with the fuselage angle of attack, contributed by both the wing and the central lifting body, resulting in a relatively large slope; above the wing stall gate, the wing lift is saturated at its upper limit of usable lift coefficient and no longer increases with the angle of attack. The increase in aerodynamic lift is entirely contributed by the central lifting body and the horizontal stabilizer, resulting in a correspondingly lower slope. However, this relationship remains monotonic, so the inverse solution remains unique above the threshold. The inverse kinematics is processed in segments below and above the wing stall threshold: First, the angle of attack corresponding to the aerodynamic lift that satisfies the lower limit of the margin is calculated based on the current airspeed and this relationship. If this angle of attack is above the wing stall threshold, it is taken as the target value of the fuselage angle of attack. If the minimum fuselage angle of attack obtained by the monotonic relationship is below the threshold, then the minimum angle of attack is taken, meaning that the lower limit of the margin can be satisfied without exceeding the threshold. The target value of the fuselage angle of attack is given by the inverse kinematics of the lower limit of the rotor lift margin, rather than by a preset modal angle of attack or airspeed schedule. The inverse kinematics result of this embodiment falls within the range of 12° to 18° of fuselage angle of attack. Within this range, the proportion of the increase in overall lift per unit increase in fuselage angle of attack is greater for the central lifting body than for the wing. The target value of the fuselage angle of attack increases the weight share borne by aerodynamic lift to above the preset share, and reduces the vertical thrust share borne by the rotor system to below the share corresponding to the lower limit of the rotor lift margin. In this state, the rotor system is in a low thrust trim state, and the vertical thrust of the rotor system only bears the vertical component corresponding to the difference between the total weight of the aircraft and the current aerodynamic lift, as well as the trim and control requirements of the aircraft.

[0037] In this embodiment, the upper limit of the angle of attack is jointly calibrated based on the effective domain of the aerodynamic characteristics of the central lifting body and the pitch trim capability of the aircraft: the effective domain of the grouped lift data of the central lifting body ends at a fuselage angle of attack of 18°. There is currently no aerodynamic basis for higher fuselage angles of attack, and higher fuselage angles of attack would exceed the pitch trim capability. Therefore, the upper limit of the angle of attack is set at 18°. The working angle of attack in the controlled stall state does not exceed this upper limit. Data for fuselage angles of attack above 18° are pending. Once the data is available, this upper limit can be relaxed accordingly. After exiting the controlled stall state, the aircraft recovers to the trim angle of attack calculated based on the current airspeed and total weight using conventional trim, and is trimmed and controlled via conventional channels.

[0038] As shown in Figure 5, under controlled stall conditions, a benefit availability criterion is also executed: the proportion of the current aerodynamic lift, after deducting the lift contributed by the central lifting body, to the total aircraft weight is calculated online to obtain the aerodynamic load share after removing the central lifting body, and the rotor load share required for the rotor system to supplement this aerodynamic load share is calculated; when the rotor load share exceeds the preset share, the benefits of the controlled stall state are determined to be unavailable and the state is exited. The elimination calculation given in this embodiment according to the design values ​​is shown in the table below, with the fuselage angle of attack in each row taken as 18°: 11.9 m / s 19.0% 13.0% 18.1% 6.0 percentage points 0.8 percentage points 18.0 m / s 43.4% 29.6% 41.5% 13.8 percentage points 1.9 percentage points 24.3 m / s 79.1% 54.0% 75.6% 25.1 percentage points 3.4 percentage points As shown in the table, the central lifting body is the main source of aerodynamic load in this state: after removing the central lifting body, the rotor system must compensate with 6.0 to 25.1 percentage points of the overall weight at the corresponding operating point; the preset share in this embodiment is 90%, which is the same as the lower limit of the rotor lift margin. At the operating point of airspeed 11.9 m / s, the rotor load share after removing the central lifting body is 87.0%, which is close to the preset share, and the benefit is close to the boundary; when the airspeed increases to 18.0 m / s and 24.3 m / s, the values ​​are 70.4% and 46.0% respectively, and the benefit is sufficient. If the central lifting body stalls along with the rotor, the benefit in this state becomes zero, and this criterion can be used to report the unusable benefit online and exit this state.

