Classifications

• • G—PHYSICS
  • G05—CONTROLLING; REGULATING
  • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
  • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
  • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
  • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
  • G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
  • G05D1/0825—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using mathematical models

Definitions

• the invention relates to a calculation system for an aircraft and to a method for determining characteristics and motion variables of an aircraft.
• Aircraft such as airplanes or helicopters
• Aircraft are exposed to various forces during their flight.
• Significant influencing factors are the buoyancy forces generated by the wings, the aerodynamic resistance of the aircraft, the weight force or gravity acting on a center of gravity of the aircraft, the thrust generated by the engines, the control forces generated on the control surfaces of the aircraft and those caused by the aircraft respective forces caused torques.
• the inertia of the aircraft or the mass inertia of the aircraft components also play a role in the abovementioned forces. Flight maneuvers and air turbulence cause structural loads on the aircraft.
• the invention provides an aircraft calculation system having at least one sensor for detecting aeroelastic and flight mechanic motion quantities of the aircraft, for detecting positions and movements of control surfaces of the aircraft or for detecting speeds of wind gusts acting on the aircraft; and with a calculation unit which calculates parameters of passenger comfort and cabin safety as well as movement quantities of the aircraft as a function of the sensor data output by the sensors and a non-linear simulation model of the aircraft.
• the calculation unit automatically adapts the non-linear simulation model based on the sensor data output by the sensors.
• sensors are provided for detecting movement variables of an on-board system of the aircraft.
• the on-board system has at least one movable mass for damping an associated aircraft part of the aircraft.
• sensors for detecting flight mechanical movement variables of the aircraft also measure deformations of aircraft parts of the aircraft.
• the sensors for detecting flight mechanical movement variables of the aircraft and for detecting aeroelastic movement variables of the aircraft acceleration or pressure sensors.
• the calculation unit is provided in the aircraft or the sensor data from the sensors of the aircraft are received by the calculation unit via a wireless radio interface.
• the linear simulation model of the aircraft can be read from a memory.
• the calculation unit is connected to an input unit for input of parameters of the simulation model of the aircraft.
• the calculation unit is connected to an output unit for outputting the characteristic and movement variables.
• the on-board system of the aircraft is automatically controlled in dependence on the parameters and movement variables calculated by the calculation unit for minimizing load forces and vibrations. In one embodiment of the calculation system according to the invention, the on-board system of the aircraft can be switched on or off for different frequency ranges.
• the invention further provides a method for determining parameters of passenger comfort and movement quantities of an aircraft, having the following steps:
• the invention further provides a computer program with program instructions for carrying out a method for determining parameters of passenger comfort and movement variables of an aircraft, having the following steps:
• the invention further provides a data carrier which stores such a computer program.
• FIG. 1 shows coordinate systems of a non-linear simulation model of an aircraft used in the inventive calculation system
• FIG. 2 shows a block diagram of a possible embodiment of the calculation system according to the invention
• FIG. 3 shows a block diagram of a further embodiment of the calculation system according to the invention.
• FIGS. 4A, 4B are diagrams for explaining the non-linear simulation model of an aircraft on which the calculation system according to the invention is based;
• FIGS. 5A-5G show special cases of the non-linear simulation model on which the calculation system according to the invention is based; characters
• 6A, 6B are diagrams illustrating application examples of the aircraft calculating system according to the invention.
• 7A, 7B are diagrams for illustrating further examples of application of the inventive calculation system for aircraft.
• Flight mechanics describes the behavior of an aircraft moving in the atmosphere using aerodynamics.
• the flight mechanics describes the behavior of the entire system or the aircraft, wherein a position, attitude and flight speed of a missile is calculated at any time. This is done with the help of equations of motion, which form a system of equations of coupled differential equations. Due to maneuvers and air turbulence maneuver loads and structural loads occur on an aircraft. Maneuver loads can be described using non-linear equations of motion and are based on databases that specify aerodynamic forces. In particular, large aircraft must take into account not only non-linear movements but also the elastic deformations of their structure.
• the specific force is a measure of the pilot's sense of acceleration in terms of magnitude and direction and is defined as the ratio of the resulting external force to the aircraft mass.
• the accelerations and velocities are measured with respect to an inertial system.
• the earth serves as an inertial system, whereby an earth-fixed coordinate system F E is defined, in which the z-axis points in the direction of the center of the earth.
• the x- and y-axes are selected such that a legal system is created.
• the axbox can z. B. are aligned to the magnetic north point.
• an aerodynamic coordinate system F A is chosen whose origin also lies in the center of gravity C of the aircraft.
• the x-axis of this coordinate system lies in the direction of the negative flow velocity, the z-axis in the direction of the negative buoyancy.
• the y-axis is chosen analogously to the previous considerations.
