CN107102547A - A kind of RLV landing phase Guidance Law acquisition methods based on sliding mode control theory - Google Patents

Abstract

A method for obtaining the guidance law of the RLV landing segment based on sliding mode control theory. Firstly, the altitude deviation and the lateral distance deviation are calculated according to the nominal trajectory of the RLV landing segment; then, an appropriate sliding mode surface is selected according to the tracking deviation of the nominal trajectory, and The RLV three-degree-of-freedom dynamic equation is equivalently transformed into a form described by a sliding surface; finally, the guidance law is designed using the Lyapunov method. In the design process, the upper bound of the uncertainty of the guidance loop is obtained by analyzing the aerodynamic data, and a compensation item is introduced to suppress it, so that the guidance system has asymptotic stability to uncertainties such as disturbances. The method of the invention can effectively overcome the influence of uncertainty on the RLV guidance system, thereby improving guidance precision; by introducing a sliding mode control method, the original second-order system is equivalently transformed into a first-order system, making the design process more intuitive.

Description

A Method for Acquiring Guidance Law of RLV Landing Phase Based on Sliding Mode Control Theory

technical field

The invention relates to a method for obtaining guidance law of RLV landing section based on sliding mode control theory.

Background technique

Reusable launch vehicles (RLV) are space shuttle vehicles that combine the features and functions of spacecraft and aircraft. They can stay in orbit to complete various space tasks, and can also return safely and accurately like an airplane. ground. Due to its reusable characteristics, RLV will become a highly reliable carrier for human beings to explore the universe at a low cost and a military weapon for competing for space supremacy. Therefore, the major powers in the world continue to invest huge resources in its research and development, and conduct new research and exploration.

The return and re-entry phase of the RLV is usually divided into the initial re-entry phase, the terminal energy management phase, and the approach and landing phase. Go-around capability, if the guidance or control method is unstable or cannot meet the accuracy requirements, it may cause the RLV to fail to land safely, and even lead to catastrophic consequences. During the landing process, the uncertainty of aerodynamic data, atmospheric density, and external disturbances such as wind all affect the flight of the RLV, so the guidance law used must be robust to these uncertainties or disturbances, Thereby improving the success rate of landing. In modern nonlinear control methods, sliding mode control theory can reduce the order of high-order systems. The control law design process is simple, intuitive, and also has strong robustness. It can provide new ideas for the design of guidance laws for RLV landing stages. .

Contents of the invention

The technical problem to be solved by the present invention is: to overcome the deficiencies of the prior art, a method for obtaining the guidance law of the RLV landing section is proposed, which fully considers the uncertainty and disturbance of the aircraft itself and the outside during the landing process, and utilizes the sliding mode control theory The guidance law is designed, and the disturbance and uncertainty compensation items are introduced in the design process to make the guidance law robust. According to the Lyapunov method and the exponential function convergence speed and range requirements corresponding to the sliding mode surface, the control parameters are determined, so that The tracking error of the landing nominal trajectory has asymptotic convergence.

The technical solution adopted in the present invention is: a method for obtaining the guidance law of the RLV landing section based on sliding mode control theory, comprising the following steps:

1) Calculate and obtain the height deviation and lateral deviation of the RLV according to the obtained current height of the RLV, the lateral distance of the RLV from the airport runway and the preset nominal trajectory of the RLV landing;

2) According to the RLV landing nominal trajectory and the RLV particle kinematics equation, establish the landing nominal trajectory tracking error differential equation;

3) Determine the sliding surface of the height channel and the side channel;

4) Select N feature points on the nominal trajectory of the RLV landing, calculate the nominal lift at each feature point according to the aircraft’s nominal aerodynamic coefficient and nominal flight state, and calculate the lift uncertainty of each feature point separately and uncertainty upper bounds;

5) According to the RLV particle dynamics equation, the differential equation of the landing nominal trajectory tracking error in step 2), and the sliding mode surface selected in step 3), the sliding mode surface s ₁ for the height channel and the sliding mode surface for the lateral channel are obtained. Differential equation of die surface s2 _;

6) Establish a virtual guidance law;

7) Utilize the saturation function to replace the sgn function in the virtual guidance law in step 6);

8) Calculate and obtain the expected lift and the expected roll angle, and obtain the expected angle of attack according to the expected lift, nominal aerodynamic data and current flight state de-interpolation;

9) Use the expected angle of attack and expected roll angle obtained in step 8) as the final guidance law to realize the tracking of the nominal landing trajectory by the RLV.