[0039] The forward thrust rotor assembly and the lift rotor assembly share the vertical thrust required by the rotors according to the thrust distribution that makes the overall pitch moment of the aircraft zero. This thrust distribution ratio is determined by the ratio of the pitch gain of the rotor disk center of the forward thrust rotor assembly and the lift rotor assembly relative to the overall center of gravity of the aircraft. In this embodiment, the design values ​​are: the pitch gain of the forward thrust rotor assembly is +0.1844 N·m / N, and the pitch gain of the lift rotor assembly is −0.3171 N·m / N. The thrust distribution ratio obtained by the ratio of the two is T_forward / T_lift = 1.7193. It should be noted that this distribution ratio is calibrated according to the condition of making the overall pitch moment zero. It belongs to two different design calibers than the equivalent caliber obtained by canceling the horizontal thrust components of the two rotors alone. This embodiment is calibrated according to the former. The vertical force balance relationship of the aircraft is the same under both calibers. The difference lies in the attribution of the pitch moment and the design point of the rotor disk relative to the ground angle determined accordingly.

[0040] In the controlled stall state, the roll channel is undertaken by the equal and opposite thrust differential of the lifting rotors symmetrically arranged on the left and right sides of the rotor system. In this embodiment, the deflection authority of the aileron is less than the roll authority of the equal and opposite thrust differential. The roll channel does not switch when entering and exiting the controlled stall state, that is, the channel does not switch laterally in this state, and the existing roll channel is still used laterally.

[0041] It should be noted that the data boundaries in this embodiment are as follows: the aerodynamic lift and load-bearing ratios mentioned above are given under symmetrical steady-state conditions. There is currently no data on the asymmetric stall condition where one wing enters stall before the other. This segment is controlled laterally by the equal and opposite thrust differential of the lifting rotor, and will be supplemented after lateral test or identification data is backfilled. In addition, the state where the fuselage angle of attack is greater than the wing stall gate after the wing stalls does not constitute an abnormal event in itself, does not trigger a stall alarm, does not trigger a gate-override degradation, and does not trigger any interventions calibrated according to the wing stall gate.

[0042] The aerodynamic lift figures in the tables above are all taken from models excluding rotor slipstream, i.e., the force amplification of the rotor slipstream on the wing and central lifting body is not considered. The aerodynamic load-bearing ratios given under this caliber are conservative values; after considering the rotor slipstream, the aerodynamic load-bearing ratio will only be higher, not lower. Regarding power consumption, using hovering power consumption as a baseline of 1.000, the relative power consumption at the low-speed operating point (airspeed 11.9 m / s, fuselage angle of attack 18°) in this state is estimated to be approximately 0.740 based on design, and the relative power consumption at the aerodynamic full-load point (airspeed 27.3 m / s) is approximately 0.135~0.194. The above power consumption estimates are based on momentum theory and anchored to hovering power; they are comparative calibers within the same family, not actual bench measurements, and are only used to illustrate ratios and trends.

[0043] The main parameters for this embodiment, based on the design values, are shown in the table below: Total takeoff weight 490.3 N Prototype level Wing reference area / Mean aerodynamic chord / Wingspan / Length 0.6350 m² / 0.3203 m / 2.1934 m / 1.0962 m Design value Wing installation angle 8° relative to the fuselage Wing stall angle of attack (wing diameter) Approximately 17° Wing chord angle of attack Wing stall door (fuselage angle of attack type) 9° When the fuselage angle of attack exceeds 9°, the wing enters a stall. Upper limit of available lift coefficient on the wing side 1.10~1.15 The corresponding overall lift coefficient is approximately 1.35~1.45. Preset threshold for low speed range Approximately 27.3 m / s Airspeed calibration based on the aerodynamic lift just enough to support the weight of the entire aircraft. Operating point of fuselage angle of attack under controlled stall conditions 12°~18° Obtained by inverse solution of the lower limit of the residual Upper limit of angle of attack 18° Central lifting body aerodynamic characteristics effective range and pitch trim capability Available caliber for power budget 1.5 times hovering power The available vertical thrust is approximately 1.31 times the total weight of the machine. Rotor lift margin: hovering / low-speed operating point Approximately 0.237 / Approximately 0.382 It increases monotonically with airspeed to approximately 0.999 upper limit of rotor load share / lower limit of margin 90% / approximately 0.313 The lower limit of the margin is set according to the upper limit of the bearing share. Vertical component coefficient of forward propulsion rotor assembly / lift rotor assembly 0.8925／0.9063 Determined by the direction of the ground thrust of each propeller disk. Forward propulsion rotor assembly / lift rotor assembly pitch gain +0.1844 / −0.3171 N·m / N The thrust ratio for zero pitch moment is 1.7193. The exit threshold of rotor vertical thrust fraction 1% Corresponding pneumatic full load area

