DE69701589T2 - Landing position modifier for an aircraft - Google Patents

Description

Field of the invention

The invention relates to aircraft control systems, and more particularly to systems for controlling movable aircraft surfaces and/or high-lift surfaces of an aircraft. In particular, the invention relates to a method for modifying the attitude of an aircraft during landing, comprising the steps described in the preamble part of claim 1. Such a method is known from US-A-5 446 666.

Background of the invention

The pitch attitude of an aircraft is the angle subtended by the aircraft's longitudinal axis and a horizontal line. Pitch attitude is an important aircraft parameter that must be considered during approach and touchdown. The aircraft's pitch attitude during landing varies according to a number of factors. For example, atmospheric disturbances, such as the presence of a wind gust, affect the pitch attitude so that adjustment is required. Pitch attitude adjustment in response to atmospheric disturbances ensures that the desired flight path angle is maintained during landing.

During landing, the pitch attitude envelope is partially limited by the shape of the aircraft. During the flare and touchdown, as the nose of the aircraft rises to increase the pitch attitude, the tail section of the aircraft comes progressively closer to the ground. The distance between the tail section of the aircraft body and the ground is known as the tail body contact margin. To minimize the possibility of the tail section contacting the ground, it is desirable that the pitch attitude not exceed a maximum value, generally referred to as the most critical tail body flight condition. The most critical tail body flight condition defines a predetermined maximum pitch attitude that provides a required clearance between the tail section of the aircraft body and the ground during landing. The required clearance is known as the tail body contact margin requirement. This requirement is given in degrees and is defined by the angle between the runway and a line extending between the main landing gear of an aircraft and the lowest portion of the tail section of the aircraft. See Fig. 1.

In addition to a maximum value, it is also desirable during landing that the pitch attitude not exceed a minimum value. During approach, the aircraft's main gear and nose gear are extended. Since the nose gear is not designed to withstand the shock of aircraft touchdown, it is critical to land the aircraft so that the main gear contacts the ground first. The distance between the nose gear and the ground when the main gear contacts the ground is known as the nose gear contact margin. To avoid nose gear touchdown first landings, a minimum pitch attitude is defined, commonly referred to as a most critical nose gear flight condition. Landing with pitch attitudes equal to or greater than the minimum value for the most critical nose gear flight condition ensures that main gear touchdown precedes nose gear touchdown by a predetermined amount. The predetermined value is known as the nose gear contact span requirement. This requirement Obstacle is also expressed in degrees and is defined by the angle between the runway and a line extending between the main gear and the nose gear on the aircraft. See Fig. 2.

As stated above, the pitch attitude of an aircraft varies during landing due to atmospheric conditions and other parameters. Sometimes the pitch attitude increases to a large value, causing the tail section of the aircraft to descend dangerously close to the ground during approach and touchdown. In some cases, the tail body contact margin requirement may be exceeded, causing the tail section of the aircraft body to contact the ground, resulting in damage to the aircraft.

Various proposals have been made to reduce either the effect or the likelihood of tail section contacts during landing. One such proposal to reduce the effect of tail section contacts involves adding a tail section skid to absorb light tail section contacts with little or no resulting damage to the aircraft structure.

The inclusion of a tail skid has several disadvantages. First, due to the addition of the tail skids to the aircraft structures, the tail skids reduce the geometric structural clearance between the aircraft and the ground. Second, the tail skids require modification of the aircraft structures. Third, the tail skids increase the weight and thereby reduce the payload capacity of an aircraft. Fourth, in the event of a tail skid contact, an inspection of the integrity of the tail skid as well as the aircraft itself is required. This inspection is economically burdensome. Furthermore, the resulting reduction in structural clearance ironically increases the probability of an aircraft contacting the ground, thereby causing the disadvantages mentioned above.

The probability of tail body contacts during approach and landing involves re-adjusting a number of flight parameters to reduce the risk of tail body contact. These adjustments include increasing landing approach speeds, increasing trailing edge flap deflections at the flap detents, and increasing symmetrical lateral control surface descent. To address the same problem, vortex generators have been added to the aircraft wings. All of these solutions are successful in reducing the landing pitch attitude of an aircraft by resulting in a larger tail body contact margin. However, the solutions simultaneously reduce the nose gear contact margin. Therefore, although the probability of tail body contacts is reduced, the probability of nose gear first contacts is concomitantly increased.