The landing nominal trajectory tracking error differential equation established in the step 2) is: Among them, v is the speed of RLV, γ is the track inclination angle of RLV, χ is the direction angle of RLV; h represents the current height of RLV, z represents the lateral distance of RLV from the airport runway, h _c represents the preset RLV landing mark called trajectory, Indicates the height deviation of the RLV.

The sliding mode surfaces of the height channel and the side channel in the step 3) are respectively with Among them, c ₁ and c ₂ are design parameters to be determined.

The selection rules of the design parameters c ₁ and c ₂ are: such that The convergence rate to converge to zero satisfies and z converge to within 1 meter within 40 seconds; and c ₁ >0, c ₂ >0.

The specific process of step 4) is: select N feature points on the nominal trajectory of the RLV landing, and calculate the lift uncertainty Δ ⁺ = | L ⁺ -L ₀ , Δ ^- = | L ^- of each feature point respectively -L ₀ |, and determine the upper bound of uncertainty Δ _M = (1+10%) Δ; where Δ is the maximum value of lift uncertainty Δ ⁺ , Δ ^- ; N is a positive integer; L ₀ is the characteristic The nominal lift corresponding to the point; L ⁺ is the maximum positive deviation of the aerodynamic data corresponding to the feature point and the lift in the case of positive wind disturbance; L ^- is the maximum negative deviation of the aerodynamic data corresponding to the feature point and there is a negative Lift in the presence of wind disturbances.

The differential equations about s ₁ and s ₂ in step 5) are

in, u ₁ (L,σ)=Lcosσ represents the longitudinal lift component, u ₂ (L,σ)=Lsinσ represents the lateral lift component, g is the acceleration of gravity, L is the lift of RLV, σ is the roll angle of RLV, Δ _γ is Disturbance force caused by stroke in the longitudinal channel and uncertain item due to uncertainty of aerodynamic data, Δ _χ is interference force produced by stroke in the transverse channel and uncertainty item due to uncertainty of aerodynamic data, m is the mass of RLV .

The virtual guidance law established in step 6) is expressed as

Among them, k ₁ , k ₂ are the design parameters to be determined and k ₁ >0, k ₂ >0; k ₁ , k ₂ according to the Lyapunov function The rate of convergence to zero is determined.

The selection rule of the design parameters k ₁ and k ₂ is as follows: V(t) ^≤e-2Kt V(0), K=min{k ₁ , k ₂ } converges to zero and the convergence speed satisfies V within 40 seconds Converge to within 1m ² /s ² .

Said step 7) in saturation function is δ is a positive number.

The specific expressions for calculating the desired lift L ^* and the desired roll angle σ ^* in the step 8) are:

The advantage of the present invention compared with prior art is:

(1) The method of the present invention introduces a first-order sliding mode surface for the height channel and the lateral channel respectively, which can guarantee that when the sliding mode surface converges to 0 in theory, the tracking error to the nominal trajectory can also converge to 0, so The sliding mode surface can be used to convert the second-order guidance dynamic system into a first-order equivalent. As long as the guidance law is designed so that the sliding mode surface converges to 0, the tracking of the nominal trajectory can be realized, making the design of the guidance law The process is simpler and more intuitive;

(2) The method of the present invention comprehensively analyzes the uncertainty of aerodynamic data and atmospheric density, and the wind disturbance that may exist, introduces the compensation item of uncertainty and disturbance, and combines with the sliding mode control method, so that the sliding mode surface can have Asymptotically converges to 0 in the case of uncertainty, which means that the nominal trajectory tracking error can also asymptotically converge to 0;

(3) Through the method for obtaining the guidance law proposed by the present invention, the guidance error (nominal trajectory tracking error) is first converted into a sliding surface, and finally it is converted into the convergence range of the Lyapunov function and the exponential function, and can be obtained according to The convergence speed and convergence range of the exponential function corresponding to the Lyapunov function and the sliding mode surface adjust the guidance coefficient to obtain a satisfactory guidance effect, which provides a basis for the selection of parameters and improves the guidance accuracy.