[0044] Alternative implementation methods: The form of the above-described rotor system can be replaced. The lifting rotor assembly can be changed from two coaxial twin rotors to a single rotor assembly. The number of rotors can be increased or decreased. The rotors can be fixed-pitch rotors or variable-pitch rotors with pitch adjustment mechanisms. The number of rotor blades can be two, three, four, or more. The forward propulsion rotor assembly can also be located at a position other than the wingtip, or it can be combined with the lifting rotor to form a unified thrust unit. The power form can be changed from pure electric to hybrid electric. The entire aircraft can also be scaled up proportionally to larger-scale models such as logistics and transportation aircraft, while the angular relationships and criterion caliber remain unchanged. The equivalent expressions of the above-described criteria can be replaced. For example, the rotor lift margin can also be equivalently expressed as the ratio of the rotor's vertical thrust required to the rotor's vertical thrust available, or as the rotor's load-bearing share. The three correspond one-to-one, and any expression can serve as the unique envelope protection criterion and exit criterion for this state. The selection method for the aforementioned lower limit of the margin can be replaced. It can be calibrated according to the upper limit of the rotor load share, or according to the upper limit of the electric power budget, or according to the pitch trim margin, and the corresponding lower limit value will change accordingly. The aforementioned exit conditions can be used in combination. The exit condition can be based solely on the rotor lift margin being lower than the lower limit, or it can be combined with any one or more of the following: fuselage angle of attack reaching the upper limit of angle of attack, airspeed entering the aerodynamic full load zone, rotor vertical thrust share being lower than the preset share, and mode exit command. The target value of the aforementioned fuselage angle of attack can also be solved by inverse solving the rotor load share corresponding to the lower limit of the rotor lift margin, or by numerically iterating to approximate the lower limit. The piecewise processing of the inverse solution can also be simplified to a single-piece solution on the same monotonic relation. The aforementioned central lifting body configuration is replaceable and can be composed of a blended fuselage, an independent lifting body section, or a combination of fuselage and nacelle. The stall gates of the central lifting body and wing can also be calibrated according to other angles of attack, as long as the relationship remains that when the fuselage angle of attack is greater than the wing stall gate, the wing lift saturates and the increasing lift demand is borne by the central lifting body. The aforementioned preset threshold for the low-speed range can also be calibrated according to the airspeed corresponding to the upper limit of the rotor load share, or according to the airspeed at which the wing alone bears the weight of the entire aircraft. The aforementioned upper limit of angle of attack can also be appropriately relaxed after the aerodynamic data of the central lifting body is backfilled. The aforementioned roll channel execution mode is replaceable and can be equal-reverse thrust differential of symmetrically arranged lifting rotors, differential thrust of symmetrically arranged thrust units, or other equivalent lateral execution modes. In this state, no new channel handover is required. The preset share, lower limit of margin, and hysteresis in the aforementioned criteria can be scaled accordingly based on the aircraft type, electrical power budget, and flight control bandwidth. All the above variants can achieve controlled stall lift distribution functions, using the rotor lift margin as the sole envelope protection criterion and exit criterion, resolving the target value of the fuselage angle of attack from the lower limit of this margin, and ensuring that an angle of attack greater than the wing stall gate does not constitute an abnormal event. The solution of this invention is not limited to vertical takeoff and landing aircraft, but can also be applied to aircraft with runway takeoff and landing or short takeoff and landing.