• This coordinate system is obtained by rotating the body-fixed main axis system by an angle of attack ⁇ about its y-axis and then by a sliding angle ⁇ about the z-axis.
• the aerodynamic coordinate system F A is body-fixed only in stationary flight states of the aircraft. The transition from the body-fixed to the earth-fixed coordinate system is done by means of a transformation matrix L EB
• the subscript indicates the coordinate system in which the vectors are displayed.
• the index B is omitted, if not absolutely necessary. In the speed must be additionally distinguished between wind and calm. In general, with the Speed Addition Law:
• the superscript specifies the frame of reference by which the respective velocities are measured.
• W E is the wind speed that can be assumed to be zero. Thus, the amounts in both frames are the same and the superscript can be omitted.
• the resulting force F is composed of the aerodynamic force R and the weight. This Be Draws are used in the above equation and then resolved to V.
• the movement of the rigid body is influenced by the elastic deformations.
• the dynamics of the elastic degrees of freedom should be considered in the equations of motion.
• the deformation of the structure can be approximately described by superposition of normal modes of free vibration:
• x ', y', z ' are the deflections from the respective rest positions Xo, Yo, Z 0 ; f n, g n and h n the mode shaping functions and ⁇ n generalized coordinates.
• the additional equations of motion for the mode ⁇ n are obtained from the equation of Lagrange as equations of forced vibrations.
• the natural frequency ⁇ n of the damping d n and the generalized moment of inertia I n are approximately equivalent
• equation (24) can be written in the following form.
• equation (24) can be formulated for all modes.
• the external forces acting on an aircraft are weight, aerodynamic forces, buoyancy, resistance and thrust.
• the point of attack of the lift is in the so-called neutral point, which is different from the center of gravity. This creates moments. The same applies to the thrust.
• the resulting forces are summarized in a vector R, the moments in a vector Q.
• Buoyancy and resistance are generated by the relative movement of aircraft and air, ie V and ⁇ . These forces also depend on the angle of attack a and the angles of the control surfaces of the primary flight control, altitude ( ⁇ E ), lateral ( ⁇ A ) and rudder ( ⁇ R ).
• additional control surfaces, spoilers, spoilers, canards are used, which are referred to below as ⁇ c .
• the angles of the control surfaces are combined together with the thrust ⁇ F in a control vector c.
• the aerodynamic effects are based on nonlinear relationships. They can be described by Taylor series, which are broken off after a certain order.
• the coefficients of the second and third order terms are one to two orders of magnitude below the first order coefficients. If the angle of attack remains below 10 °, the terms of higher order can be neglected.
• the starting point of the linear approach is a stationary flight state. The speeds and rates as well as forces and moments are split into a stationary and a fault term:
• the horizontal symmetrical straight flight can be selected.
• the quantities indicated by u and ⁇ in equation (31) describe the influence of the elastic modes on the aerodynamics. They are each vectors of length k, where k is the number of elastic modes.
• the c indexed derivatives are also vectors that describe the influence of the control variables. Its dimension is equal to the number of control variables.
• ⁇ _ and ⁇ denote the introduced elastic ⁇ modes, while the control variables contained in the vector c are.
• the symmetrical horizontal straight-ahead flight is also assumed here. All error terms are assumed to be sufficiently small, so that the linear approximation for the aerodynamics is valid. Furthermore, it is neglected. Under these Prerequisites can be written the equations of motion in the following form:
• Equations (33) and (34) are denoted by the following abbreviations:
• the matrices A 13 and B 1 are obtained by respectively replacing the index ⁇ in the matrix A 12 by u and c, respectively.
• H_ and h (x 1 ) are:
• equation (33) described non-linear simulation model contains a ⁇ efficacy matrix F, which takes into account the non-linear characteristics of system ⁇ sizes.
• the effectiveness matrix F is given in equation (42). If one extends the model to aerodynamic, Appeldyna ⁇ mix and AeroLas ti see non-linear activities i arising
• the quantities X NL , W , Z NL , W , Y NL , W , DNL, I and D NL , 2 describe the influence strength of the nonlinearity.
• Equation (33) The nonlinear simulation model presented in Equation (33) can be described more physically (in generalization of Newton's and Euler's equations of motion) as follows:
• the equation system is shown clearly in the diagram of FIG. 4A.
• the equation system illustrated in FIG. 4A comprises a dynamic model of linear differential equations which is extended by an efficiency matrix F which is multiplied by a nonlinearity vector g.
• the vector x forms a hyper-motion vector of the aircraft.
• further non-linear extensions can be represented very clearly.
• nonlinearities in engine dynamics in system behavior or in error cases extend the non-linearity vector g ⁇ x, x, p, t) and the efficiency matrix F in additional entries.