Description of drawings

Fig. 1 is a block flow diagram of the inventive method;

Fig. 2 is the height curve of RLV under the inventive method effect;

Fig. 3 is the lateral deviation curve of RLV under the effect of the method of the present invention;

Fig. 4 is the speed curve of RLV under the inventive method effect;

Fig. 5 is the track inclination curve of RLV under the effect of the method of the present invention;

Fig. 6 is the direction angle curve of RLV under the effect of the inventive method;

Fig. 7 obtains the command curve of the angle of attack guidance law for the inventive method;

Fig. 8 is the command curve of the roll angle guidance law obtained by the method of the present invention.

detailed description

The invention is based on the guidance idea of tracking the nominal trajectory of landing, and uses the sliding mode control method and the disturbance compensation idea to design the guidance law of the approach and landing section of the RLV. According to the second-order nonlinear guidance model of the RLV approach and landing phase, the guidance error is transformed into the sliding surface error by using the sliding surface, and the original second-order model is converted into the first-order form, and the disturbance compensation item is introduced to affect the aerodynamic data, atmospheric The uncertainties caused by external disturbances such as density and wind are compensated, so that the sliding surface can asymptotically converge, and a satisfactory convergence speed can also be obtained by adjusting the guidance gain according to Lyapunov theory.

As shown in Figure 1, it is a flow chart of the method of the present invention, based on the sliding mode control theory of the RLV landing section guidance law acquisition method, the specific steps are as follows:

Step 1: Establish a coordinate system for the approach and landing phase: take the projection of the starting point of approach and landing on the ground as the origin, the direction pointing to the end of the runway is the x-axis, the y-axis is perpendicular to the x-axis, and points to the sky, and the z-axis is in the same direction as the x and y axes. Right-handed. Assume that the coordinates of RLV in this coordinate system are (x, h, z);

Step 2, according to the designed RLV landing nominal trajectory hc= _f (x), and the current height h of the RLV fed back by the altimeter and GNSS (global satellite navigation system) and the lateral distance z of the RLV from the airport runway , respectively calculate the height deviation of RLV and lateral deviation The specific design method of the landing nominal trajectory can be found in the literature GHBarton and SGTragesser, Autolanding trajectory design for the X-34, AIAA-99-4161, 1999;

Step 3, according to the nominal landing trajectory designed in step 2, and the RLV particle kinematic equation shown in formula (1)

Establish the landing nominal trajectory tracking error differential equation shown in formula (2)

Where v is the velocity of the RLV, γ is the track inclination of the RLV, and χ is the direction angle of the RLV; both γ and χ are obtained by the feedback of the navigation system composed of INS+GNSS;

Step 4, select the sliding surface of the height channel and the lateral channel as

Among them, c ₁ , c ₂ are the design parameters to be determined, by adjusting the design parameters c ₁ , c ₂ so that The convergence rate to converge to zero satisfies and z converge within 1 meter within 40 seconds; c ₁ >0, c ₂ >0;

for In other words, when s→0, there is x→0 (theoretical basis can be found in the literature: Duan Guangren, Hou Mingzhe, Tan Feng. Design of adaptive integrated guidance and control law based on sliding mode method. Acta Armaments, 2010,31 (2):191-198.), so the problem of RLV tracking the nominal trajectory can be transformed into the problem of controlling the sliding surface s ₁ , s ₂ to converge to 0.

Step 5, select some feature points on the RLV landing nominal trajectory, and calculate the nominal lift L ₀ according to the nominal aerodynamic coefficient and the nominal flight state of the aircraft on each feature point;

Considering the maximum positive deviation of the aerodynamic data and the possible positive wind disturbance, the lift force L ⁺ is calculated again according to the nominal flight state at the selected feature points;

Considering the maximum negative deviation of aerodynamic data and the possible negative wind disturbance, the lift force L ^- is calculated again according to the nominal flight state at the selected feature points;

For each selected feature point, calculate the lift uncertainty Δ ⁺ ＝ |L ⁺ -L ₀ |, Δ ^- ＝ |L ^- -L ₀ |, and select the maximum value as Δ, and finally determine whether The upper bound of certainty is Δ _M = (1+10%) Δ;