Like tail body contacts, nose gear first contacts can be avoided by adjusting a number of flight parameters. Reducing landing approach speeds, reducing trailing edge flap deflections at the flap detents, and reducing symmetrical droop of the side control surfaces all increase the landing pitch attitude of an aircraft. This increase more improves the nose gear contact margin. The disadvantage of these various attempts is that an increase in pitch attitude concomitantly reduces the tail body contact margin. The resulting reduction in pitch attitude Tail body contact margin consequently increases the probability of tail body contacts.

To overcome the above-mentioned disadvantages, pitch attitude limitation techniques have been proposed. As a result, a pitch attitude limitation automatically limits pitch attitudes to a predetermined range. Although pitch attitude limitations can serve to reduce the likelihood of nose gear first contacts and tail body contacts, they unduly limit a pilot's control of the aircraft. Limiting a pilot's control of an aircraft is undesirable because it violates some fundamental aircraft design philosophies that dictate that a pilot have absolute control of the aircraft.

Other proposals have been made to reduce the likelihood of nose gear first landings without also increasing the likelihood of tail body contacts. One such proposal, commonly referred to as direct lift control for flight path control, is to provide a closed loop control law that decouples flight path control from pitch attitude control. Decoupling flight path control from pitch attitude control allows changes in flight path angle to be made with little or no change in pitch attitude. Flight path is controlled by modulating wing control surfaces and flaps, not by modifying pitch attitude.

Although the direct lift control proposal has some advantages, it has significant disadvantages. Because the direct lift control technique separates the flight path control from the longitudinal If decoupled from pitch attitude control, direct lift control, unlike the flight control systems of virtually all conventional large commercial transport aircraft, would result in aircraft handling characteristics that are unfamiliar to commercial pilots. This proposal would likely require commercial pilots to undergo additional training to learn the different handling characteristics of an aircraft incorporating direct lift control. In addition to unconventional maneuvering characteristics, direct lift control also presents logistical difficulties. Implementing direct lift control would require complex system changes to ensure acceptable pilot and flight control system interaction. Furthermore, the use of spoilers, or auxiliary wings, often used by direct lift control, could result in unacceptable airframe shake, unduly compromising passenger comfort.

From the above identified prior art document US-A-5 446 666 a flight control system for an aircraft, especially a helicopter, is known which is automatically transformed from a unique trim control mode to a displacement control mode, or vice versa, when the aircraft touches down or takes off. This transformation involves a gradual adjustment of control laws which is automatically carried out on the basis of signals from sensors monitoring the condition of the aircraft landing gear. There is no mention of any contact span at all in this document.

Another prior art document, US-A-3 887 148, describes a flight control system using a glide slope coupler for guiding an aircraft along a high frequency defined glide slope by floating out to the point of landing. The system does not contain any provisions for maintaining the pitch attitude of the aircraft between the nose gear and tail body spans.

The present invention is directed to providing a method for improving the nose gear contact margin and/or the tail body contact margin without the foregoing and other disadvantages of prior methods.

Summary of the invention

According to the present invention, a method of modifying the attitude of an aircraft during landing of the type described above is characterized in that said table or program of deflection values is selected such that a predetermined tail body contact margin and/or a predetermined nose gear contact margin for the aircraft is met or exceeded.

According to other aspects of this invention, the movable aircraft surfaces are flaperons.

According to further aspects of this invention, the predetermined flight condition parameter is the approach airspeed, the reference value is the reference approach airspeed Vref, and the current value is the current approach airspeed Vcur. The difference between the reference approach airspeed Vref and the current approach airspeed Vcur is an airspeed difference ΔV.

According to still other aspects of this invention, the flaperons are adjusted symmetrically by an amount equal to a flaperon depression change ΔδFR. The flaperons are adjusted symmetrically by an amount equal to a flaperon depression change ΔΔFR. Flaperon aileron depression change ΔδFR is selected according to a predetermined table or according to a predetermined program, wherein the flaperon or flap aileron depression change ΔδFR is a function of the airspeed difference ΔV. The predetermined table or the predetermined program provides a value for the flaperon or flap aileron depression change ΔδFR according to the equation:

ΔδFR = 31/15 (ΔV-5)

if the airspeed difference ΔV has a value between 5 and 20. If the airspeed difference ΔV is less than 5, the flaperon or flap aileron lowering or extending change ΔδFR has a value of zero. If the airspeed ΔV is greater than 20, the flaperon or flap aileron lowering or extending change ΔδFR has a value of 31.

According to still further aspects of this invention, the reference approach airspeed Vref is determined from the following equation:

wherein:

nZCG = normal load or stress factor in the center of gravity

CLref = Reference lift coefficient

CLcur = current lift coefficient.

According to still further aspects of this invention, the normal center of gravity load factor, ZCG, is limited in value to provide maneuvering compensation.

