GB2616530A - Self-adaptive iterative guidance method and device for aerospace vehicle - Google Patents
Self-adaptive iterative guidance method and device for aerospace vehicle Download PDFInfo
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Abstract
A self-adaptive iterative guidance method and device for an aerospace vehicle. The guidance method comprises the following steps: updating an equivalent specific impulse of an engine in real time according to measurement information of an inertia measurement unit (S1); and adaptively adjusting parameters related to the working state of the engine in an iterative guidance algorithm by utilizing the equivalent specific impulse of the engine updated in real time, to obtain a flight program angle adapted to a thrust descent fault of the engine (S2). According to the guidance method and the guidance device, the adaptive capacity of the iterative guidance algorithm to the thrust descent fault of the engine can be improved, the guidance precision is improved, and the ability of the aerospace vehicle to complete a flight mission under the condition of the thrust descent fault of the engine is improved.
Description
Description
SELF-ADAPTIVE ITERATIVE GUIDANCE METHOD AND DEVICE FOR AEROSPACE VEHICLE
Technical Field
The present application belongs to the field of aerospace vehicle control technology, and in particular relates to a self-adaptive iterative guidance method and a guidance device for an aerospace vehicle.
Background Art
to The cost of development, production and launch of an aerospace vehicle, as a means of transportation into space, is very high. If the engine that powers the vehicle breaks down, it will easily lead to the failure of the launch mission, and the loss will be huge. For this reason, improving the ability of the aerospace vehicle to complete the mission when the engine thrust is partially reduced is one of the key technologies that need to be solved urgently in the development of the aerospace vehicle.
The iterative guidance method is an optimal control method with the current state as the initial value, the state of the orbit entry point as the goal, and the minimum propellant consumption as the performance index. It can adjust the flight path in real time according to the flight state of the aerospace vehicle, and has a certain adaptive ability. In the traditional iterative guidance algorithm, constant values are used for parameters related to the engine. However, when the engine actually has a thrust drop failure, the key parameters of the engine often have large changes. Therefore, the traditional iterative guidance method using constant parameters has certain deficiencies, and has inadequate ability to adapt to engine thrust drop failure.
Summary of the Invention
In order to overcome the problems existing in related technologies at least to a certain extent, the present application provides a self-adaptive iterative guidance method and a guidance device for an aerospace vehicle
Description
According to a first aspect of the embodiments of the present application, the present application provides a self-adaptive iterative guidance method for an aerospace vehicle, which includes the following steps.
updating an equivalent specific impulse of an engine in real time according to a measurement information of an inertia measurement unit, and using the equivalent specific impulse of the engine updated in real time to self-adaptively adjust the parameters related to an engine working state in an iterative guidance algorithm, to obtain a flight program angle adapted to an engine thrust drop failure.
In the above self-adaptive iterative guidance method for the aerospace vehicle, a process that, updating an equivalent specific impulse of an engine in real time according to a measurement information of an inertia measurement unit, is: selecting an equivalent specific impulse estimation cycle number, and when a current cycle is less than the equivalent specific impulse estimation cycle number, taking a specific impulse in a normal state of the engine as the equivalent specific impulse in the current cycle, and taking a consumption per second in the normal state of the engine as an equivalent consumption per second in the current cycle when the current cycle is greater than or equal to the equivalent specific impulse estimation cycle number, calculating an equivalent specific impulse estimation coefficient, according to an initial mass of the aerospace vehicle, the current cycle and the equivalent consumption per second of the current cycle; using the measurement information of the inertia measurement unit to perform a navigation calculation to obtain a deviation between an apparent velocity of the aerospace vehicle during flight and a standard trajectory apparent velocity; using the equivalent specific impulse estimation coefficient and the deviation between the apparent velocity of the aerospace vehicle during flight and the standard trajectory apparent velocity, to estimate an equivalent specific impulse deviation and the equivalent consumption per second of the current cycle; and using the equivalent specific impulse deviation and the equivalent consumption per second of the current cycle and the specific impulse and the consumption per second of the normal state of the engine to update the equivalent specific impulse and the equivalent
Description
consumption per second of the current cycle.
