Patent Application
20170139425
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The technology herein relates to a flight control protection methodology for an aircraft. More particularly, example techniques herein limit the wing lift coefficient as a function of dynamic pressure (or velocity) to create a limitation for the maximum lift produced by the wing.
Commonly-assigned U.S. Pat. No. 8,214,089 (incorporated herein by reference) discloses a flight control system that moves control surfaces such as elevators according to a pilot command summed with an automatic command. The flight control system monitors a set of flight parameters such as angle of attack (AOA) to determine if the flight vehicle is operating inside a permitted envelope. The flight control system incorporates automatic protections through automatic commands if the flight vehicle is close to its envelope limits.
While the techniques disclosed in that patent are highly useful, further improvements are possible and desirable.
The following detailed description of exemplary non-limiting illustrative embodiments is to be read in conjunction with the drawings, of which:
The example non-limiting methodology and apparatus modifies conventional angle of attack (AOA) envelope protection such as disclosed in U.S. Pat. No. 8,214,089 to perform additional structural protection, such as to protect maximum lift.
The example non-limiting embodiments build the AOA_max tables as a function of a set of parameters that comprehend dynamic pressure. The dynamic pressure can be calculated by using different combinations of parameters, for example: Qdyn (dynamic pressure), KCAS (calibrated airspeed in knots), KEAS (effective air speed in knots), Mach and Altitude.
In the definition of the operational envelope of cargo aircraft, it is common to limit the maximum load factor as a function of the aircraft weight in order to optimize the aircraft performance for different mission objectives.
The usual process to limit the load factor is based on the use of mass estimators. As shown in
1. Fly-By-Wire (FBW) aircraft (closed loop): The flight control computer implements a control law (12) that limits the maximum load factor available based on the mass.
2. Conventional aircraft (open loop): The system informs the pilot (14) of the allowable load factor. The pilot is then responsible for maneuvering the aircraft without exceeding the informed limits.
Both approaches result in a limited load factor 14 so the aircraft always operates within the load factor limitation the weight and CG envelope E (
However, the particular scheme shown in
In the prior art, it is known to calculate the mass by estimating the lift coefficient CL (that is interpolated using an aerodynamic database of the aircraft), based on flight conditions according to equations 1 and 2 below:
CL=f (alpha, flap, Mach, elevator, Stabilizer position) (equation 1)
Estimated mass=Qdyn*S*CL/(nz*g) (equation 2)
Where:
Qdyn=Dynamic pressure
S=Aircraft reference wing area
nz=aircraft longitudinal load factor
g=gravity.
A problem of the estimation process can result from uncertainty in the aerodynamic database and imprecision of angle of attack measurements. Mass estimators 10 used in civil aircrafts can present perceptual errors on the order of 10%. When applying this technology to cargo planes, the scheme becomes more complex when needed to cover sensitive missions such as a cargo drops or firefighting.
It would be desirable to overcome difficulties in estimating aircraft mass by providing an alternative method for developing a load limiting factor such as shown in
The example non-limiting technology herein presents functionality that limits the maximum lift force that can be produced by the aircraft. This functionality limits the maximum load factor achieved by the aircraft as a function of weight (without the need to estimate it) respecting the maximum lift force that designs the wing structure. The functionality also allows adds other features to reduce the loads more effectively, since the controlled parameter represents more faithfully the phenomena that design the structures.
The present methodology is applicable for example in cases where the structural limitation is mainly governed by the maximum lift force produced by an aircraft wing. Example non-limiting methods and apparatus herein limit the wing lift coefficient as a function of dynamic pressure (or velocity) to create a limitation for the maximum lift produced by the wing.
The Mach effect on maximum lift coefficient also can be observed on
In a conventional angle of attack protection, the reference maximum values allowed by the control law are calculated based on aircraft configuration (such as flap/slat deflections, landing gear position . . . ) and a set of parameters related to flight condition (such as Mach number, ice condition . . . ). Both configuration and flight condition parameters are selected based on relevance to the aerodynamic stall phenomena.
In one example non-limiting embodiment, the lift coefficient limitation can be implemented through angle of attack protection where the reference maximum values used by the functionality are additionally limited by a parameter related to the dynamic pressure, such as equivalent airspeed, calibrated airspeed, dynamic pressure and other combinations.