• the matrix entries of F describe the influence strength of nonlinearities, as "effective force or moment" in the generalized Newton “see and Euler” see equations of motion.
• the mass matrix M, the damping matrix D and the stiffness matrix K are extended matrices that take into account the aerodynamics.
• Fig. 4B illustrates the structure of such an extended matrix.
• the coupling describes the influence of a parameter on the aircraft.
• the mass matrix M, the damping matrix D and the stiffness matrix K describe linear influences, while the effectiveness matrix F describes non-linear properties of system variables.
• These parameters are flight mechanics parameters, parameters of the on-board system and parameters of aeroelasticity.
• FIGS. 5A, 5B, 5C, 5D, 5E show special cases of the general nonlinear simulation model illustrated in FIG. 4A.
• the non-linear efficiency matrix as well as the nonlinearity vector g and the input variable vector p are zero. In this way one arrives at the special case of the purely linear equation system of differential equations.
• the non-linear efficiency matrix and the nonlinearity vector g are zero, while the input variable vector p is not zero, for example, for representing a gust of wind.
• the simulation model shown in FIG. 5B is thus suitable, for example, for analyzing wind gusts acting on the aircraft.
• the integral model is suitable for analysis of non-linear gusts, safety and passenger comfort.
• the integral simulation model is suitable for analyzing the system dynamics of the on-board system.
• FIG. 2 shows an exemplary embodiment of a calculation system 1 according to the invention for an aircraft 2, for example for an aircraft.
• sensors 3 are provided at the aircraft or aircraft 2 .
• the sensors 3 are used to detect aeroelastic and flight mechanic movement magnitudes of the aircraft 2.
• sensors for detecting positions and movements of control surfaces of the aircraft 2 and for detecting speeds of wind gusts acting on the aircraft 2 are provided.
• the sensors 3 thus form control surface sensors, flymechani- see sensors and aeroelastic sensors.
• the sensors 3 for detecting flight mechanical movement variables of the aircraft 2 and for detecting aeroelastic movement variables of the aircraft have, for example, acceleration and pressure sensors.
• the sensors 3 for acquiring aircraft mechanical movement variables can also measure deformations of aircraft parts of the aircraft 2.
• the calculation system 1 contains a calculation unit 4 which calculates parameters of passenger comfort and cabin safety as well as movement quantities of the aircraft 2 as a function of the sensor data output by the sensors 3 and a non-linear simulation model of the aircraft 2.
• This non-linear simulation model is read from a memory 5 in the embodiment shown in FIG.
• the calculation unit 4 has at least one microprocessor for executing a simulation software for the integral simulation model.
• the non-linear simulation model is automatically adapted based on the sensor data output by the sensors 3 and written back into the memory 5.
• the calculation unit 4 is located in the aircraft 2 and receives the sensor data via an internal data bus from the sensors 3.
• the calculation unit 4 is not located in the aircraft 2, but receives the sensor data via a wireless Radio interface of the sensors 3. The calculation unit 4 can then be located, for example, in a ground station.
• the calculation system 1 furthermore contains an input unit 6 for inputting parameters of the simulation model for the aircraft 2.
• the characteristic and motion variables calculated by the calculation unit 4 are shown in FIG Embodiment output via an output unit 7.
• the output unit 7 is, for example, a display or a display.
• the input unit 6 is, for example, a keyboard for inputting data.
• the input unit 6 and the output unit 7 together form a user interface. This user interface may be used, for example, by an aircraft design optimization engineer.
• the parameters and movement quantities calculated by the calculation unit 4 are fed back to a comparison unit 8, in which a deviation between predicted values and values determined from the tests is calculated.
• the predicted quantities can be input via a control unit 9, for example, which are compared with the simulated parameters.
• the aircraft 2 is then controlled as a function of the difference or the deviation between the predicted and the simulated parameters.
• the calculation system 1 allows the integral dynamic calculation of loads, movement quantities of the roelastik, the flight mechanics and thus makes it possible for flight-telemetry and development engineers to use the sensor data to determine the time histories of all loads, aeroelastic and flight-mechanical motion variables occurring at the aircraft 2, and to compare them with the measured sensor data. This allows a targeted aircraft design optimization.
• the calculation system 1 according to the invention is suitable for carrying out targeted pilot training for this purpose.
• a pilot may compare control inputs and the resulting comfort, safety and load characteristics of the aircraft 2.
• line, flight test and simulator pilots can avoid peak loads in dangerous flight situations or maneuvers, reduce fatigue loads, avoid vibration-critical conditions, reduce high accelerations throughout the cabin area of the aircraft, increase passenger and crew safety, and to be trained in comfort.
• FIG. 3 shows a further exemplary embodiment of the inventive calculation system 1 for an aircraft 2.
• the aircraft 2 has a so-called on-board system 10, preferably different modes of operation being adjustable.