Further, a process that, when the current cycle is greater than or equal to the equivalent specific impulse estimation cycle number, calculating the equivalent specific impulse estimation coefficient, according to the initial mass of the aerospace vehicle, the current cycle and the equivalent consumption per second of the current cycle, is: carrying out Taylor expansion on a Tsiolkovsky formula W=tr, ln[rno /(mo -rry)] , and obtaining an apparent velocity increment as: AW=[ In mo -In(mo -mot)]Azt, + ft; I (mo -met)Ant, simplifying the equation of the apparent velocity increment as: AW=crAti +bAtn, wherein, W represents the apparent velocity, tie represents the equivalent specific impulse, Au represents the equivalent specific impulse deviation, me represents the equivalent consumption per second, and Am, represents an equivalent consumption per second deviation coefficients a and b are ja = ln -1n(rtio -met) lb = tt t I (mo -m01) calculating the apparent velocity increment of the k-ith cycle according to simplified equation of the apparent velocity increment, AW(k -1)=a(k -1)Au + b(k -1)Am; calculating equivalent specific impulse estimation coefficients a(k -i) and b(k -1) of the k-ith cycle according to the equations of coefficients a and b as, a(k -i)= In mo -In[mo (k -i)t(k -i)] lb(k -i)=u 0(k -01(k -[mo -mc(k -01(k -0] , wherein, i = 0, and t(k -i) represents a time corresponding to the k-ith cycle, me(k -) represents the equivalent consumption per second of the current cycle.
Description
Furthermore, a process that, using the measurement information of the inertia measurement unit to perform a navigation calculation to obtain a deviation between an apparent velocity of the aerospace vehicle during flight and a standard trajectory apparent velocity, is: AW(k)= W (k)-W (k), wherein, A W (k) represents the deviation between the apparent velocity of the aerospace vehicle during flight and the standard trajectory apparent velocity, W (k) represents a navigation apparent velocity of a km cycle, and 147(k) represents the standard trajectory apparent velocity of the km cycle, where k = 1, 2, L. Furthermore, a process that, using the equivalent specific impulse estimation coefficient and the deviation between the apparent velocity of the aerospace vehicle during flight and the standard trajectory apparent velocity, to estimate the equivalent specific impulse deviation and the equivalent consumption per second of the current cycle, is: using the apparent velocity increments of the last n cycles to estimate the equivalent specific impulse deviation and the equivalent consumption per second deviation online, where an estimation index is: i 1 = [AW (k a(k -0Au b(k)Arner from extremum conditions at -o and =0 obtaining: eAu eAnic {[frtia2(k -0]-Atte+ ft[a(k -i)* h(k -i)] l.,=i) [a(k -i)* b(k - Am = [a(k -.0 * AW(k J c i=0 * All e +[rh2 (k -0]*Am9 =T-[h(k AW (k -0] i=0 solving above equation set to obtain the equivalent specific impulse deviation Aid "(k) and the equivalent consumption per second A112,(k) of the kth cycle respectively: n-1 Ih2 (k -i)-I[a(k -i)* W (k -0]-I[a(k -i)* h(k -0]-I[h(k -0 * A W (k Ati e(k) = 7-0 7-0 Tir (k -*Ib2 (k - Jl[a(k)-b(k -=0
Description E /
a (k -*[b(k -)* A W(k - [a(k -i)* b(k -i)]*[a(k -)* A W (k -c(k) - ' 7 (k - (k -i)-2 [a(k -i)* b(k Furthermore, a process that, using the equivalent specific impulse deviation and the equivalent consumption per second of the current cycle and the specific impulse and the consumption per second of the normal state of the engine to update the equivalent specific impulse and the equivalent consumption per second of the current cycle, is: Ard, (k) when la-IC(01> C (k), when I Au + Arica) when Ain, (k) j> nic (k) =
_
when lArn,(k) IE Inc wherein, Au (k) represents the equivalent specific impulse deviation of the km cycle, Ainc(k) represents the equivalent consumption per second of the kii. cycle, i represents the specific impulse of the normal state of the engine, Pk represents the consumption per second of the normal state of the engine, it, (k) represents the equivalent specific impulse of the kih cycle, in,(k) represents the equivalent consumption per second of the kin cycle, e " represents a specific impulse deviation percentage allowed for a normal engine operation, and e represents a consumption per second deviation percentage allowed for the normal engine operation.
Furthermore, a process that, using the equivalent specific impulse of the engine updated in real time to self-adaptively adjust the parameters related to the engine working state in the iterative guidance algorithm, to obtain the flight program angle adapted to the engine thrust drop failure, is: using the equivalent specific impulse of the engine to update a complete combustion time, i using the updated equivalent specific impulse and the updated complete combustion time of the current cycle to update a remaining flight time,
Description
using the updated equivalent specific impulse, the updated complete combustion time and the updated remaining flight time of the current cycle to update integral parameters required for the iterative guidance; and obtaining an iterative guidance program angle according to the updated equivalent specific impulse, the updated complete combustion time and the updated remaining flight time of the current cycle, the updated integral parameters required for the iterative guidance, and the iterative guidance algorithm.