According to still other further aspects of this invention, a signal indicative of flaperon aileron pitch change ΔδFR is low pass filtered to dampen high frequency turbulence.

According to still further other aspects of this invention, the flaperon pitch or extension change is held constant when the aircraft touches the ground.

According to alternative aspects of this invention, the predetermined flight state parameter is the approach attitude, the reference value is the reference approach attitude θref, the current value is the current approach attitude θcur, and the difference value is an attitude difference Δθ.

According to further alternative aspects of this invention, the predetermined flight condition parameter is the approach angle of attack, the reference value is the reference approach angle of attack αref, the current value is the current approach angle of attack αcur, and the difference value is an angle of attack difference Δα.

As can be readily seen from the foregoing outline, the invention provides an aircraft landing attitude modifier (LAM) that improves nose gear contact margins and/or tail body contact margins. The improved margins result from the automatic, symmetrical variation of aircraft movable surfaces and/or high lift surfaces, including, for example, the flaperons. Because the LAM provides increased tail body contact margin, the LAM avoids the need for tail skids designed to protect against tail skids. Accordingly, increased weight, degradation in structural clearance, and economic expense associated with tail skids are avoided. Furthermore, because the LAM can improve both the nose gear contact margin and tail gear contact margin of an aircraft, the limitations associated with the conventional techniques of merely adjusting landing approach speeds, trailing edge flap deflections to the flap stops, and symmetrically lowering or extending the lateral control surface are avoided. By allowing improved contact margins at both extreme ends of an aircraft's landing pitch attitude envelope, the LAM overcomes the limitations of these adjustments which merely improve tail gear contact margin at the expense of nose gear contact margin, or vice versa. Because the LAM does not artificially limit the available pitch attitude for an aircraft, absolute pilot control of the aircraft is maintained, unlike some conventional pitch limitation methods. The LAM also provides an open-loop control law that maintains the coupling between flight path control and pitch attitude control, thus avoiding the anomalous maneuvering characteristics and complexity of direct lift control methods that use closed-loop control laws.

Short description of the drawings

The foregoing aspects and many of the attendant advantages of this invention will be more readily understood when the same is considered with reference to the following detailed description, taken in conjunction with the accompanying drawings, is better understood, wherein:

Fig. 1 illustrates the tail body contact margin requirement of an aircraft;

Fig. 2 illustrates the nose gear contact margin requirement of an aircraft;

Fig. 3 is a control law diagram of a landing attitude modifier for an aircraft constructed in accordance with this invention;

Fig. 4 is a functional diagram of a Vref determination suitable for use in the embodiment of the invention shown in Fig. 3 ;

Fig. 5 is a control law diagram of an alternative embodiment of the invention;

Fig. 6 is a functional diagram of another alternative embodiment of the invention;

Fig. 7 is a functional diagram of an αref determination suitable for use in the embodiment of the invention shown in Fig. 6;

Figures 8A to 8C are diagrams illustrating the performance of the invention shown in Figures 3 to 7;

Fig. 9 is a control law diagram of yet another alternative embodiment of the invention; and

Fig. 10 is a diagram illustrating the performance of the invention shown in Fig. 9.

Detailed description of the preferred embodiment

As is conventional, the invention is illustrated and described in the form of a control law containing discrete blocks designed to accomplish specific functions. However, it will be understood that the invention may actually be embodied in various ways. For example, the various functions of the illustrated control law may be performed by a suitably programmed computer. Alternatively, the functions may be performed by digital or analog circuits.

Fig. 3 illustrates a first embodiment of a landing attitude modifier (LAM) 10 constructed in accordance with this invention in the form of a control law comprising a plurality of blocks containing elements configured to accomplish various functions. The LAM 10 positions the flaperons of an aircraft to improve the nose gear contact margin and the tail body contact margin during landing of an aircraft. The LAM 10 adjusts the flaperon extension symmetrically from the nominal position in response to the difference between the current approach condition and the reference approach condition of the aircraft. The adjustment provides a reduced pitch attitude variation for landings of the aircraft. Although the preferred embodiment provides adjustment of the Flaperons or landing flap ailerons of an aircraft, it is understood that the LAM 10 could also be used to adjust other lift-generating movable surfaces or combinations of lift-generating movable surfaces of an aircraft just as well.