• the on-board system of the aircraft 2 is automatically controlled in response to the characteristic and motion quantities calculated by the calculation unit 4 to minimize load forces and vibrations.
• the on-board System 10 of the aircraft 2 for different frequency ranges on or off.
• the on-board system 10 has various masses attached to aircraft parts, which can be activated depending on the mode of operation of the on-board system 10.
• the on-board system 10 is used to improve comfort and cabin safety as well as to reduce loads on parts of the aircraft 2.
• sensors 3 detect the time evolution of loads and movement quantities of the aeroelastic, the flight mechanics and the on-board systems of the aircraft 2 as well as the control surface inputs and the effect of wind gusts the aircraft 2.
• the calculation unit 4 which may be, for example, a computer, calculated by means of a simulation software and the imported simulation model characteristics of passenger comfort and motion of the aircraft 2.
• On the calculation unit 4 also runs an input and output software for the input of Parameters of the simulation model as well as the output of the calculated quantities.
• the integrated design of the calculation system 1 with the on-board system 10 leads to an improvement in passenger comfort, the safety of the cab, the aeroelastic and vibration properties and the reduction of loads.
• an identification software it is possible to physically and physically identify partial and overall models of a high-dimensional parameter space, that is the submodels, the aerodynamics of the structure, etc., using the available sensor data.
• the simulation software and the identification software are integrated together with the input software for the sensor data and the input or output software for the user interface in a software system.
• FIGS. 6A, 6B show diagrams for application examples of the calculation system 1 according to the invention.
• tion of the calculation system a fully measured transfer function of an aileron on the lateral load factor on the front fuselage of an aircraft.
• the transfer function I in Fig. 6A shows the case that no onboard system 10 is used to improve passenger comfort.
• the on-board system 10 is turned on.
• the on-board system 10 which is turned on in response to the operation selection signal thus increases the aeroelastic damping in a frequency range of 2 to 3 Hz as shown in Fig. 6A. However, this improves the vibration characteristics, ease of passage and safety and reduces direct hull loads due to hull movement.
• FIG. 6B shows by way of example the transfer functions of an aileron with respect to a lateral load factor at the front fuselage of an aircraft 2 determined by the calculation system 1 according to the invention.
• the lateral load factor describes the load on the front fuselage, the comfort and the crew or passenger safety in the case of wind gust or extreme Flight maneuvers as well as aeroelastic and vibration properties.
• the transfer function III in FIG. 6B shows the case that no on-board system 10 is used for improving passenger comfort and safety for reducing the loads.
• the on-board system 10 is turned on.
• FIGS. 7A, 7B show diagrams for further application examples for explaining the calculation system 1 according to the invention and the method according to the invention for determining parameters of passenger comfort and movement variables of an aircraft 2.
• FIG. 7A shows the time evolution of a parameter, for example a load in the form of a scaled bending moment, on an outer wing of an aircraft 2 in a spiral-shaped turn.
• a vertical load factor NZ is on A focus of the aircraft 2 of Ig increased to 1.5g.
• the course I in FIG. 7A shows the time course according to a conventional integral simulation model without the use of the sensor-based calculation system 1 according to the invention.
• the curve II in FIG. 7A shows a time characteristic for a simulation model using the sensor-based calculation system 1 according to the invention.
• the curve III in FIG 7A shows an actually measured load on an outer wing of the aircraft 2 for validation.
• the real measured curve III is reproduced very well, that is to say the simulated curve almost corresponds completely the actual measured course.
• the load factor is a parameter which also describes the comfort, the aerodynamics and the safety of the aircraft 2.
• the load factor characterizes the acceleration at the center of gravity of the aircraft 2 again. It can also be seen from FIG. 7B that the course calculated using the calculation system 1 according to the invention agrees well with the actual measured course.
• the non-linear simulation model is automatically adapted on the basis of an initial model which corresponds, for example, to the profile I in FIGS. 7A, 7B, on the basis of the sensor data.
• This adaptation can be carried out iteratively in one possible embodiment.
• the adaptation or validation of the non-linear simulation model to the sensor data supplied by the sensors 3 takes place by means of a least square algorithm (LSA).
• LSA least square algorithm
• a holistic optimization of various parameters can be achieved.
• an engineer can simultaneously optimize passenger comfort parameters and aeroelastic characteristics taking into account the loads of the on-board system 10 and the flight mechanics.
• the acceleration forces acting on the passenger seats can be minimized, while at the same time aeroelastic parameters for minimizing material wear and for maximizing flight safety are calculated.
• the invention thus provides an integral sensor-based calculation system 1 for loads, parameters of passenger comfort and cabin safety as well as for movement variables of aerodynamics, structural dynamics, stationary and transient aerodynamics and on-board systems of aircraft 2.

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