[0045] In another embodiment, the aircraft is a runway takeoff and landing aircraft equipped with wheeled landing gear: during the runway takeoff and landing phase, the aircraft's wheels touch down and bear the vertical load corresponding to the difference between the total weight of the aircraft and the current aerodynamic lift. The wings, the central lifting body, and the rotor system share the lift requirements according to the aforementioned lift distribution relationship. The fuselage angle of attack is determined by the wheel load and the nose wheel lift speed during the runway takeoff and landing phase. The rotor lift margin is still used as the envelope protection criterion and exit criterion for the runway takeoff and landing phase. The aforementioned entry conditions for controlled stall, the inverse solution method for the angle of attack target, and the exit conditions are also applicable to runway takeoff and landing aircraft.

[0046] The specific embodiments of the present invention have been described above with reference to the accompanying drawings. However, these descriptions should not be construed as limiting the scope of the present invention. The scope of protection of the present invention is defined by the appended claims. Any modifications based on the claims of the present invention are within the scope of protection of the present invention.

Claims

1. A controlled stall lift distribution control method for an aircraft, the aircraft comprising a fuselage, wings, and a rotor system, the fuselage constituting the central lifting body of the aircraft, the difference between the stall angle of attack of the wing and the angle of installation of the wing relative to the fuselage constituting the stall gate of the wing, the stall gate being a threshold value in the form of the fuselage angle of attack; the aircraft is in a controlled stall state at low speeds below a preset threshold, in which the fuselage angle of attack is greater than the wing stall gate, the lift of the wing is saturated at the upper limit of the available lift coefficient of the wing, and any further increase in lift demand is borne by the central lifting body; characterized in that, Includes the following steps: Step 1: Real-time acquisition of the airspeed, fuselage angle of attack and total weight of the aircraft; calculation of the current aerodynamic lift of the aircraft based on the airspeed and fuselage angle of attack; and calculation of the available vertical thrust of the rotor system based on the power budget and rotor disk status of the rotor system. Step 2: Calculate the rotor lift margin. The rotor lift margin is the ratio of the difference between the rotor's vertically available thrust and the rotor's vertically required thrust to the rotor's vertically available thrust. The rotor's vertically required thrust is determined by the difference between the overall weight of the aircraft and the current aerodynamic lift. The rotor lift margin is used as the sole envelope protection criterion and the sole exit criterion for the controlled stall state. The aircraft does not use the fuselage angle of attack or the wing angle of attack as the envelope protection criterion, and does not set an intervention boundary based on the wing stall gate or the stall speed. The criterion chain composed of wing angle of attack stall alarms and gate-crossing degradation links is removed from the envelope protection of the controlled stall state and replaced by the rotor lift margin. The fuselage angle of attack being greater than the wing stall gate does not constitute an abnormal event. Step 3: The target value of the fuselage angle of attack is obtained by inversely solving the relationship between the lower limit of the rotor lift margin and the monotonic change of aerodynamic lift with the fuselage angle of attack. The target value of the fuselage angle of attack is given by inversely solving the lower limit of the rotor lift margin, rather than by a preset modal angle of attack or airspeed plan table. Step 4: Control the pitch of the aircraft according to the target value of the fuselage angle of attack, so that the fuselage angle of attack tracks the target value of the fuselage angle of attack; and exit the controlled stall state when the rotor lift margin is lower than the lower limit.

2. The controlled stall lift distribution control method according to claim 1, characterized in that: In step 3, the inverse solution takes the aerodynamic lift corresponding to the lower limit of the rotor lift margin as the inverse solution target. According to the monotonic relationship between the aerodynamic lift and the fuselage angle of attack at the current airspeed, the minimum fuselage angle of attack that makes the aerodynamic lift not less than the inverse solution target is obtained, and the minimum fuselage angle of attack is used as the target value of the fuselage angle of attack.

3. The controlled stall lift distribution control method according to claim 1, characterized in that: In step 2, the wing angle-of-attack stall alarm and the gate-crossing degradation link are removed from the envelope protection of the controlled stall state. Specifically, during the controlled stall state, the aircraft sets a modal declaration flag, which prevents the wing angle-of-attack stall alarm, the gate-crossing degradation link, and the intervention boundary calibrated according to the wing stall gate from participating in the calculation of the envelope protection criteria and exit criteria, and prohibits the above items from rearming during the controlled stall state. During the controlled stall state, only one envelope protection chain with the rotor lift margin as input is retained, and the output of the envelope protection chain is the command to exit the controlled stall state. The entry and exit of the controlled stall state are determined according to the lower limit of the rotor lift margin and its hysteresis.