The embodiment of the invention illustrated in Fig. 3 includes a reference approach speed Vref determination block 12, a summer 14, a flaperon aileron extension change ΔδFR lookup block 16, a gain block 18, a low pass filter 20, and a LAM on/off and command freeze determination block 24. The functions of the reference approach speed Vref determination block 12 are preferably performed in the manner illustrated in Fig. 4. Essentially, the reference approach speed determination block 12 implements the following formula:

wherein:

Vref = reference approach airspeed

nZCG = normal load or stress factor in the center of gravity

Vcur = current approach airspeed

CLcur = current lift coefficient

CLref = reference lift coefficient.

Fig. 4 comprises a table or program block 26 for the current lift coefficient CLcur, a table or program block 28 for the reference lift coefficient CLref, a first divider 30, a first square root block 32, a multiplier 34, a second divider 36, a limiter 38 and a second square root block 40. The inputs to the current lift coefficient CLcur table block 26 are an aircraft angle of attack α and a flap detent or angle δF. The current lift coefficient CLcur may be determined mathematically in a conventional manner as a linear approximation based on the angle of attack α for a given flap detent or angle δF. Alternatively, values of the current lift coefficient CLcur may be stored in a three-dimensional table and interpolation used to determine current values.

The Mach number M and a flap detent or flap angle δF are the inputs to the reference lift coefficient CLref table block 28. Based on the Mach number M and the flap detent or flap angle δF, the reference lift coefficient CLref table block 28 determines a reference lift coefficient CLref. Preferably, the reference lift coefficient CLret is 1.53 for a flap detent of 30 and 1.39 for a flap detent of 25 for the Boeing 777. As with the current lift coefficient CLcur block 26, this can be done mathematically or using a three-dimensional lookup table. As is well known to those skilled in aerodynamics, the reference lift coefficient CLref is a reference value related to the tail body contact span, nose gear contact span and landing field length requirements of an aircraft.

The first divider 30 divides the current lift coefficient CLcur by the reference lift coefficient CLref. The square root of the coefficient is determined (block 32), and the result is multiplied by the current approach airspeed Vcur in the multiplier 34.

The normal center of gravity load factor, nZCG, provides maneuvering compensation by maintaining the baseline stability and maneuvering sensitivity of the aircraft relative to the normal commercial transport maneuvering range. The normal center of gravity load factor, nzcG, is applied to and limited by the limiter 38. The maximum and minimum values of the limiter 38 are selected to define the range of the normal center of gravity load factor, nZCG, over which baseline stability is maintained and to ensure that there is no degradation in the absolute maneuverability of the aircraft. Preferably, the normal center of gravity load factor, nZCG, is limited to a maximum value of 1.15 g and a minimum value of 0.85 g. The square root of the constrained value of the normal center of gravity load factor, nZCG, is determined by the second square root block 40. The output of the multiplier 34 is divided by the output of the second square root block 40 in the second divider 36. The output of the second divider is the reference approach airspeed Vref. As will be appreciated by those skilled in the art and others, the reference approach speed Vref may also be determined by other methods. For example, the reference approach speed Vref may be based on a table or program that depends on the gross weight of the aircraft at a given flap position.

Returning to Fig. 3, the reference approach speed Vret is subtractively summed with the current approach airspeed Vcur in the summer 14. The airspeed difference ΔV is entered on the table or Program block 16 for the flaperon or landing flap aileron extension or lowering change ΔδFR. In the preferred embodiment of the invention, the flaperon or landing flap aileron extension or lowering change ΔδFR is a function of the airspeed difference ΔV. If the airspeed difference ΔV is less than 5, the flaperon or landing flap aileron extension or lowering change ΔδFR has a value of zero. If the airspeed difference ΔV is between 5 and 20, the flaperon or landing flap aileron extension or lowering change ΔδFR is determined according to the following equation:

If the airspeed difference ΔV is greater than 20, the flaperon extension or lowering change ΔδFR has a value of 31. The flaperon extension or lowering change table or program block may vary from aircraft to aircraft. The flaperon extension or lowering change ΔδER is applied to the gain block 18. As will be described in more detail below, the gain block 18 applies a gain of either 0 or 1 to ΔδFR, depending on the LAM on/off and command freeze determination 24. Accordingly, the gain block is essentially an on/off switch. The output of the gain block 18 is applied to the low pass filter 22. The Laplace transform in the low pass filter 22 is:

1/τs + 1

where τ may vary from aircraft to aircraft. Preferably τ = 2 seconds for the Boeing 777. The low pass filter 20 attenuates the high frequency signal content due to turbulence and preserves the short period dynamic response of the aircraft while maintaining responsiveness to moderate wind shear. The bandwidth of the low pass filter 20 also allows flaperon ailerons to be repositioned to their nominal position prior to a stall warning. The output of the low pass filter 20, a flaperon aileron extension or deduction change command, ΔδFR command, is applied to another summer 22. A nominal flaperon aileron extension or deduction signal is also applied to the summer 22. The summer subtracts the flaperon extension or down change command, ΔδFR command, from the nominal flaperon extension or down signal. The output of summer 22 is a flaperon position command.