Furthermore, a process that, using the updated equivalent specific impulse and the updated complete combustion time of the current cycle to update a remaining flight time, is: the updated remaining flight time (k) is: s(k)= r(k)(1-e wherein r (k) represents the updated complete combustion time, AV represents a velocity increase, AV =[V (k)-g(k)f [Vs s(k) - (k)] 2 ± [l; - gst s(k)]2 wherein, [V,:, V,Fs] represents components of an expected velocity at an orbit entry point in an orbital coordinate system, [V (k), 1,7,(k), Tç (k)] represents components of a current velocity in the orbital coordinate system, and [4, jk-J represents components of an average value of gravitational accelerations at a current position and the orbit entry point in the orbital coordinate system Furthermore, a process that, using the updated equivalent specific impulse, the updated complete combustion time and the updated remaining flight time of the current cycle to update integral parameters required for the iterative guidance, is: the updated integral parameters 4,(k), A1 (k), A2 (k), 4(k) required for the iterative guidance are respectively:
Description
4"(k)= ite(k)ln r(k) r (k) -t g(k) A,(k)= e(k)t g(k)+ r(k)A,(k) A,(k)= Ao(k)t g(k)-A,(k) 4(k) = -0.5. e(k)Et g(k)T r(k) /42(k) Furthermore, in a process that obtaining the iterative guidance program angle according to the updated equivalent specific impulse, the updated complete combustion time and the updated remaining flight time of the current cycle, the updated integral parameters required for the iterative guidance, and the iterative guidance algorithm, when the iterative guidance program angle is a pitch program angle, calculating a reference program angle co(k) considering only a velocity constraint at the orbit entry point: -V, (k)-k,t,(k) 0(k)-atan[ -V, calculating program angle corrections Ky, and Ku,, considering the constraints of velocity and position at the orbit entry point: Ai(k)[11 -A,(k)sin[rp(k)] -0.5,[1 g(k)f -11(k)1,(k)-Y (k)} K y, - [-A,(k)A,(k)+ A. (k)A,(k)]cos[(k)] K 4(k){Y -A"(k)sin[5(k)]-0.5. k, g (k)f (k)1 8(k) -Y (k)) (k)A,(k)+ 4 (k)Ao(k)] cos[(0(k)] wherein, 17 represents a component of the position of the orbit entry point on a y-axis of the orbital coordinate system, and Y(k) represents a component of the current position on the y-axis of the orbital coordinate system; obtaining the pitch program angle tp(k) according to the reference program angle c7(k) of the velocity constraint at the orbit entry point and the program angle corrections Kc" and K4,2 for the constraints of velocity and position at the orbit entry point: (p(k)= 0(k)-K y, + K wherein, T represents an iterative guidance calculation cycle,
Description
performing a coordinate transformation and an amplitude limiting process on the pitch program angle co(k) to obtain a program angle instruction for output by the guidance system and for use by an attitude control system.
According to a second aspect of the embodiments of the present application, the present application also provides a self-adaptive iterative guidance device for an aerospace vehicle, which includes a memory and a processor, and the processor is configured to execute the steps in the self-adaptive iterative guidance method for the aerospace vehicle as described in any one of the above, based on an instruction stored in the memory.
According to the above specific embodiments of the present application, at least the following beneficial effects are achieved: the present application updates the equivalent specific impulse of the engine in real time, and utilizes the updated equivalent specific impulse of the engine to self-adaptively adjust the parameters related to the working state of the engine, such as the complete combustion time and the remaining flight time in the iterative guidance algorithm, thus obtains the flight program angle of the engine thrust drop failure, makes it possible to improve the adaptability of the iterative guidance algorithm to the engine thrust drop failure, improve the guidance accuracy, and improve the ability of the aerospace vehicle to complete flight missions in the event of the engine thrust drop failure. In the present application, by conducting online in real time updating on the equivalent specific impulse of the engine, the inertial navigation information can be used to estimate the working performance of the engine online.
It should be understood that the above general description and the following specific embodiments are only exemplary and explanatory, and are not intended to limit the scope of the present application.
Brief Description of the Drawings
The accompanying drawings, which are part of the specification of the present application, illustrate exemplary embodiments of the present application and, together with the description of the specification, serve to explain the principles of the present application Fig. 1 is a flow chart of a self-adaptive iterative guidance method for an aerospace vehicle provided in an embodiment of the present application.
Description
Fig. 2 is a flow chart of updating the equivalent specific impulse of the engine in real time in the self-adaptive iterative guidance method for the aerospace vehicle provided in the embodiment of the present application.