Step 6, according to the RLV particle dynamics equation shown in formula (5)

and the nominal trajectory tracking error equation (2) in step 3, the differential equation of the sliding surface is obtained as

in, u ₁ (L,σ)=Lcosσ represents the longitudinal lift component, u ₂ (L,σ)=Lsinσ represents the lateral lift component, g is the acceleration of gravity, L is the lift of RLV, σ is the roll angle of RLV, m is the RLV Δ _γ is the disturbance force generated by the stroke in the longitudinal channel and the uncertainty item due to the uncertainty of the aerodynamic data, Δ _χ is the disturbance force generated by the stroke in the transverse channel and the uncertainty item due to the uncertainty of the aerodynamic data , satisfy Δ _γ ≤ Δ _M , Δ _χ ≤ Δ _M , and use this as a compensation item to offset the influence of uncertainty on RLV, and obtain a guidance law with disturbance suppression performance;

Step seven, in order to make the sliding surface s ₁ , s ₂ asymptotically converge to 0, select the Lyapunov function

Taking the derivative with respect to V gives

Step eight, in order to make the Lyapunov function V converge, design the control law according to (8)

Among them, k ₁ , k ₂ are the design parameters to be determined, k ₁ >0, k ₂ >0; k ₁ , k ₂ according to the Lyapunov function The convergence speed to converge to zero is determined; substituting (9) into (8) to get

As described in step 6, it can be seen that |Δ _γ |≤Δ _M , |Δ _χ |≤Δ _M , after substituting into (10), we can get

Among them, K=min{k ₁ ,k ₂ }, according to formula (11), it can be seen that the virtual control law of formula (9) can make system (6) asymptotically stable (for specific concepts, please refer to literature Khalil, HK, Nonlinear Systems, 3rd ed., Prentice-Hall, Upper Saddle River, NJ, 2002, Chapter 4), that is, s ₁ , s ₂ can asymptotically converge to zero point. At this time, it can be known from the properties of sliding mode s ₁ , s ₂ that Tracking error of RLV to nominal trajectory z can also asymptotically converge to 0.

It can be seen from formula (11) that increasing the design parameters k ₁ and k ₂ can increase the convergence speed of the system, so that the tracking error of the nominal landing trajectory can quickly converge to zero. Therefore, select c ₁ , c ₂ designed in the fourth step, so that and s can converge to within 1 meter within 40 seconds, and make V(t) ^{≤e -2Kt} V(0), K=min{k ₁ ,k ₂ } converge to zero by adjusting design parameters k ₁ , k ₂ The convergence speed meets the requirement that V converges to within ^{1 m2} /s2 within 40 seconds, and the next design step can be entered.

Step 9, in order to avoid the discontinuity of the sign function sgn, use the saturation function instead of the sgn function, namely

in, δ is a small positive number, usually 0.1;

Step ten, solve the expected lift L ^* and the expected roll angle σ ^* according to the obtained control law (12), namely

Afterwards, according to the expected lift L ^* , the expected angle of attack α ^* is obtained by using the nominal aerodynamic data and the current flight state inverse interpolation;

Step 11, the desired angle of attack α ^* and desired roll angle σ ^* obtained in step 10 are the final guidance law designed. After inputting it to the attitude control system, as long as it is effectively tracked, the RLV pair landing can be realized Tracking of the nominal trajectory.

Example

The effectiveness of the method of the present invention is illustrated below through simulation.

The trajectory of the RLV approach and landing segment is divided into steep glide segment, arc segment, exponential transition segment and shallow glide segment. The specific off-line trajectory design method can be found in the literature (G.H.Barton and S.G.Tragesser, Autolanding trajectorydesign for the X-34, AIAA -99-4161,1999.), this simulation example only gives the relevant parameters of the designed trajectory.

The coordinate system is established with the projection of the starting point of approach and landing on the ground as the origin, the x-axis points to the touchdown point, the y-axis is perpendicular to the x-axis and points to the sky, and the z-axis is determined according to the right-hand rule. The position of the aircraft in the coordinate system is defined by (x, h, z) said. Let the coordinates of the starting point of approach and landing be (0,3000,0) meters, the coordinates of the touchdown point be (13800,0,0) meters, the coordinates of the center of the arc segment be (13526,7015.5,0) meters, and the starting point of the arc segment be The coordinates of the starting point are (11626, 208.9, 0) meters, the coordinates of the starting point of the exponential transition section are (12873, 26.2, 0) meters, the decay rate of the exponential function is 264, the proportional coefficient of the exponential function is 10, and the track angle of the steep glide section is - 13.5°, and the shallow glide segment track angle is -1°.