The LAM on/off and command freeze determination block 24 receives the flap detent or angle δF input, a composite signal validity input, and a discrete air/ground input. Based on the flap detent or angle input, δF input, and the state of the signal validity and discrete air/ground inputs, the LAM on/off and command freeze determination block 24 turns the LAM 10 on and off. For normal system operations, the state of the signal validity inputs indicates that a set of conditions described below is met, and the LAM on/off and Command freeze determination block 24 produces an input output when the flap detent or angle input, δF input, indicates that the aircraft's flaps are in a flap detent position. Preferably, this requires a flap detent of 25º or 30º for the Boeing 777.

If the flap detent or angle input, δF input, and the signal validity and discrete air-to-ground inputs are all favorable, the LAM on/off and command freeze determination block 28 outputs an on signal to the gain block 18 which causes the gain block 18 to have a value of 1. If the gain block 18 has a value of 1, the LAM 10 becomes operative. If any of the flap detent or angle, δF, the signal validity or the discrete air-to-ground inputs are not favorable, the LAM on/off and command freeze determination block 24 outputs an off signal which causes the gain block 18 to have a zero value. The zero gain value deactivates the LAM 10 and returns the flaperons to their nominal extension or lowering position.

The state of signal validity is determined by checking a series of conditions. The check includes a flaperon operation check, a LAM compensated speed band replay check, an airspeed check, an angle of attack check, α check, and a Mach number check, M check. The flaperon operation check determines whether the left and right flaperons are operating correctly. The LAM compensated speed band replay check Check determines whether the aircraft's warning electronics system has compensated the LAM 10 in its speed band response and function. The airspeed check, the angle of attack check, α-check, and the Mach number check, M-check, involve checking a variety of sources using redundancy management algorithms to ensure signal validity, as well as determining whether each parameter falls within a desired range of values. If all of the checks are satisfactory, the signal validity state is valid. If any of the checks fail, the signal validity state is invalid, causing the gain block 10 to have a value of zero.

The state of the discrete air-to-ground input to the LAM on/off and command freeze determination block 24 is based on a check of the main gear landing gear position to determine whether the aircraft is on the ground or in flight. Once the landing gear position indicates touchdown, the LAM on/off and command freeze determination block 24 outputs an aircraft on ground command freeze to the low pass filter 50. The aircraft on ground command freeze holds the output of the low pass filter 20 constant after touchdown, effectively freezing the flaperon aileron depression at its touchdown position. Until the aircraft on the ground freeze command is received, the LAM 10 allows adjustment of the flaperon or landing flaps lowering or extension in response to the current approach airspeed Vcur.

Fig. 5 illustrates an alternative embodiment of a LAM 41 constructed in accordance with this invention. The LAM 41 illustrated in Fig. 5 comprises: a table block 42 for the reference approach attitude ref, a summer 44, a flaperon aileron extension change ΔδFR lookup block 46, a gain factor block 48, a low pass filter 50, and a LAM on/off and command freeze determination block 54. The reference approach attitude θref lookup block 42 receives a flap detent or angle input, δF input. Based on the flap detent or angle input, δF input, the reference approach attitude θref, the lookup block 42 determines a reference approach attitude θref. The reference approach attitude θref represents the pitch attitude of an aircraft when flying at the desired landing flight path angle of the reference approach airspeed Vref and a given flap detent or flap angle δF. Preferably, the reference approach attitude θref has a value of 3.0 degrees for a flap detent of 25 and a value of 2.8 degrees for a flap detent of 30. The flap detent positions of 25 and 30 are normal flap detent positions for the Boeing 777.

The output of the table or program block 42 for the reference approach attitude θref is fed to the summer 44. The current approach attitude θcur of the aircraft is also fed to the summer 44. The summer 44 subtracts the reference approach attitude θref from the current approach attitude θcur. The output of the summer 44 is an attitude difference Δθ.

The flight attitude difference Δθ is given to the table or program block 46 for the flaperon or landing flap aileron extension change ΔδFR. The table or program block 46 for the flaperon extension change ΔδFR generates a flaperon extension change ΔδFR depending on the attitude difference Δθ. More specifically, the flaperon extension change ΔδFR has a value of 31 degrees when the attitude difference Δθ has a value of -3 degrees or less. The flaperon extension change ΔδFR has a value of 0 degrees when the attitude difference Δθ has a value of -1 degree or greater. When the attitude difference Δθ has a value between -3 degrees and -1 degree, the value of the flaperon extension change ΔδFR is determined according to the following equation:

ΔδFR = -31/2 (Δθ + 1)

These flaperon extension change values, ΔδFR values, are preferred values for the Boeing 777, but may be different for other aircraft.