Generally, the disclosed protection system embodiment can be implemented to limit other load parameters relevant to the structural design, such as wing lift and wing bending moment. Additionally, other relevant parameters such as wing fuel weight can be used to improve system performance. The wing fuel weight is particularly relevant if the wing bending moment is the parameter to be limited.
The structure of one example non-limiting implementation of the methodology is similar to a stall protection. A difference consists in using any combination of parameters that permits calculation of the dynamic pressure of the flight condition in order to protect maximum lift (Lmax).
In more detail,
The example non-limiting embodiments thus use a mass estimator that can estimate CL under different flight conditions. Because of this, the mass estimator is exposed to aerodynamic database errors in all these conditions. The Lmax protection thus depends on the accuracy of the database in a smaller envelope and can be calibrated with flight test data. Example: Building the alpha (AOA) vs KEAS (equivalent airspeed) boundary by making windup turns thereby sweeping speeds of interest for three different altitudes.
In some example non-limiting embodiments, Lmax protection is subject to errors in alpha readings at high alpha conditions where the relative error becomes lower. Additionally, the lift curve slope usually decreases at high angle of attack; this is reduced which also contributes to decreased effect of errors in the alpha readings. The weight estimator process needs to deal with conditions where the angle of attack is small, which makes the alpha sensor readings relatively greater. These effects can be seen for example in
Example Non-Limiting Implementation
The exemplary illustrative non-limiting implementations herein relate to systems, apparatuses and methods to be used in a flight vehicle equipped with pitch control, such as elevators and a pilot inceptor such as a side-stick or a column yoke.
An exemplary illustrative non-limiting flight control system is shown in
Further, the exemplary illustrative non-limiting system receives signals from sensors 218, 219, 220, 221, 222. In this exemplary implementation, the sensors provide: angle-of-attack (α), angle-of-attack rate ({dot over (α)}) airspeed (u), airspeed rate ({dot over (u)}), the flap position (δF), gear position (δG), pitch attitude (θ), pitch rate (q), height above ground (hAGL) ice detection bit (bIce), engine throttle lever position (δTLA), Mach number (Mach) and altitude (h). Other sensors are also possible. For example, it is possible to measure barometric air pressure using a pitot tube or other sensor to determine dynamic air pressure as discussed above.
According to this exemplary implementation, the information flows via a means of transmitting multiple data such as a bus 205. All the data, i.e. pilot commands and sensors, is sent to a processor 204 that is operable to compute output based, for example, on a programmable code. The processor 204 is able, for example, to compute an elevator command based on the input data received.
This command is sent to a mechanism to actuate a flight control surface 207 to control or limit lift. Resultantly, the control surfaces are deployed according to the command computed by the processor 204.
In the exemplary illustrative non-limiting implementation, a pilot inceptor command is transformed into an alpha (α) command and/or pitch angle (θ) command when protections are active. The relation between the variable to be controlled (α or θ) and pilot command is depicted as command shaping 308. The output of the command shaping (δlaw) is used as a reference to manipulate the elevators to track the variables α or θ. When the pilot moves the inceptor to the stop (i.e., the mechanical limit of the inceptor), command shaping produces a maximum α or θ in order to preclude the airplane from exceeding the maximum allowed α or θ for the current airplane configuration.
The control law is calculated using the pitch states of the aircraft dynamic 307 which are fed back to the closed loop control law. Airspeed (u), pitch rate (q), pitch angle (θ) and angle of attack (α) are multiplied by the gains listed as 301, 302, 303, 304, respectively. The feed-forward command is produced based on the feed-forward gain 309 multiplied by the reference generated by the command shaping output 308.
The error (e) is calculated as the result of the difference between the reference and the angle-of-attack or pitch angle. The angle-of-attack is used when the stall, low speed and/or buffeting protections are engaged. The pitch angle (θ) is used when the high attitude protection is engaged. The integral of the error (e) is multiplied by the integral gain in order to produce the integral command.
The gains values depend upon which protection is active. For example, when the low speed protection is active, the pitch angle gain 303 and true airspeed gain 301 are increased when compared to the pitch angle gain 303 and true airspeed gain 301 used in the stall protection function. Also, the gains are scheduled according to the Mach number and altitude the airplane is flying at the moment the protection is engaged.
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiments, it is to be understood that the invention is not to be limited to the disclosed embodiments, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.