Assuming that the uncertainty of aerodynamic data and the uncertainty caused by external disturbances such as wind are And take the guidance coefficients c ₁ =5, c ₂ =,1, Δ _M =2, k ₁ =1.5, k ₂ =1 in the steep glide section, k ₁ =1.5, k ₂ =1.5 after the beginning of the arc section, the lift satisfies L=QSC _L ≈QS(l _α α+l ₀ )=QSl _α α+QSl ₀ l _α ＝0.1,l ₀ ＝0.35, the resistance meets D=QSC _D ≈QS(d _α α+d ₀ )=QSd _α α+QSd ₀ , d _α =-0.02, d ₀ =-0.1, S=5.454, the acceleration of gravity is g=9.8m/s ² , the mass of the aircraft is m=3700kg, and the standard atmospheric density model is used. Consider the scenario: the initial position in the established coordinate system is (-500, 3200, 300) meters, the initial velocity is 156 m/s, the initial track inclination is -13°, the direction angle is -3°, the attack angle is 2°, and the initial values of other variables are all zero.

Figure 2 is the height curve, the abscissa is the horizontal distance x of RLV flight, the ordinate is the height h of RLV and the nominal height h _c , it can be seen that RLV can eliminate the initial height deviation before the height of 2500 meters, so that the actual height of RLV can be tracked The nominal trajectory of upper landing; Figure 3 is the lateral deviation curve, the abscissa is time, and the ordinate is the lateral distance z of the RLV. It can be seen that after about 40 seconds of flight, the lateral deviation can basically be kept near zero; Figure 4 is the speed curve, the abscissa is time, the ordinate is the speed v of RLV; Fig. 5 is the track inclination curve, the abscissa is time, the ordinate is the track inclination γ of RLV and the track inclination γ _c corresponding to the reference track, to eliminate Altitude deviation, the actual track of the RLV is slightly steeper than the track corresponding to the reference track under the action of the guidance law within 0 to 15 seconds. When the altitude deviation is eliminated, γ and the track inclination angle γ _c corresponding to the reference track basically overlap, thereby ensuring that the height of the RLV can track the reference trajectory; Figure 6 is the direction angle curve, the abscissa is time, and the ordinate is the direction angle χ of the RLV. Due to the existence of the lateral deviation at the initial moment, the direction angle χ of the RLV is at Under the action of the guidance law, adjustments are made to reduce the lateral deviation. When the lateral deviation tends to zero, χ is also maintained near zero; Figure 7 is the angle of attack curve, and α is the angle of attack of the RLV. It can be seen that at the initial moment of the simulation In order to eliminate the large change in the angle of attack of the height deviation, when the height deviation is basically eliminated, the angle of attack curve is relatively flat, until it enters the arc pull-up section (about 90 seconds), the angle of attack is quickly pulled up so that the actual height of the RLV can be tracked Landing nominal trajectory; Figure 8 is the roll angle curve, σ is the roll angle of the RLV, it can be seen that when there is a lateral position deviation, the roll angle changes drastically, and when the lateral deviation is eliminated, the roll angle is basically maintained near zero.

It can be seen from the simulation results that under the action of the guidance law acquisition method proposed in the present invention, the RLV can cope with a certain initial position deviation and uncertainty, and realize robust tracking of the landing nominal trajectory, and the resulting angle of attack , The sideslip angle guidance command is relatively smooth, and can be directly input into the attitude control system.

The content that is not described in detail in the description of the present invention belongs to the well-known technology of those skilled in the art.

Claims (10)

1. An RLV landing segment guidance law obtaining method based on a sliding mode control theory is characterized by comprising the following steps:

1) calculating and obtaining the height deviation and the lateral deviation of the RLV according to the obtained current height of the RLV, the lateral distance between the RLV and the airport runway and a preset RLV landing nominal track;

2) establishing a landing nominal track tracking error differential equation according to the RLV landing nominal track and an RLV mass point kinematic equation;

3) determining the sliding mode surfaces of the height channel and the lateral channel;

4) n characteristic points are selected on the RLV landing nominal track, nominal lift force is obtained on each characteristic point through calculation according to the nominal aerodynamic coefficient and the nominal flight state of the aircraft, and lift force uncertainty and uncertainty upper bound of each characteristic point are calculated respectively;

5) obtaining a sliding mode surface s related to the height channel according to an RLV particle dynamics equation, a landing nominal track tracking error differential equation in the step 2) and the sliding mode surface selected in the step 3)₁And the slip-form surface s of the lateral channel₂A differential equation of (2);

6) establishing a virtual guidance law;

7) replacing the sgn function in the virtual guidance law in the step 6) with a saturation function;

8) calculating to obtain an expected lift force and an expected roll angle, and obtaining an expected attack angle according to the expected lift force, the nominal aerodynamic data and the reverse interpolation of the current flight state;

9) and (3) taking the expected attack angle and the expected roll angle obtained in the step 8) as a final guidance law, and realizing the tracking of the RLV on the landing nominal track.