The output of the flaperon extension change lookup block 46 ΔδFR is applied to the gain block 48. The operation of the gain block 48, the low pass filter 50, and the LAM on/off and command freeze determination 54 functions in the same manner as these elements function in the embodiment of the invention illustrated in Figure 3 and described above. Consequently, these blocks will not be discussed further. The output of the LAM 41, illustrated in Figure 5, is a flaperon extension change command, ΔδFR command, which, as in the embodiment of the invention shown in Figure 3, is subtractively summed with a nominal flaperon extension in a summer 52 to produce a flaperon position command.

Figure 6 illustrates another alternative embodiment of a LAM 81 constructed in accordance with this invention. The LAM 81 illustrated in Figure 6 includes: a reference approach angle of attack determination block 80, a summer 82, a flaperon extension change (ΔδFR) lookup block 84, a gain block 86, a low pass filter 88, and a LAM on/off and command freeze determination block 92. Figure 7 functionally illustrates how a reference approach angle of attack αref is determined using the reference approach angle of attack αref determination block 80. The determination block 80 for the reference approach angle of attack αref comprises a table or program block 94 for the reference lift coefficient C1ref and a table or program block 96 for the current lift coefficient CLcur.

The reference lift coefficient CLref table block 94 receives a Mach number M input and a flap detent or angle δF input. Based on these inputs, a reference lift coefficient CLref is determined. Preferably, the reference lift coefficient CLref has a value of 1.53 for a flap detent of 30 and a value of 1.39 for a flap detent of 25 for the Boeing 777. As noted above, the flap detent of 25 and 30 are normal flap detent landing settings for the Boeing 777. The reference lift coefficient CLref is entered into the current lift coefficient CLcur table block 96. The current lift coefficient CLcur table block 96 determines a reference approach angle of attack αref based on CLref and the flap detent or angle δF input.

Referring to Fig. 6, the reference approach angle of attack αref is input to summer 82. The current approach angle of attack αcur of the aircraft is also input to summer 82. The summer subtracts the reference approach angle of attack αref from the current approach angle of attack αcur. The output of summer 82 is an angle of attack difference Δα. The angle of attack difference Δα is input to flaperon extension change ΔδFR table block 84. Flaperon extension change ΔδFR table block 84 produces a flaperon extension change ΔδFR in response to the angle of attack difference Δα.

More specifically, the flaperon extension change ΔδFR has a value of 31 degrees when the angle of attack difference Δα has a value of -3 degrees or less. The flaperon extension change ΔδFR has a value of 0 degrees when the angle of attack difference Δα has a value of -1 degrees or greater. When the angle of attack difference Δα has a value between -3 degrees and -1 degrees, the value of the flaperon extension change ΔδFR is determined according to the following equation:

ΔδFR = -31/2 (Δα + 1)

These flaperon extension change values, ΔδFR values, are preferred values for the Boeing 777. They may be different for other aircraft types. The flaperon extension change ΔδFR is input to the gain block 86. The gain block 86, the low pass filter 88 and the LAM on/off and command freeze determination block 92 function in the same manner as the corresponding elements in the embodiment of the invention illustrated in Fig. 3 and described above. Therefore, its operation will not be discussed here. The output of the LAM 81 illustrated in Fig. 6 is a flaperon extension change command, ΔδFR command, which, as in the embodiment shown in Fig. 3, is subtractively summed with a nominal flaperon extension in a summer 90 to produce a flaperon position command.

As indicated above, the LAM can increase both the nose gear contact margin and/or the tail body contact margin. Figures 8A through 8C illustrate how a LAM 10 of the type shown in Figure 3 improves the nose gear contact margin and the tail body contact margin compared to an aircraft not having such a LAM. Figure 8A is a graph illustrating the variation (line 60) of a lift coefficient CL with approach pitch attitude (θ) for an aircraft without a LAM for a constant approach flight path angle. The abscissa is the variation in approach pitch attitude (θ) in degrees and the ordinate indicates the change in lift coefficient CL. Line 60 has a constant positive pitch. Two spaced vertical lines represent a nose gear initial contact and a tail body contact. As recognized by those skilled in the art and others, to compensate for wind gusts, the speed of an aircraft will normally be between Vref and Vref+20 during approach and touchdown. Point A on line 60 corresponds to when the aircraft is flying at Vref. Point B on line 60 corresponds to when the aircraft is flying at Vref+20. The horizontal distance between point A and the line indicating tail body contact defines a tail body contact span when flying at Vref. Similarly, the horizontal distance defines The horizontal distance between point B and the line indicating nose gear initial contact defines a nose gear contact margin when flown at Vref+20. The horizontal distance between point A and point B on line 60 defines a pitch attitude variation Aθ1. Pitch attitude variation Aθ1 represents the total approach pitch attitude variation for an aircraft not having a LAM at a center of gravity.