Fig. 3 is a flow chart of obtaining a flight program angle adapted to engine thrust drop failure in the self-adaptive iterative guidance method for the aerospace vehicle provided in the embodiment of the present application.
Detailed Description of Embodiments
In order to make the purposes, technical solutions and advantages of the embodiments of the present application clearer, the following will clearly illustrate the spirit of the content disclosed in the application with the accompanying drawings and detailed descriptions. Any person skilled in the art can change and modify the content of the application on the basis of the technology taught by the content of the present application after understanding the embodiments of the content of the present application, which does not depart from the spirit and scope of the content of the present application.
The exemplary embodiments and descriptions thereof of the present application are used to explain the present application, but not to limit the present application. In addition, elements/members with the same or similar numerals used in the drawings and the embodiments are used to represent the same or similar parts.
The terms "first", "second", . etc. used herein do not specifically refer to a sequence or order, nor are they used to limit the present application, but are only used to distinguish elements or operations described with the same technical terms.
As used herein, "comprising", "including", "having", "containing" and so on are all open terms, meaning including but not limited to.
As used herein, "and/or" includes any or all combinations of the stated things.
The "plurality" herein includes "two" and "two or more"; the "multiple groups" herein includes "two groups" and "two or more groups".
Certain terms used to describe the present application are discussed below or elsewhere in this specification to provide those skilled in the art with additional guidance in describing the present application.
Description
In the traditional iterative guidance algorithm for an aerospace vehicle, because constant values are adopted for the parameters related to the engine, the iterative guidance algorithm has insufficient adaptability to the engine thrust drop failure. In view of this technical problem, the present application uses the measurement information of the inertia measurement unit (abbreviated as the inertia unit) to obtain the deviation between the apparent velocity of the aerospace vehicle and the standard apparent velocity through navigation calculations, and uses the deviation to update the equivalent specific impulse of the engine in real time. Then, the parameters related to the working state of the engine, such as the complete combustion time and the remaining flight time in the iterative guidance algorithm, are self-adaptively adjusted to obtain the flight program angle that adapts to the engine thrust drop failure.
Fig. 1 is a flow chart of a self-adaptive iterative guidance method for an aerospace vehicle provided in an embodiment of the present application.
As shown in Fig. 1, the self-adaptive iterative guidance method for the aerospace vehicle provided by the embodiment of the present application includes the following steps: Si, as shown in Fig. 2, updating the equivalent specific impulse of the engine in real time according to the measurement information of the inertia measurement unit, the specific process is: Sit, selecting an equivalent specific impulse estimation cycle number n, when a current cycle k is less than the equivalent specific impulse estimation cycle number n, that is, when k<n, the equivalent specific impulse u(k) and the equivalent consumption per second me(k) of the kih cycle are the normal state calibration values provided by the engine, namely: ite(k)= (1) in,(k)= hi (2) in equation (1), represents the specific impulse of the engine in normal state, and in equation (2), Mc represents the consumption per second of the engine in normal state. Usually, the value of the equivalent specific impulse estimation cycle number n s: 25
Description
S12, when the current cycle k is greater than or equal to the equivalent specific impulse estimation cycle number n, that is, when k?-n, calculating equivalent specific impulse estimation coefficients a(k -) and b(k -i) according to an initial mass rn0 of the aerospace vehicle, the current cycle and the equivalent consumption per second of the current cycle, the specific process is: carrying out Taylor expansion on the Tsiolkovsky formula W=i), ln[ni, 1 (m0-ni0t)], and obtaining an apparent velocity increment as: A =[ I n aro -I n(n, -m01)]Aiç. + tic! 1 (tn,-m,t)Am, (3) equation (3) can be simplified as: AW=a47 +/),Am (4) in equation (3) and equation (4), W represents the apparent velocity, u, represents the equivalent specific impulse, An, represents the equivalent specific impulse deviation, In, represents the equivalent consumption per second, and Am, represents a deviation of the equivalent consumption per second.