2. The RLV landing segment guidance law acquisition method based on the sliding mode control theory as claimed in claim 1, characterized in that: the landing nominal trajectory tracking error differential equation established in the step 2) is as follows:wherein v is the speed of the RLV, gamma is the track inclination angle of the RLV, and chi is the direction angle of the RLV; h represents the current height of the RLV, z represents the lateral distance of the RLV from the centerline of the airport runway, h_cRepresents a preset nominal trajectory for the RLV landing,indicating the height deviation of the RLV.

3. The RLV landing segment guidance law acquisition method based on the sliding-mode control theory as claimed in claim 2, characterized in thatIn the following steps: the slip form surfaces of the high-degree channel and the lateral channel in the step 3) are respectivelyAndwherein, c₁,c₂Are the design parameters to be determined.

4. The RLV landing segment guidance law acquisition method based on the sliding mode control theory as claimed in claim 3, characterized in that: the design parameter c₁,c₂The selection rule is as follows: so thatConvergence rate to zero point is satisfiedAnd z converges to within 1 meter in 40 seconds; and c is₁>0,c₂>0。

5. The RLV landing segment guidance law acquisition method based on the sliding mode control theory as claimed in claim 1, 2, 3 or 4, wherein: the specific process of the step 4) is as follows: selecting N characteristic points on the RLV landing nominal track, and respectively calculating the lift uncertainty delta of each characteristic point⁺＝|L⁺-L₀|、Δ^-＝|L^--L₀And determining an upper uncertainty bound Δ_MΔ (1+ 10%); where Δ is the lift uncertainty Δ⁺、Δ^-Maximum value of (1); n is a positive integer; l is₀A nominal lift corresponding to the feature points; l is⁺The maximum positive deviation of the pneumatic data and the lift force under the condition of positive wind disturbance are considered for the feature point; l is^-And considering the maximum negative deviation of the pneumatic data and the lift force under the condition of negative wind disturbance for the feature point.

6. The RLV landing segment guidance law acquisition method based on the sliding mode control theory as claimed in claim 5, characterized in that: in step 5) with respect to s₁,s₂Is a differential equation of

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Wherein,u₁(L, σ) ═ Lcos σ represents the longitudinal lift component, u₂(L, σ) ═ Lsin σ represents the lateral lift component,g is the gravitational acceleration, L is the lift of the RLV, σ is the roll angle of the RLV, Δ_γDisturbance forces for wind in longitudinal channels and uncertainty due to uncertainty in the aerodynamic data, Δ_χM is the mass of the RLV for the disturbance forces caused by the wind in the cross channel and for the uncertainty due to the uncertainty of the aerodynamic data.

7. The RLV landing segment guidance law acquisition method based on the sliding mode control theory as claimed in claim 6, characterized in that: the virtual guidance law established in the step 6) is expressed as