Fig. 8B illustrates how the nose gear contact margin is improved by the addition of the LAM to the aircraft represented in Fig. 8A. Line 60 of Fig. 8A is shown in dashed form and is included for reference purposes. Line 62 illustrates the new performance provided by the addition of a LAM. Point C on line 62 indicates the situation when the aircraft is flying at Vref+20. The horizontal distance separating point B on line 60 and point C on line 62 defines a nose gear contact margin improvement θimp,n. Points D and E on line 62 correspond to the maximum and minimum values produced by block 16 for the flaperon extension change ΔδFR shown in Fig. 3. The end portions 64, 66 on line 62 represent the operating region of the LAM corresponding to the maximum and minimum constant values of the flaperon extension variation ΔδFR where the flaperons are held at fixed extreme deflections. The region on line 62 between point D and point E represents the aircraft configuration change caused by the installation of a LAM. The horizontal distance between the points on line 62 corresponding to Vref and Vref+20 defines a pitch attitude variation Δθ2. A comparison of Δθ1 in Fig. 4A with Δθ2 in Fig. 4B illustrates that Δθ2 is less than Δθ1. The relatively smaller Δθ₂ value indicates an improvement in the pitch attitude variation and, more specifically, in the nose gear contact margin.

Fig. 8C illustrates an improved tail body contact margin effected by installing a LAM 10 of the type shown in Fig. 3 at the reference approach speed Vret. Line 70 (partially dashed) illustrates the variation of the lift coefficient CL with the approach pitch attitude θ without a LAM. The horizontal distance between point F on line 70 and the line indicating tail body contact defines a tail body contact margin for an aircraft without a LAM. The horizontal distance between point G on line 70 and the line indicating nose gear initial contact represents a nose gear contact margin for an aircraft without a LAM. Line 72 (partially dashed) illustrates the variation of the lift coefficient CL with the approach pitch attitude θ with an increase in the landing approach speed. At the reference approach speed, i.e. at Vref, the tail body contact margin is improved. The horizontal distance between point H on line 72 and the line indicating the tail body contact is greater than the distance between point F and the line indicating the tail body contact. The improvement in the tail body contact margin is given by θimp,a.

If no LAM were included, line 72 (partially dashed part) would intersect the line indicating nose gear initial contact at Vref+20. As a result, there would be no nose gear contact margin. This does not occur. Line 74 illustrates how a LAM changes the situation. The LAM causes a change in pitch to occur (line 74) when the reference speed changes. Line 74 intersects line 70 before reaching the nose gear contact margin. As a result, the nose gear contact margin is maintained or secured without a LAM.

As will be readily appreciated by an understanding of the foregoing discussion, a LAM can also be used to provide both a tail body contact margin improvement and a nose gear contact margin improvement. Improvement of both the nose gear contact margin and the tail body contact margin can be achieved by providing a LAM that combines the features depicted in Figures 8B and 8C.

Figure 9 illustrates an alternative method for improving the tail body contact margin by modifying a LAM 10 of the type shown in Figure 3. The LAM 10 shown in Figure 9 is identical to the LAM 10 shown in Figure 3 except for the substitution of a flaperon extension change ΔδFR table block 102 and a summer 100 which replace the flaperon extension change ΔδFR table block 16 and summer 22 of Figure 3, respectively. As shown in Figure 9, the flaperon extension change ΔδFR has a maximum value when the airspeed difference ΔV has a value that is less than a predetermined low value. When the airspeed difference ΔV has a value between a predetermined lower value and a predetermined higher value, the value of the flaperon extension change ΔδFR has a value which continuously decreases from the maximum value to a minimum value. When the airspeed difference is greater than the predetermined higher value, the flaperon extension change ΔδFR has the minimum value.

The flaperon extension change command, ΔδFR command, is applied to the summer 100 along with the nominal flaperon extension signal. The flaperon extension change signal, ΔδFR signal, is added to the nominal flaperon extension signal. The output of the summer 100 is a flaperon position command. Because the function of the LAM illustrated in Fig. 9 is otherwise identical to the function of the LAM illustrated in Fig. 3, no further description is given here.