In equation (4), the coefficients a and b are: Ja = ln in, -Inc° th =tt t 1 (mo-m,t) calculating the apparent velocity increment of the cycle according to equation (4), AW (k -1)-a(k -0411 + h(k (6) calculate the equivalent specific impulse estimation coefficients a(k -) and b(k -1) of the cycle according to equation (5), ja(k 1)-In mo-ln[mo -tnc(k -i)t(k (7) 1h(k -1)=71,(k -1)t(k -i)l[nk-me(k -1)t(k -i)] in equation (6) and equation (7), i = 0, 1, 2, L, n-1, and r(k -/) represents the time corresponding to the k-ith cycle, me (k -1) represents the equivalent consumption per second of the current cycle. (5)
Description
S13, using the measurement information of the inertia measurement unit to perform navigation calculations to obtain a deviation AW(k) between the apparent velocity of the aerospace vehicle during flight and the standard trajectory apparent velocity: AW(k)= W (k)-W (k) (8) in equation (8), 147(k) represents the navigation apparent velocity of the ki1 cycle, and I/17(k) represents the standard trajectory apparent velocity of the k111 cycle, where k = 1, 2, L. S14, using the equivalent specific impulse estimation coefficients a(k -0 and b(k -) and the deviation,A147(k) between the apparent velocity of the aerospace vehicle during flight and the standard trajectory apparent velocity, to estimate the equivalent specific impulse deviation Au,(k) and the equivalent consumption per second Am,(k) of the kih cycle, the specific process is: using the apparent velocity increments of the last n cycles to estimate the equivalent specific impulse deviation and the equivalent consumption per second deviation online, where an estimation index is: 1 =ii[AW (k -)-a(k -i)Au -b(k -1)Am "1- (9) =0 from extremum conditions ci! 61 -0 and = 0, obtaining: (-Azle ciAm, i (12(k -i)] * Ali, ± [a( k -i) * b(k - =E [ [a(k -0 * AW (k -i)]
J
1E[a(k -i)* b(k -i)] n 1 1 J Atte ± solving the above equation set (10) to obtain the equivalent specific impulse deviation Au (/c) and the equivalent consumption per second Ain, (k) of the kih cycle respectively: (k -i)-1[a(k -i)* W (k -i)]-I[a(k -i)* b(k -0]-i[b(k -i)* W (k Au (k)= z=() (k - (k -i)- [a(k -b(k -i)] 1-0 i-UL 1-0 (10)
Description E i
a (k -* [b(k -) * A W (k -I)] [a(k -i) * b(k -i)]*/[6-1(k -) * A W (k -c(k) - ' 7 (k - (k -i)-2 [a(k -i)* b(k S15, using the equivalent specific impulse deviation Au e(k) and the equivalent consumption per second Am(k) of the IQ, cycle and the specific impulse We and the consumption per second tk of the normal state of the engine to update the equivalent specific impulse u(k) and the equivalent consumption per second me(k) of the kth cycle, the specific process is: Au (k) whenlAudj(k)1> 6." * when c(k)liL e When A/17(k) N when lArn, * th (12) in equation (11) and equation (12), c, represents the specific impulse deviation percentage allowed for a normal engine operation, and s", represents the consumption per second deviation percentage allowed for the normal engine operation.
S2, as shown in Fig. 3, using the equivalent specific impulse of the engine updated in real time to self-adaptively adjust the parameters related to an engine working state in an iterative guidance algorithm, to obtain a flight program angle adapted to the engine thrust drop failure, the specific process is: 521, using the equivalent specific impulse of the engine to update a complete combustion time r(k), where, the updated complete combustion time r(k) is: c(k) (13) r(k) -
PFYL i r(k)
I in(k) = + At in equation (13), 14/4 represents the average apparent acceleration.
Description
S22, using the updated equivalent specific impulse tre(k) and the updated complete combustion time r(k) of the kth cycle to update a remaining flight time 15(k), where, the updated remaining flight time t g(k) is: AT" tg(k)= r(k)(1-e (14) in equation (14), AV represents a velocity increase, AV =[[ç -V(k)/g(k)I2 +[11, (k) -Th.! (k)I2 +[v, -11(k)-grig(k)r (15) in equation (15), [r.," r.,.,17,] represents components of an expected velocity at an orbit entry point in an orbital coordinate system, [V, (/c),V(k),1/(k)] represents components of a current velocity in the orbital coordinate system, and [kr, kr] represents components of an average value of gravitational accelerations at a current position and the orbit entry point in the orbital coordinate system S23, using the updated equivalent specific impulse iir(k) , the updated complete combustion time r(k) and the updated remaining flight time t g(k) of the kth cycle to update integral parameters required for the iterative guidance, where, the updated integral parameters 4,(k), A, (k), A, (k), 4(k) required for the iterative guidance are respectively.
4,(k) = lc (k)ln r(k) (16) 7-(k)-tg (k) A, (k)= -lle(k)t g(k) + r(k)4(k) 4(k) = Ao(k)t g(k) -A1(k) 4(k) = -0.5 -II, (k)Lt g(k)12 + r(k)4(k) S24, obtaining an iterative guidance program angle according to the updated equivalent specific impulse it e(k) , the updated complete combustion time 7-(k) and the updated remaining flight time t g(k) of the kth cycle, the updated integral parameters required for the iterative guidance, and the iterative guidance algorithm.
Description
The following takes a calculation of a pitch program angle as an example to describe step S24 in detail.
Calculating a reference program angle)(k) considering only a velocity constraint at the orbit entry point: c70(k)-atan[ g V -V (k)-gt (k) -V x(k)-g(k) (17) calculating program angle corrections K1,1 and K,,, considering the constraints of velocity and position at the orbit entry point: A,(k){Y -At,(k)sin[(k)]-0.5. glig(k)12 -V,.(k)tg(k)-Y (k)} K 1,,- [-A1(k)24.2(k)+ 4,(k)240(k)]cos[(9(k)] 4"(k)1Y -Ai, (k) sink:6(k)] -0.5-ltg(k)r. -V-,.(k)t, (k) -Y (k)} 102 - [-A1(k)A2(k)+ A3(k)4(k)] cos[(p(k)] (18) K in equation (18), 37 represents a component of the position of the orbit entry point on a y-axis of the orbital coordinate system, and Y(k) represents a component of the current position on the y-axis of the orbital coordinate system.
Obtaining the pitch program angle q(k) according to the reference program angle (o(k) of the velocity constraint at the orbit entry point and the program angle corrections K921 and lc for the constraints of velocity and position at the orbit entry point: qt(k)=(k)- +K92,T (19) in equation (19), T represents an iterative guidance calculation cycle.
Performing a coordinate transformation and an amplitude limiting process on the pitch program angle cy(k) to obtain a program angle instmction for output by the guidance system and for use by an attitude control system The self-adaptive iterative guidance method for an aerospace vehicle of the present application updates the equivalent specific impulse of the engine online in real time, and utilizes the updated equivalent specific impulse of the engine to self-adaptively adjust the iterative guidance parameters, makes it possible to improve the adaptability of the iterative guidance algorithm to the engine thrust drop failure, and improve the guidance accuracy.
Description
In an exemplary embodiment, the embodiment of the present application also provides a self-adaptive iterative guidance device for an aerospace vehicle, which includes a memory and a processor, wherein the processor is configured to execute the self-adaptive iterative guidance method for an aerospace vehicle in any one of the embodiments in the present application, according to the instructions stored in the memory.
Wherein, the memory may be a system memory, a fixed non-volatile storage medium or the like, and the system memory may store an operating system, an application program, a boot loader, a database, and other programs In an exemplary embodiment, the embodiment of the present application also provides a computer storage medium, which is a computer-readable storage medium, for example, a memory including a computer program, and the above computer program can be executed by a processor to complete the self-adaptive iterative guidance method for an aerospace vehicle in any one of the embodiments in the present application The above is only an illustrative specific embodiment of the present application.
Without departing from the concept and principle of the present application, any equivalent changes and modifications made by those skilled in the art shall fall within the protection scope of the present application.
Claims (1)
- Claims 1. A self-adaptive iterative guidance method for an aerospace vehicle, characterized by comprising the following steps: updating an equivalent specific impulse of an engine in real time according to a measurement information of an inertia measurement unit, a process of this step is: selecting an equivalent specific impulse estimation cycle number, and when a current cycle is less than the equivalent specific impulse estimation cycle number, taking a specific impulse in a normal state of the engine as the equivalent specific impulse in the current cycle, and taking a consumption per second in the normal state of the engine as an equivalent consumption per second in the current cycle; when the current cycle is greater than or equal to the equivalent specific impulse estimation cycle number, calculating an equivalent specific impulse estimation coefficient, according to an initial mass of the aerospace vehicle, the current cycle and the equivalent consumption per second of the current cycle; using the measurement information of the inertia measurement unit to perform a navigation calculation to obtain a deviation between an apparent velocity of the aerospace vehicle during flight and a standard trajectory apparent velocity; using the equivalent specific impulse estimation coefficient and the deviation between the apparent velocity of the aerospace vehicle during flight and the standard trajectory apparent velocity, to estimate an equivalent specific impulse deviation and the equivalent consumption per second of the current cycle; using the equivalent specific impulse deviation and the equivalent consumption per second of the current cycle and the specific impulse and the consumption per second of the normal state of the engine to update the equivalent specific impulse and the equivalent consumption per second of the current cycle, and using the equivalent specific impulse of the engine updated in real time to self-adaptively adjust the parameters related to an engine working state in an iterative guidance algorithm, to obtain a flight program angle adapted to an engine thrust drop failure Claims 2. The self-adaptive iterative guidance method for the aerospace vehicle according to claim 1, wherein a process that, when the current cycle is greater than or equal to the equivalent specific impulse estimation cycle number, calculating the equivalent specific impulse estimation coefficient, according to the initial mass of the aerospace vehicle, the current cycle and the equivalent consumption per second of the current cycle, is: carrying out Taylor expansion on a Tsiolkovsky formula W=heln[m, I (m, -mat)] , and obtaining an apparent velocity increment as: AW=[ In m0 -In(m -att)]Azte +net I (m0-met)Amc. simplifying the equation of the apparent velocity increment as: AW =a6at + b Am, . wherein, W represents the apparent velocity, he represents the equivalent specific impulse, Au, represents the equivalent specific impulse deviation, In, represents the equivalent consumption per second, and Am, represents an equivalent consumption per second deviation, coefficients a and b are: la = In mo -In(mo -met) b =11,1 I (mo-in)) calculating the apparent velocity increment of the k-ith cycle according to simplified equation of the apparent velocity increment, AW(k -I)=a(k -OAP, + b(k -0Am. ; calculating equivalent specific impulse estimation coefficients a(k -i) and h(k -I) of the cycle according to the equations of coefficients a and b as, a(k -i)= In m, -In[mo -mc(k -i)t(k -i)] lb(k -i) = u,e(k -i)I(k -i) I [mo -Inc(k -1)1(k -i)] wherein, i = 0, 1, 2, L, n and t(k -1) represents a time corresponding to the k-i111 cycle, Claims m 6(k -1) represents the equivalent consumption per second of the current cycle.3. The self-adaptive iterative guidance method for the aerospace vehicle according to claim 2, wherein a process that, using the measurement information of the inertia measurement unit to perform a navigation calculation to obtain a deviation between an apparent velocity of the aerospace vehicle during flight and a standard trajectory apparent velocity, is: AW(k) = W (k)-W(k), wherein, AW (k) represents the deviation between the apparent velocity of the aerospace vehicle during flight and the standard trajectory apparent velocity, W (k) represents a navigation apparent velocity of a km cycle, and IF(k) represents the standard trajectory apparent velocity of the km cycle, where k = 1, 2, L. 4. The self-adaptive iterative guidance method for the aerospace vehicle according to claim 3, wherein a process that, using the equivalent specific impulse estimation coefficient and the deviation between the apparent velocity of the aerospace vehicle during flight and the standard trajectory apparent velocity, to estimate the equivalent specific impulse deviation and the equivalent consumption per second of the current cycle, is: using the apparent velocity increments of the last n cycles to estimate the equivalent specific impulse deviation and the equivalent consumption per second deviation online, where an estimation index is: 1 =1[AW (k -i)-a(k -i)Alle -b(k -)Atne1-i=0 from extremum conditions ai -0 and -0, obtaining: DAtte DAm, Claims { [1a2 (k -0]. An, +{[a(k -i)* h(k -solving above equation set to obtain the equivalent specific impulse deviation Au e(k) and the equivalent consumption per second Arne(k) of the k111 cycle respectively: (k -)-[a(k -0 * AW (k [a(k -0 * b(k -[b(k -i)* AW (k -0] AUe (k)- ,=0,,0 z=0 (k -i)*1,2 (k -i)-i[a(k -0 * b(k -0 i 0, 0 / -i)-[b(k -A W (1/ -0] -/[a(k -b(k -0]-1[6i(k -0 * AI (k -0] Am c(k)= a2 (k --162 (k - ))1[a(k -b(k 5. The self-adaptive iterative guidance method for the aerospace vehicle according to claim 4, wherein a process that, using the equivalent specific impulse deviation and the equivalent consumption per second of the current cycle and the specific impulse and the consumption per second of the normal state of the engine to update the equivalent specific impulse and the equivalent consumption per second of the current cycle, is Sue(k) when 14*(01> s" * when I Ai t"(k)l-s: 6,, 17, in" +Arnu(k) when AM, 00 I> En -k) = Irk when lAm (k) Ic e * wherein, Au (k) represents the equivalent specific impulse deviation of the kii, cycle, Am, (k) represents the equivalent consumption per second of the kth cycle, k represents the specific impulse of the normal state of the engine, ik represents the consumption per second of the normal state of the engine, u e(k) represents the equivalent specific impulse of the kth {r[a(k -i)* b(k t, in =1[a(k -0 * Aff:(k Au, +[b2 (k -0]*=1[b(k (k -0] =0 i=0 )1 i=Cl i=Cl i= Claims cycle, tric(k) represents the equivalent consumption per second of the kw cycle, s" represents a specific impulse deviation percentage allowed for a normal engine operation, and s", represents a consumption per second deviation percentage allowed for the normal engine operation.6. The self-adaptive iterative guidance method for the aerospace vehicle according to claim 1, wherein a process that, using the equivalent specific impulse of the engine updated in real time to self-adaptively adjust the parameters related to the engine working state in the iterative guidance algorithm, to obtain the flight program angle adapted to the engine thrust drop failure, is: using the equivalent specific impulse of the engine to update a complete combustion time; using the updated equivalent specific impulse and the updated complete combustion time of the current cycle to update a remaining flight time, using the updated equivalent specific impulse, the updated complete combustion time and the updated remaining flight time of the current cycle to update integral parameters required for the iterative guidance; and obtaining an iterative guidance program angle according to the updated equivalent specific impulse, the updated complete combustion time and the updated remaining flight time of the current cycle, the updated integral parameters required for the iterative guidance, and the iterative guidance algorithm.7. The self-adaptive iterative guidance method for the aerospace vehicle according to claim 6, wherein a process that, using the updated equivalent specific impulse and the updated complete combustion time of the current cycle to update a remaining flight time, is: the updated remaining flight time tg(k) is: Claims Air t (k)= r(k)(1-wherein, r(k) represents the updated complete combustion time, AV represents a velocity increase, AV = \AI; (k)- (k)] ,(k)-,t f(k)] 2 ± [1-; -(k) -g_tf(k)]2 wherein, , Vs] represents components of an expected velocity at an orbit entry point in an orbital coordinate system, [I 1 (0,1/"(k),V(k)] represents components of a current velocity in the orbital coordinate system, and [k." , Tr: represents components of an average value of gravitational accelerations at a current position and the orbit entry point in the orbital coordinate system.8. The self-adaptive iterative guidance method for the aerospace vehicle according to claim 7, wherein a process that, using the updated equivalent specific impulse, the updated complete combustion time and the updated remaining flight time of the current cycle to update integral parameters required for the iterative guidance, is: the updated integral parameters (k), (k), A,(k), (k) required for the iterative guidance are respectively: Ao(k)= t (k)In r(k) r(k) -t g(k) (k) = A,(k) = Ao(k)t g(k) (k) (k)= -0.5. tt,(k)[1. g(k)12 + r(k)4(k) 9. The self-adaptive iterative guidance method for the aerospace vehicle according to claim 8, wherein, in a process that obtaining the iterative guidance program angle according to the updated equivalent specific impulse, the updated complete combustion time and the updated Claims remaining flight time of the current cycle, the updated integral parameters required for the iterative guidance, and the iterative guidance algorithm, when the iterative guidance program angle is a pitch program angle, calculating a reference program angle co(k) considering only a velocity constraint at the orbit entry point: V, -V, (k)-,t,,(k) (o(k)=atan[ ] 17, -(k)-Lt g(k) calculating program angle corrections 1(1,, and K1,2 considering the constraints of velocity and position at the orbit entry point: {.1 c.,,, A,(k){Y -A, (k) sink T 3 (01 -0.5. T crt g(k)r -V (k)t g(k)-Y (k)) [-A1(k)4(k)+ As(k)A0(k)]cos[o(k)] 4,(k1) Y -Ai, (k)sinp(k)]-0 5. Tr, [t g(k)r -I-, (k)c(k)-Y (k)) wherein, 17 represents a component of the position of the orbit entry point on a y-axis of the orbital coordinate system, and Y (k) represents a component of the current position on the y-axis of the orbital coordinate system; obtaining the pitch program angle co(k) according to the reference program angle @(k) of the velocity constraint at the orbit entry point and the program angle corrections K1,1 and for the constraints of velocity and position at the orbit entry point: y(k) = (k) -K + wherein, T represents an iterative guidance calculation cycle; performing a coordinate transformation and an amplitude limiting process on the pitch program angle (4k) to obtain a program angle instruction for output by the guidance system and for use by an attitude control system K [-4(k)/4",(k)+ A;(k)A0(k)]e0s[cO(k)] Claims 10. A self-adaptive iterative guidance device for an aerospace vehicle, characterized by comprising a memory and a processor, the processor is configured to execute the steps in the self-adaptive iterative guidance method for the aerospace vehicle as in any one of claims 1-9, based on an instruction stored in the memory.
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| CN112810834B (en) * | 2020-12-23 | 2022-11-11 | 北京航天自动控制研究所 | Ground test method for simultaneously checking inertial navigation and simulating flight |
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