4. The controlled stall lift distribution control method according to claim 1, characterized in that, It also includes a benefit availability criterion: under the controlled stall state, the proportion of the aerodynamic lift after deducting the lift contributed by the central lifting body to the total weight of the aircraft is calculated online to obtain the aerodynamic load share after removing the central lifting body, and the rotor load share required by the rotor system to make up the aerodynamic load share is calculated; when the rotor load share exceeds the preset share, it is determined that the benefit of the controlled stall state is unavailable and the controlled stall state is exited.

5. The controlled stall lift distribution control method according to claim 1, characterized in that: In step 3, the inverse solution also makes the target value of the fuselage angle of attack fall into the following angle of attack range, in which the proportion of the central lifting body in the total lift of the aircraft for each unit increase in fuselage angle of attack is greater than the proportion of the wing.

6. The controlled stall lift distribution control method according to claim 1, characterized in that: In step 4, the conditions for exiting the controlled stall state also include the fuselage angle of attack reaching the upper limit of the angle of attack calibrated according to the effective domain of the aerodynamic characteristics of the central lifting body and the pitch trim capability of the aircraft; When the airspeed enters the aerodynamic full load zone and the vertical thrust share borne by the rotor system is lower than a preset share, the controlled stall state is exited; and when a modal exit command is received from the mission layer or the pilot, the controlled stall state is exited.

7. The controlled stall lift distribution control method according to claim 1, characterized in that: The rotor system includes a forward propulsion rotor assembly and a lift rotor assembly. In steps 1 and 2, the available vertical thrust of the rotor and the required vertical thrust of the rotor are decomposed into vertical components according to the ground thrust direction of each rotor disk of the forward propulsion rotor assembly and the lift rotor assembly, and then synthesized. The forward propulsion rotor assembly and the lift rotor assembly share the required vertical thrust of the rotor according to the thrust ratio that makes the pitch moment of the entire aircraft zero. The thrust ratio is determined by the ratio of the pitch gain of the rotor disk center of the forward propulsion rotor assembly and the lift rotor assembly relative to the center of gravity of the entire aircraft.

8. The controlled stall lift distribution control method according to claim 1, characterized in that: In the controlled stall state, the roll channel is undertaken by the equal and opposite thrust differential of the lifting rotors symmetrically arranged on the left and right sides of the rotor system. The aileron deflection authority of the aircraft is less than the roll authority of the equal and opposite thrust differential. The roll channel does not switch when entering and exiting the controlled stall state.

9. The controlled stall lift distribution control method according to claim 1, characterized in that: In steps 3 and 4, the target value of the fuselage angle of attack increases the weight share borne by the aerodynamic lift to above a preset share, and reduces the vertical thrust share borne by the rotor system to below the share corresponding to the lower limit of the rotor lift margin; in the controlled stall state, the rotor system is in a low thrust trim state, and the vertical thrust of the rotor system only bears the vertical component corresponding to the difference between the total weight of the aircraft and the current aerodynamic lift, as well as the trim and control requirements of the aircraft.

10. A controlled stall lift distribution control system for an aircraft, the aircraft comprising a fuselage, a wing, and a rotor system, the fuselage constituting the central lifting body of the aircraft, the difference between the stall angle of attack of the wing and the angle of installation of the wing relative to the fuselage constituting the stall gate of the wing, the stall gate being a threshold value in the form of the fuselage angle of attack, characterized in that, include: Flight control computer; The attitude inertial measurement unit, airspeed sensor, and angle of attack sensor are respectively connected to the signal input terminal of the flight control computer; A thrust feedback device is used to read back the actual thrust of each rotor in the rotor system. The thrust feedback device is at least one of the following: telemetry feedback of the electronic speed controller of the rotor system, current sensor, voltage sensor, speed sensor and thrust sensor. The signal output terminal of the thrust feedback device is connected to the signal input terminal of the flight control computer. Each rotor of the rotor system is driven by an electronic speed controller, and the control input terminal of each electronic speed controller is connected to the control output terminal of the flight control computer. The flight control computer is configured to perform the controlled stall lift distribution control method according to any one of claims 1 to 9.

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