Figure 10 illustrates how the tail body contact margin is improved by the addition of the LAM depicted in Figure 9. As shown in Figure 10, line 80 corresponds to when the aircraft is flying at some nominal deployment position. As will be appreciated by those skilled in the art and others, the nominal deployment position may have different values for different aircraft. The distance between point R and the line indicating tail body contact is the tail body contact margin for the aircraft flying at Vref.

Line 82 corresponds to when the aircraft undergoes such a change that it has a flaperon deployment position with a deflection that produces maximum lift. Of course, the maximum lift deflection value varies from aircraft to aircraft. The area on line 84 between point B and point Q represents the operating range of the LAM corresponding to the maximum and minimum constant value of the flaperon deployment change ΔδFR where the flaperons are held at fixed extreme deflections. The horizontal distance between point S on line 82 and point R on line 80 represents the improvement in the tail body contour. clock margin caused by the addition of the LAM as given by θimp,a.

Claims (12)

1. A method for modifying the attitude of an aircraft during landing, comprising: Generating a reference value for a predetermined parameter which defines a flight state of the aircraft; Generating a current value for the predetermined flight condition parameter; subtractively combining the reference value and the current value to produce a difference value for the predetermined flight condition parameter; generating a table or program of deflection values for at least one movable surface capable of generating lift, depending on the difference value; Determining the deflection value for the at least one movable surface from the table or the program, based on the difference value; Adjusting the position of the at least one movable surface according to an amount equal to the deflection value, characterized in that the table or program of deflection values is selected such that a predetermined tail body contact margin and/or a predetermined nose gear contact margin for the aircraft is met or exceeded. 2. A method according to claim 1, wherein the predetermined flight state parameter is the approach airspeed, the reference value is the reference approach airspeed Vref, the current value is the current approach airspeed Vcur and the difference value is an airspeed difference ΔV is. 3. A method according to claim 1 or 2, wherein the at least one movable surface is a flaperon. 4. The method of claim 3, wherein the deflection value is the flaperon pitch change ΔδFR and the tabulating step further comprises: Provide a value of zero for the flaperon or landing flap aileron lowering or extension change ΔδFR when the airspeed difference ΔV is less than 5; Provide a value for the flaperon or flap aileron lowering or extension change ΔδFR according to the equation: ΔδFR = 31/15(ΔV - 5) if the airspeed difference ΔV is between five and 20, and Provide a value of 31 for the flaperon or landing flap aileron lowering or extension change ΔδFR if the airspeed difference ΔV is greater than 20. 5. The method of claim 4, comprising fixing the flaperon deployment change ΔδFR when the aircraft contacts the ground. 6. A method as claimed in any one of claims 2 to 5, wherein generating a reference value comprises determining Vref from the following equation: wherein: Vref = reference approach airspeed nZCG = normal load or stress factor at the centre of gravity Vcur = current approach airspeed cLcur = current lift coefficient cLref = reference lift coefficient 7. A method according to claim 6, comprising limiting the normal center of gravity load factor nZCG so as not to provide any degradation in the absolute maneuverability of the aircraft. 8. A method as claimed in any one of claims 2 to 5, wherein the generating step comprises determining Vref for the following equation: wherein: Vref = reference approach airspeed Vcur = current approach airspeed cLcur = current lift coefficient cLref = reference lift coefficient 9. A method as claimed in any one of claims 2 to 8, wherein the generating step comprises determining the reference approach airspeed Vref from a table of values depending on the gross weight and flap detent or position of the aircraft. 10. A method according to any one of the preceding claims, wherein the predetermined flight condition parameter is the approach attitude , said reference value is the reference approach attitude Θref, said current value is the current approach attitude Θcur and said difference value is an attitude difference ΔΘ. 7.1. A method as claimed in any one of claims 1-10, wherein the predetermined flight condition parameter is the approach angle of attack, said reference value is the reference approach angle of attack αcur and said difference value is an angle of attack difference Δα. 12. The method of claim 10 or 11, wherein the deflection value is a flaperon pitch change ΔδFR and the tabulating step further comprises: providing a value of zero for the flaperon or landing flap aileron lowering or extending variation ΔδFR if the said difference value is greater than -1; Provide a value for the flaperon or flap aileron lowering or extension change AδFR according to the equation: ΔδFR = -31/2(difference value + 1) if the difference value is between -3 and -1; and Provide a value of 31 for the flaperon or flap aileron lowering or extending change ΔδFR if the difference value is less than -3.