<mrow> <mfenced open = "{" close = ""> <mtable> <mtr> <mtd> <mrow> <msub> <mi>u</mi> <mn>1</mn> </msub> <mrow> <mo>(</mo> <mi>L</mi> <mo>,</mo> <mi>&amp;sigma;</mi> <mo>)</mo> </mrow> <mo>=</mo> <mi>L</mi> <mi> </mi> <mi>c</mi> <mi>o</mi> <mi>s</mi> <mi>&amp;sigma;</mi> <mo>=</mo> <mfrac> <mi>m</mi> <mrow> <mi>c</mi> <mi>o</mi> <mi>s</mi> <mi>&amp;gamma;</mi> </mrow> </mfrac> <mrow> <mo>(</mo> <mo>-</mo> <msub> <mi>k</mi> <mn>1</mn> </msub> <msub> <mi>s</mi> <mn>1</mn> </msub> <mo>-</mo> <msub> <mi>c</mi> <mn>1</mn> </msub> <mo>(</mo> <mrow> <mi>v</mi> <mi> </mi> <mi>s</mi> <mi>i</mi> <mi>n</mi> <mi>&amp;gamma;</mi> <mo>-</mo> <msub> <mover> <mi>h</mi> <mo>&amp;CenterDot;</mo> </mover> <mi>c</mi> </msub> </mrow> <mo>)</mo> <mo>-</mo> <msub> <mi>f</mi> <mi>h</mi> </msub> <mo>-</mo> <mi>sgn</mi> <mo>(</mo> <mrow> <msub> <mi>s</mi> <mn>1</mn> </msub> <mi>v</mi> <mi> </mi> <mi>c</mi> <mi>o</mi> <mi>s</mi> <mi>&amp;gamma;</mi> </mrow> <mo>)</mo> <msub> <mi>&amp;Delta;</mi> <mi>M</mi> </msub> <mo>)</mo> </mrow> </mrow> </mtd> </mtr> <mtr> <mtd> <mrow> <msub> <mi>u</mi> <mn>2</mn> </msub> <mrow> <mo>(</mo> <mi>L</mi> <mo>,</mo> <mi>&amp;sigma;</mi> <mo>)</mo> </mrow> <mo>=</mo> <mi>L</mi> <mi> </mi> <mi>s</mi> <mi>i</mi> <mi>n</mi> <mi>&amp;sigma;</mi> <mo>=</mo> <mfrac> <mi>m</mi> <mrow> <mi>c</mi> <mi>o</mi> <mi>s</mi> <mi>&amp;chi;</mi> </mrow> </mfrac> <mrow> <mo>(</mo> <mo>-</mo> <msub> <mi>k</mi> <mn>2</mn> </msub> <msub> <mi>s</mi> <mn>2</mn> </msub> <mo>-</mo> <msub> <mi>c</mi> <mn>2</mn> </msub> <mi>v</mi> <mi> </mi> <mi>cos</mi> <mi>&amp;gamma;</mi> <mi>sin</mi> <mi>&amp;chi;</mi> <mo>-</mo> <msub> <mi>f</mi> <mi>z</mi> </msub> <mo>-</mo> <mi>sgn</mi> <mo>(</mo> <mrow> <msub> <mi>s</mi> <mn>2</mn> </msub> <mi>v</mi> <mi> </mi> <mi>c</mi> <mi>o</mi> <mi>s</mi> <mi>&amp;gamma;</mi> <mi>c</mi> <mi>o</mi> <mi>s</mi> <mi>&amp;chi;</mi> </mrow> <mo>)</mo> <msub> <mi>&amp;Delta;</mi> <mi>M</mi> </msub> <mo>-</mo> <mi>sgn</mi> <mo>(</mo> <mrow> <msub> <mi>s</mi> <mn>2</mn> </msub> <mi>v</mi> <mi> </mi> <mi>s</mi> <mi>i</mi> <mi>n</mi> <mi>&amp;gamma;</mi> <mi>s</mi> <mi>i</mi> <mi>n</mi> <mi>&amp;chi;</mi> </mrow> <mo>)</mo> <msub> <mi>&amp;Delta;</mi> <mi>M</mi> </msub> <mo>)</mo> </mrow> </mrow> </mtd> </mtr> </mtable> </mfenced> <mo>,</mo> </mrow>

Wherein k is₁,k₂To be determinedK is a design parameter of₁>0,k₂>0；k₁,k₂According to the Lyapunov functionThe convergence rate to zero is determined.

8. The RLV landing segment guidance law acquisition method based on the sliding mode control theory as claimed in claim 7, characterized in that: the design parameter k₁,k₂The selection rule is as follows: so that V (t) is less than or equal to e^-2KtV(0),K＝min{k₁,k₂The convergence rate at which V converges to 1m in 40 seconds²/s²Within.

9. The RLV landing segment guidance law acquisition method based on the sliding-mode control theory as claimed in claim 1 or 7, wherein: the saturation function in the step 7) isIs a positive number.

10. The RLV landing segment guidance law acquisition method based on the sliding mode control theory as claimed in claim 8, characterized in that: calculating to obtain the desired lift L in the step 8)^*And desired roll angle σ^*The specific expression of (A) is as follows: