CN102393200B - General inertial navigation test method based on flight simulation - Google Patents

Abstract

The invention discloses a general inertial navigation test method based on flight simulation. The method comprises the following steps: establishing a flight control simulation module based on a control rate design, designing the flight path, aiming inertial navigation initial conditions, implementing a flight simulation process, and outputting the flight state data of flights; inversely solving and calculating the specific force and angular rate data by utilizing the flight state data; solving and calculating the flight state data of the flights in real time according to the specific force and angular rate; and comparing the flight state data output by the flight control simulation module and the flight state data output by an inertial navigation system according to consistent time shot, and comparing the different information of the data, thus detecting the inertial navigation performance. Through the method provided by the invention, not only can dynamic continuous inertial navigation performance test be intensified, but also the method accords with the test requirements of the complex working environment in the air on the inertial navigation system, and has great significance to engineering practical application of the inertial navigation system performance testing.

Description

General inertial navigation method of testing based on flight simulation

Technical field

The present invention relates to a kind of inertial navigation system method of testing under continuous, the aerial practical flight of dynamic similation, is the method for testing performance of a general inertial navigation system of using towards engineering.

Background technology

Test for the inertial navigation system performance, generally need carry out the test of every index at multiaxis turntable or testing apparatus, as the personal error at traditional inertial navigation test such as Liu Yuhao (Chinese inertial technology journal, 2008), a kind of automatic test and calibration system based on turntable proposed.Chen Cai etc. (systems engineering and electronic technology, 2010) have proposed based on the Inertial navigation platform error coefficient discrimination method of hydro-extractor test according to the turntable error etc.Tradition inertial navigation test can only be carried out the test under the discrete state condition, and the inertial navigation performance is tested in the time of can not getting off the plane the non-stop flight state to true aerial various factors condition.

And in actual taking a flight test, carry out the inertial navigation performance test, though can carry out the non-stop flight test of aerial various complex conditions, there are problems such as cost height, test period length, very flexible.Therefore, study a kind of artificial actual aircraft flight, can be continuously, the method and system dynamically the inertial navigation performance tested has important significance for theories and engineering using value.

Summary of the invention

In order to overcome the deficiencies in the prior art, the present invention proposes a kind of based under the flight simulation excitation of control rate design the inertial navigation performance test methods, not only strengthened dynamic continuous inertial navigation performance test, and meet aerial complicacy working environment to the requirement of inertial navigation system test, have very important significance for the engineering practical application of inertial navigation system performance test.

The technical solution adopted for the present invention to solve the technical problems comprises the steps:

The first step is set up the flight control emulation module based on the control rate design, for the inertial navigation system performance test provides the data driving source.Utilization is carried out nonlinear equation by known model aircraft parameter and is derived and calculating based on the design of aircraft control law, and carries out the vertical and side direction control law design of aircraft, realizes the motion under the aircraft typical Mode.

Second step, carry out flight path setting, inertial navigation starting condition aligning, implement the flight simulation process, and output flight status data, specifically may further comprise the steps:

1, carry out flight path and set, flight path comprises aircraft initial state information (position, attitude, speed) and other boat

Mark dot position information, flight path are set to contain in the air and are climbed, glide, spiral, turn and comprehensive non-stop flight mode.It is poor that flight mode is characterised in that whether the relevant position is arranged between two track points.As climb flight mode basis for estimation be whether two track points have difference in height, whether other also is respectively two track points longitude, latitude difference.If two track points have corresponding flight path coordinate difference, then namely represent corresponding flight mode.

2, inertial navigation initial alignment, the starting condition data are consistent during namely the Flight Condition Data under the inertial navigation starting condition and flight path setting;

3, real-time resolving and output is by the Flight Condition Data of the aircraft of flight control emulation module output, comprises aircraft state information such as speed (m/s), the angle of attack (rad), yaw angle (rad), rate of pitch (rad/s), roll angle speed (rad/s), yaw rate (rad/s), the angle of pitch (rad), roll angle (rad), crab angle (rad), aircraft position (longitude, latitude, highly).

The 3rd step is by anti-specific force and the angular rate data of calculating of Flight Condition Data.By longitude, latitude and the height (L, λ, H) that the flight simulation module provides, reach geographic coordinate system and get off the plane east orientation, north orientation and sky to speed (V _E, V _N, V _U) calculating inertial coordinates system, to tie up to navigation coordinate with respect to terrestrial coordinates be the angular speed projection

Terrestrial coordinate system ties up to the angular speed projection down of navigation coordinate system with respect to navigation coordinate

Inertial coordinates system is with respect to navigation coordinate angular speed projection under navigation coordinate system Aircraft pitch, roll and the course angle (θ, γ, ψ) that provide by the flight simulation module calculate angular velocity under the body axis system according to the Eulerian angle differential equation

Attitude matrix

Utilize

Calculate angular speed under the gyro angular speed output body axis system

Utilize the specific force equation to calculate specific force f under acceleration and the output body coordinate ^b

In the 4th step, inertial navigation system goes out the flight status data according to specific force and angular speed real-time resolving.Specific force and angular rate data are transferred to the inertial navigation system that will test by communication line, and inertial navigation system carries out corresponding data solver, by communication line Flight Condition Data is exported.

The 5th step, according to the time clap the consistent Flight Condition Data of flight control emulation module output and the Flight Condition Data that inertial navigation system is exported of contrast, compare data differences information between the two, thereby the inertial navigation performance detected.

Inertial navigation test system hardware based on flight simulation among the present invention encourages communication line between computing machine, supervisory control comuter, communication card, the computing machine, computing machine and inertial navigation system communication line to form by emulation, and communication card is settled at emulation excitation computing machine, supervisory control comuter respectively.Finish flight control emulation at emulation excitation computing machine, instead resolve, the functions such as transmission of data, finish the functions such as reception, demonstration, record, analysis of data at supervisory control comuter.Communication line and communication card communicate with computing machine and inertial navigation system respectively, realize between the computing machine, the transfer function of data between computing machine and the inertial navigation system.The Flight Condition Data that emulation excitation computing machine produces directly is transferred to supervisory control comuter by communication card and communication line; Specific force, angular rate data that emulation excitation computing machine produces are transferred to inertial navigation system by communication card and communication line; The Flight Condition Data that inertial navigation system calculates receives data by communication line and the integrated circuit board on supervisory control comuter.

The invention has the beneficial effects as follows:

The switching that the turntable test of tradition ground can not be simulated aerial complicated non-stop flight state and state of flight.This programme can be simulated aerial Live Flying based on the flight simulation of control rate design and be controlled and produce dynamically continuous track, can realize the performance test of inertial navigation system under many flight mode, continuous dynamically flight.

The method of testing that this programme provides can provide checking in the early stage of ground data and data-optimized and improvement in performance work in early stage for truly taking a flight test in the air.

The enforcement of this programme has solved problems such as the cost height of directly taking a flight test, test period length, very flexible.

It is convenient, flexible that this programme is implemented, and has development prospect and engineering using value.

Description of drawings

Fig. 1 forms for overall system hardware.

Fig. 2 forms for system software module.

Fig. 3 forms structure for flight control emulation module.

Fig. 4 is the overall calculation flow chart of method

Fig. 5 is course line and its perspective view.

Fig. 6 is track initiation point and impact point height relativeness.

Aircraft and course line relation when Fig. 7 is flight.

Embodiment

The present invention is further described below in conjunction with drawings and Examples.

Be the overall system hardware design as Fig. 1, hardware encourages communication line between computing machine, supervisory control comuter, communication card, the computing machine, computing machine and inertial navigation system communication line to form by emulation, and communication card is settled at emulation excitation computing machine, supervisory control comuter respectively.

The overall system software module as shown in Figure 2.Software module is formed and has been comprised emulation excitation subsystem, emulation Monitor And Control Subsystem.In the excitation subsystem, control emulation module, the anti-compositions such as module, data transmission, communication module that resolve by flight; Monitor And Control Subsystem is made up of modules such as communication module, data reception module, data demonstration, data recording, calculating and analyses.

Fig. 4 is method concrete calculation flow chart when implementing.When implementing the present invention program, can at first set up correlation function modules such as flight control module, guide module processed.Carry out the inertial navigation performance test according to Fig. 4 flow process then.When concrete test, at first carry out the flight path setting, utilize to set up the guidance of flying of good flight control analogue system, produce height, crab angle control, carry out vertical, horizontal side direction control law design respectively, control elevating rudder, yaw rudder, aileron model move, and add that the thrust that Atmospheric models are calculated, thrust meter is calculated drives model aircraft, thereby produce Flight Condition Data.Flight Condition Data is compared with the required position calculation in real-time course line in the guide module processed, constantly produces guidance information.Simultaneously, Flight Condition Data will resolve that module produces specific force, angular rate data offers inertial navigation system by inertial navigation is counter.The 3rd, Flight Condition Data also directly is transferred to supervisory control comuter, and inertial navigation system produces other one group of Flight Condition Data according to specific force, angular rate data, the two is compared, thereby can detect the inertial navigation performance.

The concrete implementation step details of each several part is as follows:

1, sets up flight control analogue system

Design to realize flying to control the foundation of analogue system based on the aircraft control law, its construction module and commutative relation are as shown in Figure 3.In this structure, mainly consist of the following components:

1.1 flight guidance

Positional information according to flight path data message and current aircraft, judge whether aircraft is on the course line, place, carry out height, crab angle calculating, thereby produce vertical and horizontal side direction control respectively, give to fly to control module all kinds of mode control informations are provided, make the aircraft can be according to predetermined airline operation.

1.1.1 highly control

In flight path control, highly change, the current plane projection position according to prebriefed pattern and aircraft calculates the height that aircraft should reach in current location in real time, this is carried out real-time follow-up by flying to control module.

Suppose that A-B is the course line according to 1.1 described settings, C (X _C, 0, Z _C), D (X _D, 0, Z _D) be respectively A (X _A, Y _A, Z _A), B (X _B, Y _B, Z _B) projection on surface level XOZ, have: X _A=X _C, Z _A=Z _C, X _B=X _D, Z _B=Z _D

Aspect height control, can calculate the assigned altitute in course line by projection.As shown in Figure 4, suppose that the current point of aircraft is the E point, calculate to such an extent that aircraft projection in the plane should be the F point by non-level flight range so.In order to obtain the height that aircraft should reach, elder generation to the C-D of course line, is the G point with the F spot projection.Obtain and just can find behind the G point coordinate course line A-B to go up the coordinate H point that G is ordered that is projected as in the XOZ plane, then can obtain the height that aircraft should reach.

Among Fig. 5, C, D and F coordinate are known, calculate the G point coordinate, wherein (X _D-X _C) and (Z _D-Z _C) can not be zero simultaneously, because the course line of planning can not be vertical with ground level.Calculate X _G, Z _G

$\left\{\begin{array}{c}{X}_G=\frac{{(X_D-X_C)}^2{X}_F+{(Z_D-Z_C)}^2{X}_C+(Z_F-Z_C)(Z_D-Z_C)(X_D-X_C)}{{(X_D-X_C)}^2+{(Z_D-Z_C)}^2}\\ Z_G=\frac{{(Z_D-Z_C)}^2{Z}_F+{(X_D-X_C)}^2{Z}_C+(X_F-X_C)(Z_D-Z_C)(X_D-X_C)}{{(X_D-X_C)}^2+{(Z_D-Z_C)}^2}\end{array}\right.---(1-1)$

ABDC separates with the plane, and as shown in Figure 6, A, B, C, D, G coordinate are known.Can be released the relation of each straight-line segment by Fig. 6:

$\frac{\mathrm{AJ}}{\mathrm{HJ}}=\frac{\mathrm{AI}}{\mathrm{BI}}\⇔\frac{\mathrm{CG}}{\mathrm{HJ}}=\frac{\mathrm{CD}}{\mathrm{BI}}\⇒\mathrm{HJ}=\frac{\mathrm{CG}}{\mathrm{CD}}\mathrm{BI}---(1-2)$

And then draw the height value that H is ordered, the height h that namely should arrive.

When the course line starting point is lower than impact point:

$h=Y_H=Y_A+\frac{\sqrt{{(X_C-X_G)}^2+{(Z_C-Z_G)}^2}}{{(X_C-X_D)}^2+{(Z_C-Z_D)}^2}(Y_B-Y_A)=Y_A+K(Y_B-Y_A)---(1-3)$

When in like manner starting point was higher than impact point, the height value that H is ordered was the height h that reaches:

$h=Y_H=Y_A+\frac{\sqrt{{(X_C-X_G)}^2+{(Z_C-Z_G)}^2}}{{(X_C-X_D)}^2+{(Z_C-Z_D)}^2}(Y_A-Y_B)=Y_A-K(Y_A-Y_B)---(1-4)$

When starting point is identical with the impact point height, the height h=Y that should reach _B

1.1.2 crab angle control

The driftage distance as the course controlled quentity controlled variable, participates in the control of aircraft level navigation, as judging whether to arrive impact point in navigation.The principal element that influences flight track is the course angle of aircraft, the course angle of aircraft should be consistent with the prebriefed pattern flight-path angle when the aircraft flight flight path is on prebriefed pattern, when producing driftage, should there be deviation at the vector angle with the prebriefed pattern flight-path angle, by driftage is obtained given crab angle apart from carrying out PID control.

According to set route in the plane projection and the physical location of aircraft relatively, calculate the driftage distance, and then calculate height and the course controlled quentity controlled variable is flown to control module, fly to control the real-time follow-up of the course realization prebriefed pattern of module controls aircraft.

Model aircraft resolve obtain the aircraft current position coordinates will be for the calculating of driftage distance, driftage apart from navigation as the course controlled quentity controlled variable, participate in the control of aircraft level navigation, can also conduct judge whether to arrive impact point simultaneously.

By shown in Figure 5, aircraft and course line relation during navigation, driftage apart from the computing formula of Δ d is:

$\mathrm{\Δd}=\frac{|(Z_D-Z_C)X_F-(X_D-X_C)Z_F+Z_C{X}_D-Z_D{X}_C|}{\sqrt{{(X_D-X_C)}^2+{(Z_D-Z_C)}^2}}---(1-5)$

In Fig. 7, the course line vector

Angle (course line flight-path angle) ψ with direct north ₀For:

In the formula, Δ z=z ₂-z ₁, Δ x=x ₂-x ₂

The principal element that influences flight track is the course angle of aircraft, and the course angle of aircraft should be consistent with the prebriefed pattern flight-path angle when the aircraft flight flight path is on prebriefed pattern, and when producing driftage, should there be deviation at the vector angle with the prebriefed pattern flight-path angle.So, desirable course angle (given course angle) ψ of aircraft during driftage _gShould be:

${\mathrm{\ψ}}_g={\mathrm{\ψ}}_0+K_P\mathrm{\Δd}+K_I{\∫}_0^t\mathrm{\Δddt}+K_D\frac{d\left(\mathrm{\Δd}\right)}{\mathrm{dt}}---(1-7)$

In the formula: K _PBe scale-up factor; K _IBe integral coefficient; K _DBe differential coefficient, obtain by debugging.

Need discretize when actual computation, T is computation period.Be shown below after discrete:

${\mathrm{\ψ}}_g\left(k\right)={\mathrm{\ψ}}_0+K_p\mathrm{\Δd}\left(k\right)+K_{l}T\underset{i=0}{\overset{k}{\mathrm{\Σ}}}\mathrm{\Δd}\left(i\right)+K_D\frac{\mathrm{\Δd}\left(k\right)-\mathrm{\Δd}(k-1)}{T}---(1-8)$

In flight course, at first, can determine the direction ψ in course line according to two voyages point ₀When flying behind impact point, calculate the direction ψ in the course line that makes new advances again ₀, the rest may be inferred.

1.2 the flight control based on the control rate design

Flying to control Module Design mainly realizes the emulation of typical state of flight (flat flying climbed, and glides, and turns left, and turns right and waits each mode and comprehensive mode) is solved the logic switching problem of complicated flight control.

The Flight Control Law modular design mainly is divided into vertical control law design and the design of side direction control law, carrying out control law resolves, this module is according to the aeroplane performances such as aerodynamic parameter of all kinds of aircrafts own, transfer ginseng to finish the control rate design according to PID, the value that control law resolves is out carried out deflection as the controlled quentity controlled variable control rudder face of topworks.

1.2.1 vertically control rate design

The main basic function that lengthwise movement will realize is as follows:

(1) waits the angle of pitch to climb with the given angle of pitch or glide, keep pitching stable.

(2) put down with assigned altitute and fly, keep the flying height of aircraft.

A. pitch loop control law design

Aircraft uses pitch attitude to keep/control mode under downslide, climb mode, is the damping torque that increases aircraft, introduces angle rate signal, with the variable of angle of pitch increment Delta θ as control.

B. highly keep the design of circuit controls rule

Highly stable control module is input to angle of pitch control module with the difference in height signal, changes the flight path pitch angle, and the lifting of control aircraft is zero until difference in height, and aircraft is got back to till the predetermined height.

1.2.2 side direction control law design

The purpose of sideway movement control is at different mode of motion, takes different measures to guarantee that aircraft has good performance.The main basic function that sideway movement will realize is as follows:

(1) carries out the uniform slope orbit with given roll angle.

(2) course keeps mode, gives vectoring, aircraft is forwarded to vectoring flight, and realize coordinate turn.

A. the design of the mode of spiraling control law

Aircraft changes course or spirals when turning, and aircraft is general all by means of the pitch angle control module, imports given control signal γ _g, make craft inclination, change the flight path deflection angle by the side force that produces behind the craft inclination, reach the purpose that changes vector.

B. course retentive control rule design

Course control is coordinated control with aileron and yaw rudder simultaneously, stable and control course, and course signal is sent the aileron passage, the lift-over signal incoming direction rudder passage that simultaneously the aileron passage is produced.

1.3 set up topworks

By yaw rudder, elevating rudder, steering wheel models such as aileron are formed, and this module adopts the design of reaction type rudder loop, the instruction that flies to control module can compare situation with the planning flight path by state of flight and produce angle of rudder reflection, offers aircraft six degree of freedom model (1.7 is described) as the input controlled quentity controlled variable.

1.4 set up Atmospheric models

Be atmospheric gas pressure, temperature with the height change model, calculate air pressure and the temperature of place height by height value, calculating for thrust provides required temperature, barometric information.Account form is undertaken by international atmosphere ISA model in this module.

1.5 thrust is calculated

According to place height, temperature, air pressure, and corresponding aircraft-position information, suitable throttle thrust calculated to drive aircraft flight.This programme is taked the motor power testing experiment, the corresponding relation of accelerator open degree and thrust under the friction speed of acquisition, and the engine gear that provides according to instruction control and model resolve the velocity amplitude that obtains and carry out two-dimensional interpolation, obtain the thrust magnitude of engine.Thrust result of calculation is exported to model aircraft, makes flight model obtain corresponding thrust.

1.6 set up model aircraft

Aircraft is regarded as rigid body, and aircraft has six-freedom degree in the motion in space, i.e. three of barycenter one-movement-freedom-degrees and around three rotational freedoms of barycenter.The degree of freedom of three movements of barycenter is the increase and decrease campaign of speed, oscilaltion campaign and left and right sides shifting movement.Three rotational freedoms are pitch movement, crab angle motion and roll angle motion.The non-linear full dose equation of the six degree of freedom of aircraft is made up of 12 differential equation of first orders, comprises six kinetics equations and six kinematical equations.

When the non-linear full dose equation of the six degree of freedom of derivation aircraft, need carry out following hypothesis to aircraft:

Think that aircraft is not only rigid body, and quality is constant;

Suppose that ground is inertial reference system, namely coordinate is inertial coordinate hypothetically;

Ignore earth curvature, looking ground is the plane;

Suppose that acceleration of gravity does not change with flying height;

Suppose the x of body axis system _b-O _b-z _bThe plane is the symmetrical plane of aircraft, and aircraft geometric shape symmetry not only, and internal soundness is distributed also symmetry.So product of inertia I _Xy=I _Zy=0.

The computing formula of three-axis force and three-axis force square is as follows:

$\left\{\begin{array}{c}Q=C_{x}q{S}_w\\ Y=C_{y}q{S}_w\\ C=C_{z}q{S}_w\\ M_x=m_{x}q{S}_{w}L\\ M_y=m_{y}q{S}_{w}L\\ M_z=m_{z}q{S}_w{b}_A\end{array}\right.---(1-9)$

Wherein, Q, Y, C are respectively resistance, lift, side force.M _x, M _y, M _zBe respectively rolling moment, yawing, pitching moment.

Be air hydrodynamic, S _wBe wing area, L is the span, b _ABe mean aerodynamic chord.

In order to calculate three-axis force and the three-axis force square in (1-1) formula, calculate three-axis force coefficient and three-axis force moment coefficient earlier, the computing formula that has provided each coefficient below is as follows:

Resistance coefficient: $C_x=C_{x0}+\mathrm{\Δ}{C}_{x}H+{\mathrm{AC}}_y^2---(1-10)$

Lift coefficient: $C_y=c_{y0}+c_y^{\mathrm{\α}}*\mathrm{\α}+c_y^{{\mathrm{\δ}}_z}*{\mathrm{\δ}}_z---(1-11)$

Lateral force coefficient: $C_z=c_z^{\mathrm{\β}}*\mathrm{\β}+c_z^{\stackrel{\‾}{{\mathrm{\ω}}_y}}*\stackrel{\‾}{{\mathrm{\ω}}_y}+c_z^{{\mathrm{\δ}}_y}*{\mathrm{\δ}}_y---(1-12)$

The rolling moment coefficient: $m_x=m_x^{\mathrm{\β}}*\mathrm{\β}+m_x^{\stackrel{\‾}{{\mathrm{\ω}}_x}}*\stackrel{\‾}{{\mathrm{\ω}}_x}+m_x^{\stackrel{\‾}{{\mathrm{\ω}}_y}}*\stackrel{\‾}{{\mathrm{\ω}}_y}+m_x^{{\mathrm{\δ}}_x}*{\mathrm{\δ}}_x+m_x^{{\mathrm{\δ}}_y}*{\mathrm{\δ}}_y---(1-13)$

Yawing moment coefficient: $m_y=m_y^{\mathrm{\β}}*\mathrm{\β}+m_y^{\stackrel{\‾}{{\mathrm{\ω}}_x}}*\stackrel{\‾}{{\mathrm{\ω}}_x}+m_y^{\stackrel{\‾}{{\mathrm{\ω}}_y}}*\stackrel{\‾}{{\mathrm{\ω}}_y}+m_y^{{\mathrm{\δ}}_x}*{\mathrm{\δ}}_x+m_y^{{\mathrm{\δ}}_y}*{\mathrm{\δ}}_y---(1-14)$

Pitching moment coefficient: $m_z=m_{z0}+m_z^{{\mathrm{\δ}}_z}*{\mathrm{\δ}}_z+m_z^{\mathrm{\α}}*(\mathrm{\α}-{\mathrm{\α}}_0)+m_z^{\stackrel{\‾}{{\mathrm{\ω}}_z}}*\stackrel{\‾}{{\mathrm{\ω}}_z}+m_z^{\stackrel{\‾}{\mathrm{\α}}}*\stackrel{\‾}{\mathrm{\α}}---(1-15)$

In the following formula, the derivative of each component of resistance coefficient all is the function of Mach number, wherein C _X0, c _Y0Be zero resistance, lift coefficient, Δ C _xH is that the thunder of height is such counted correction, and A is the induced drag factor.

With Be two derivatives of lift coefficient, δ _x, δ _y, δ _zBe respectively aileron, direction, elevator angle.α is aircraft angle of attack, α ₀For zero-lift angle, β are yaw angle, δ _yThe rudder kick angle,

With

Derivative for resistance coefficient.

Be horizontal static-stability derivative,

Be the lift-over damping derivative,

Be the intersection dynamic derivative of yawing rotation,

Be the control derivative of aileron,

Be the cross control derivative,

Be course static-stability derivative,

Be the cross derivative of rolling movement,

Be the course damping derivative,

Be the cross control derivative,

It is the rudder control derivative.m _Z0Be the zero lift moment coefficient,

Aerodynamic derivative,

The aerodynamic derivative that causes for tailplane,

For aircraft around OZ _tAxle rotates and causes aerodynamic derivative,

Be zero dimension lift-over, driftage, rate of pitch.

For zero dimension change in angle of attack speed,

Be the dynamic derivative of taking offence.

Q, Y, C and M _x, M _y, M _zAll known situation gets off to resolve 12 differential equation of first orders in the aircraft flight process.

$\left\{\begin{array}{c}\mathrm{\Σ}{F}_x=m({\stackrel{\·}{V}}_x+V_z{\mathrm{\ω}}_y-V_y{\mathrm{\ω}}_z)\\ \mathrm{\Σ}{F}_y=m({\stackrel{\·}{V}}_y+V_x{\mathrm{\ω}}_z-V_z{\mathrm{\ω}}_x)\\ \mathrm{\Σ}{F}_z=m({\stackrel{\·}{V}}_z+V_y{\mathrm{\ω}}_x-V_x{\mathrm{\ω}}_y)\end{array}\right.---(1-16)$

$\left\{\begin{array}{c}\mathrm{\Σ}{M}_x=I_x{\stackrel{\·}{\mathrm{\ω}}}_x+(I_z-I_y){\mathrm{\ω}}_z{\mathrm{\ω}}_y+I_{\mathrm{xy}}({\mathrm{\ω}}_x{\mathrm{\ω}}_z-{\stackrel{\·}{\mathrm{\ω}}}_y)\\ \mathrm{\Σ}{M}_y=I_y{\stackrel{\·}{\mathrm{\ω}}}_y+(I_x-I_z){\mathrm{\ω}}_x{\mathrm{\ω}}_z-I_{\mathrm{xy}}({\mathrm{\ω}}_y{\mathrm{\ω}}_z-{\stackrel{\·}{\mathrm{\ω}}}_x)\\ \mathrm{\Σ}{M}_z=I_z{\stackrel{\·}{\mathrm{\ω}}}_z+(I_y-I_x){\mathrm{\ω}}_y{\mathrm{\ω}}_x-I_{\mathrm{xy}}({\mathrm{\ω}}_y^2-{\mathrm{\ω}}_x^2)\end{array}\right.---(1-17)$

$\left\{\begin{array}{c}\stackrel{\·}{\mathrm{\γ}}={\mathrm{\ω}}_x-\mathrm{tg\θ}({\mathrm{\ω}}_y\cos \mathrm{\γ}-{\mathrm{\ω}}_z\sin \mathrm{\γ})\\ \stackrel{\·}{\mathrm{\ψ}}=({\mathrm{\ω}}_y\cos \mathrm{\γ}-{\mathrm{\ω}}_z\sin \mathrm{\γ})/\cos \mathrm{\θ}\\ \stackrel{\·}{\mathrm{\θ}}={\mathrm{\ω}}_y\sin \mathrm{\γ}+{\mathrm{\ω}}_z\cos \mathrm{\γ}\end{array}\right.---(1-18)$

$\left\{\begin{array}{c}{\stackrel{\·}{x}}_d=V_x\cos \mathrm{\ψ}\cos \mathrm{\θ}+V_y(\sin \mathrm{\ψ}\sin \mathrm{\γ}-\cos \mathrm{\ψ}\sin \mathrm{\θ}\cos \mathrm{\γ})+V_z(\sin \mathrm{\ψ}\cos \mathrm{\γ}+\cos \mathrm{\ψ}\sin \mathrm{\θ}\sin \mathrm{\γ})\\ {\stackrel{\·}{y}}_d=V_x\sin \mathrm{\θ}+V_y\cos \mathrm{\θ}\cos \mathrm{\γ}-V_z\cos \mathrm{\θ}\sin \mathrm{\γ}\\ {\stackrel{\·}{z}}_d=-V_x\sin \mathrm{\ψ}\cos \mathrm{\θ}+V_y(\cos \mathrm{\ψ}\sin \mathrm{\γ}+\sin \mathrm{\ψ}\sin \mathrm{\θ}\cos \mathrm{\γ})+V_z(\cos \mathrm{\ψ}\cos \mathrm{\γ}-\sin \mathrm{\ψ}\sin \mathrm{\θ}\sin \mathrm{\γ})\end{array}\right.---(1-19)$

∑ F in the following formula _x, ∑ F _y, ∑ F _zBe respectively making a concerted effort on x, y, the z axle,

Be respectively speed differential component on x, y, the z axle.ω _x, ω _y, ω _zBe respectively angular velocity component on x, y, the z axle, other footmark implication therewith is identical.M is Aircraft Quality. γ, ψ are respectively pitching, roll and course angle.I _x, I _y, I _zBody is around the moment of inertia of each, I _XyBe the product of inertia.

Under the output rudder of 1.3 steering wheel models believed one side only breath and 1.5 thrust information functions, model aircraft was exported corresponding Flight Condition Data.

2, flight path setting, inertial navigation starting condition are aimed at, flight simulation process, the output of flight status data

2.1 flight path setting

It generally is to arrange according to the trajectory planning requirement that flight path arranges.Trajectory planning refers to seek movable body satisfies certain performance index optimum from the initial point to the impact point movement locus under specific constraint condition.For ease of saying something, among the design the flight path setting by different situation set that aircraft is initial, the position coordinates of state transition point and termination track points, the initial attitude, position, the velocity information that comprise aircraft simultaneously, each modal information that flies calculates judgement according to coordinate and the aircraft of adjacent track points in current position, and whether basis for estimation i.e. two track points has difference in height, longitude, difference of latitude to produce corresponding flight guidance and control to select suitable flight mode.

2.2 inertial navigation initial alignment

The initial information that namely guarantees inertial navigation is consistent with the original state that flight path is set, i.e. settings such as identical attitude, position, speed.

2.3 flight simulation

More than setting up, under 2.1, the 2.2 described prerequisites, begin the flight simulation process.

2.4 output Flight Condition Data

Satisfy and successful implementation communication condition, namely realized the correlation function of communication module.Guaranteed to carry out under the prerequisite of data communication protocol, the communication protocol between the computing machine as inertial navigation system and simulation computer the output of real-time Flight Condition Data.These data are divided into two branch roads to be exported simultaneously, and a branch road is directly exported to monitoring computing system (supervisory control comuter) by relevant communication protocol and the corresponding interface, and another branch road is exported to the anti-module of resolving, and its data flow is seen Fig. 2.

3, inertial navigation is counter resolves

Anti-specific force and the angular speed of calculating of Flight Condition Data by the output of flight control emulation module.

The longitude that provides by the flight simulation module, latitude and height (L, λ, H), and geographic coordinate system get off the plane east orientation, north orientation and day to speed (V _E, V _N, V _U) calculating inertial coordinates system, to tie up to navigation coordinate with respect to terrestrial coordinates be the angular speed projection Terrestrial coordinate system ties up to the angular speed projection down of navigation coordinate system with respect to navigation coordinate

Inertial coordinates system is with respect to navigation coordinate angular speed projection under navigation coordinate system

${\mathrm{\ω}}_{\mathrm{en}}^n={\left[\begin{array}{cccc}-\frac{v_N^n}{R_e}& \frac{v_E^n}{R_e}& \frac{v_E^n}{R_e}\tan & L\end{array}\right]}^T---(3-1)$

${\mathrm{\ω}}_{\mathrm{ie}}^n={\left[\begin{array}{ccc}0& {\mathrm{\ω}}_{\mathrm{ie}}\cos L& {\mathrm{\ω}}_{\mathrm{ie}}\sin L\end{array}\right]}^T---(3-2)$

${\mathrm{\ω}}_{\mathrm{in}}^n={\mathrm{\ω}}_{\mathrm{ie}}^n+{\mathrm{\ω}}_{\mathrm{en}}^n---(3-3)$

R in the following formula _eBe earth radius,

For navigation coordinate is east orientation, north orientation speed, L is latitude, ω _IeBe the spin velocity of the earth under inertial coordinates system.

Aircraft pitch, roll and the course angle that provides by the flight simulation module (

γ, ψ) calculate angular velocity under the body axis system according to the Eulerian angle differential equation

Inertial navigation attitude transition matrix

$\left[\begin{array}{c}{\mathrm{\ω}}_{\mathrm{nbx}}^b\\ {\mathrm{\ω}}_{\mathrm{nby}}^b\\ {\mathrm{\ω}}_{\mathrm{nbz}}^b\end{array}\right]=C_{\mathrm{\γ}}{C}_{\mathrm{\θ}}\left[\begin{array}{c}0\\ 0\\ -\stackrel{\·}{\mathrm{\ψ}}\end{array}\right]+C_{\mathrm{\γ}}\left[\begin{array}{c}\stackrel{\·}{\mathrm{\θ}}\\ 0\\ 0\end{array}\right]+\left[\begin{array}{c}0\\ \stackrel{\·}{\mathrm{\γ}}\\ 0\end{array}\right]=>\left[\begin{array}{c}{\mathrm{\ω}}_{\mathrm{nbx}}^b\\ {\mathrm{\ω}}_{\mathrm{nby}}^b\\ {\mathrm{\ω}}_{\mathrm{nbz}}^b\end{array}\right]=\left[\begin{array}{ccc}\cos \mathrm{\θ}\sin \mathrm{\γ}& \cos \mathrm{\γ}& 0\\ -\sin \mathrm{\θ}& 0& 1\\ -\cos \mathrm{\θ}\cos \mathrm{\γ}& \sin \mathrm{\γ}& 0\end{array}\right]\left[\begin{array}{c}\stackrel{\·}{\mathrm{\ψ}}\\ \stackrel{\·}{\mathrm{\θ}}\\ \stackrel{\·}{\mathrm{\γ}}\end{array}\right]---(3-4)$

$C_b^n=\left[\begin{array}{ccc}\cos \mathrm{\γ}\cos \mathrm{\ψ}+\sin \mathrm{\γ}\sin \mathrm{\ψ}\sin \mathrm{\θ}& \sin \mathrm{\ψ}\cos \mathrm{\θ}& \sin \mathrm{\γ}\cos \mathrm{\ψ}-\cos \mathrm{\γ}\sin \mathrm{\ψ}\sin \mathrm{\θ}\\ -\cos \mathrm{\γ}\sin \mathrm{\ψ}+\sin \mathrm{\γ}\cos \mathrm{\ψ}\sin \mathrm{\θ}& \cos \mathrm{\ψ}\cos \mathrm{\θ}& -\sin \mathrm{\γ}\sin \mathrm{\ψ}-\cos \mathrm{\γ}\cos \mathrm{\ψ}\sin \mathrm{\θ}\\ -\sin \mathrm{\γ}\cos \mathrm{\θ}& \sin \mathrm{\θ}& \cos \mathrm{\γ}\cos \mathrm{\θ}\end{array}\right]---(3-5)$

In the following formula

Be respectively

At three direction of principal axis components, C _γ,

Be respectively lift-over, pitching transition matrix,

Be respectively the angular velocity of pitching, driftage, roll angle.Utilize

Calculate angular speed under the gyro angular speed output body axis system

${\mathrm{\ω}}_{\mathrm{ib}}^b={\mathrm{\ω}}_{\mathrm{nb}}^b+{\left(C_b^n\right)}^T{\mathrm{\ω}}_{\mathrm{in}}^n---(3-6)$

The proportion of utilization equation calculates the specific force f under acceleration and the output body coordinate ^b

$f^n={\stackrel{\·}{v}}^n-(2{\mathrm{\ω}}_{\mathrm{ie}}^n+{\mathrm{\ω}}_{\mathrm{en}}^n)\×v^n-g^n---(3-7)$

$f^b={\left(C_b^n\right)}^T{f}^n---(3-8)$

V in the following formula ⁿThe following speed of navigation coordinate system,

Be speed differential, g under the navigation coordinate system ⁿBe acceleration of gravity, f ⁿFor navigation coordinate is specific force.

The computational algorithm of attitude transition matrix also can the service orientation Method of Cosine, hypercomplex number method, equivalent rotating vector method etc.

4, inertial navigation system calculates Flight Condition Data

4.1 specific force, angular rate data are transferred to inertial navigation

According to relevant communication protocol, as realization related communication functions of modules such as ARINC429, specific force, angular rate data that above-mentioned 3 trifles are calculated are transferred to actual inertial navigation system.

4.2 inertial navigation system calculates the flight status data

Inertial navigation system calculates other one group of Flight Condition Data according to performance own.

4.3 the output of Flight Condition Data

According to relevant communication protocol, realize the communication module function, the Flight Condition Data that inertial navigation system is produced is transferred to Monitor And Control Subsystem (supervisory control comuter).

5, the comparative analysis of data

In Monitor And Control Subsystem, realize the related communication function, namely finish relevant data transmission communication function respectively, can receive the Flight Condition Data that is produced by inertial navigation system and excitation subsystem respectively, respectively being compared by flight control analogue system Flight Condition Data that directly export and that exported by inertial navigation system, as data such as speed, attitude, positions.During contrast according to the time clap the principle of correspondence, namely satisfying under 2.2 conditions, according to corresponding each time data of clapping be analyzed.Thereby every performances such as the error of each data item of check inertial navigation system, variance.

Claims (1)

1. general inertial navigation method of testing based on flight simulation is characterized in that may further comprise the steps:

The first step, foundation is based on the flight control emulation module of control law design, for the inertial navigation system performance test provides the data driving source, utilization is based on the design of aircraft control law, carrying out nonlinear equation by known model aircraft parameter derives and calculating, and carry out the vertical and side direction control law design of aircraft, realize the motion under the aircraft typical Mode;

Second step, carry out flight path setting, inertial navigation starting condition aligning, implement the flight simulation process, and output flight status data, specifically may further comprise the steps:

1, carry out flight path and set, flight path comprises aircraft initial state information and other track points positional information, and flight path is set to contain in the air and climbed, glides, spirals, turns and comprehensive non-stop flight mode;

2, inertial navigation initial alignment, the starting condition data are consistent during the Flight Condition Data under the inertial navigation starting condition and flight path setting;

3, real-time resolving and output are by the Flight Condition Data of the aircraft of flight control emulation module output;

The 3rd step, by anti-specific force and the angular rate data of calculating of Flight Condition Data, the longitude that provides by the flight simulation module, latitude and height, and geographic coordinate system get off the plane east orientation, north orientation and day calculate inertial coordinates system to velograph to tie up to navigation coordinate with respect to terrestrial coordinates be that angular speed projection, terrestrial coordinate system tie up to the angular speed projection down of navigation coordinate system, inertial coordinates system with respect to navigation coordinate angular speed projection under navigation coordinate system with respect to navigation coordinate

Aircraft pitch, roll and course angle by the flight simulation module provides calculate angular velocity under the body axis system according to the Eulerian angle differential equation

Attitude matrix

Utilize

Calculate angular speed under the gyro angular speed output body axis system; Utilize the specific force equation to calculate specific force under acceleration and the output body coordinate;

The 4th step, inertial navigation system goes out the flight status data according to specific force and angular speed real-time resolving, specific force and angular rate data are transferred to the inertial navigation system that will test by communication line, and inertial navigation system resolves the back and by communication line Flight Condition Data is exported;

The 5th step, according to the time clap the consistent Flight Condition Data of flight control emulation module output and the Flight Condition Data that inertial navigation system is exported of contrast, compare data differences information between the two, thereby the inertial navigation performance detected.

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