CA2591645C - Variable gradient control stick force feel adjustment system - Google Patents
Variable gradient control stick force feel adjustment system Download PDFInfo
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- CA2591645C CA2591645C CA2591645A CA2591645A CA2591645C CA 2591645 C CA2591645 C CA 2591645C CA 2591645 A CA2591645 A CA 2591645A CA 2591645 A CA2591645 A CA 2591645A CA 2591645 C CA2591645 C CA 2591645C
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Abstract
A gradient actuator is operationally coupled to lateral and longitudinal springs of a tilt rotor aircraft's latitudinal and longitudinal control mechanism. The gradient actuator is further coupled to a Nacelle position sensor for tracking the Nacelle position of the tilt rotor aircraft during operation. The force feel of a cyclic controller in an aircraft and the aircraft's operational control are obtained when the gradient actuator is configured to change the moment arm length of the lateral spring cartridge assembly with respect to a lateral actuator and the longitudinal spring cartridge assembly with respect to a longitudinal actuator. The change in moment arm length is based on input to the gradient actuator from the Nacelle position sensor.
Description
VARIABLE GRADIENT CONTROL STICK
FORCE FEEL ADJUSTMENT SYSTEM
This is a division of co-pending Canadian Patent Application No. 2,380,346 filed August 4, 2000.
TECHNICAL FIELD
The present invention is directed to the field of control stick force adjustment systems as used in aircraft and, more particularly, to an improved variable gradient control stick force feel adjustment system.
BACRGROIIND OF TEE INVENTION
Control stick force adjustment systems are used in the aircraft aviation field to provide pilots with a better 'feel' and control over their aircraft by adjusting the tension of the manual s control system (e.g, control stick, cyclic stick, steering, peddles, etc.,) at varying air speeds. For example, In a conventional force trim mechanism for a helicopter cyclic system, only a fixed force gradient is provided. In simple terms, for every increment of cyclic displacement, the pilot -feels a proportional force. It is of course desirable that a certain amount of force is encountered in any direction a pilot moves the controller (be it left or right, forward or backward). Force on a cyclic stick provides the pilot, and ultimately the aircraft, with stability during airborne operations. The force, typically, is is generated by a four bar linkage that compresses or extends a spring cartridge. Two linkage assemblies are utilized, one for lateral motion and another for longitudinal motion. By. moving the spring cartridge grounding points, the position where the pilot using the cyclic stick feels zero force can be moved. Actuators called Force Trim Actuators are also used to move the spring cartridge grounding points. Because the linkages of the conventional lateral or longitudinal force trim system move in a fixed plane, these linkages are considered two-dimensional.
It should be appreciated that the force encountered in the typical helicopter operation is a substantially linear relationship. When operating an airplane, however, a pilot normally encounters a much stiffer control stick because a much higher spring force is required as the aircraft travels at higher airspeeds. Instead of moving the stick forward and backward, or s the steering assembly left or right, with a normal force of one pound per inch, a pilot should encounter approximately 3 pounds per inch. Without the additional force, an aircraft flying at high speeds could undergo very erratic and dangerous aircraft movement.
Many prior controller force adjustment systems utilize electric motors to place a higher torque on the control stick, resulting in a higher tensioned feel. Although the force trim systems for some aircraft incorporate a spring tension against any force exerted by the pilot against the pilot-controlled directional gear, automated control is the predominant technology in later is model aircraft. For example, in current tilt rotor aircraft applications, a variable force field actuator takes a given parameter (e.g., tilt rotor position or airspeed) and uses an electric motor to in-turn cause an increasing or decreasing force against the pilot-controlled directional system, based on inputs to the electric motor by a controller. Such a system is not only heavy but also very expensive because of the electronics in controlling the motor and the redundancy that may be required with automated systems in order to safeguard against potential system failures.
is Many problems in achieving variable tension on manual controllers are unique to a tilt rotor aircraft because it functions as both an airplane and a helicopter. Because a tilt rotor aircraft operates as both, it is desirable to have the feel of the tilt rotor aircraft change as it is converted from an airplane to a helicopter, and vice a versa, during flight. The way s that the 'feel' and resulting handling capabilities are accomplished currently in tilt rotor aircraft systems (such as the Bell XV1S and the V22 tilt rotor aircraft), is to use the heavier, more expensive variable force field actuator systems, as described above. It would be more desirable in tilt rotor aircraft applications, and for the aircraft industry as a whole, to have access to a less complicated, lighter and more reliable variable gradient cyclic force feel system, such as disclosed in the present invention.
FORCE FEEL ADJUSTMENT SYSTEM
This is a division of co-pending Canadian Patent Application No. 2,380,346 filed August 4, 2000.
TECHNICAL FIELD
The present invention is directed to the field of control stick force adjustment systems as used in aircraft and, more particularly, to an improved variable gradient control stick force feel adjustment system.
BACRGROIIND OF TEE INVENTION
Control stick force adjustment systems are used in the aircraft aviation field to provide pilots with a better 'feel' and control over their aircraft by adjusting the tension of the manual s control system (e.g, control stick, cyclic stick, steering, peddles, etc.,) at varying air speeds. For example, In a conventional force trim mechanism for a helicopter cyclic system, only a fixed force gradient is provided. In simple terms, for every increment of cyclic displacement, the pilot -feels a proportional force. It is of course desirable that a certain amount of force is encountered in any direction a pilot moves the controller (be it left or right, forward or backward). Force on a cyclic stick provides the pilot, and ultimately the aircraft, with stability during airborne operations. The force, typically, is is generated by a four bar linkage that compresses or extends a spring cartridge. Two linkage assemblies are utilized, one for lateral motion and another for longitudinal motion. By. moving the spring cartridge grounding points, the position where the pilot using the cyclic stick feels zero force can be moved. Actuators called Force Trim Actuators are also used to move the spring cartridge grounding points. Because the linkages of the conventional lateral or longitudinal force trim system move in a fixed plane, these linkages are considered two-dimensional.
It should be appreciated that the force encountered in the typical helicopter operation is a substantially linear relationship. When operating an airplane, however, a pilot normally encounters a much stiffer control stick because a much higher spring force is required as the aircraft travels at higher airspeeds. Instead of moving the stick forward and backward, or s the steering assembly left or right, with a normal force of one pound per inch, a pilot should encounter approximately 3 pounds per inch. Without the additional force, an aircraft flying at high speeds could undergo very erratic and dangerous aircraft movement.
Many prior controller force adjustment systems utilize electric motors to place a higher torque on the control stick, resulting in a higher tensioned feel. Although the force trim systems for some aircraft incorporate a spring tension against any force exerted by the pilot against the pilot-controlled directional gear, automated control is the predominant technology in later is model aircraft. For example, in current tilt rotor aircraft applications, a variable force field actuator takes a given parameter (e.g., tilt rotor position or airspeed) and uses an electric motor to in-turn cause an increasing or decreasing force against the pilot-controlled directional system, based on inputs to the electric motor by a controller. Such a system is not only heavy but also very expensive because of the electronics in controlling the motor and the redundancy that may be required with automated systems in order to safeguard against potential system failures.
is Many problems in achieving variable tension on manual controllers are unique to a tilt rotor aircraft because it functions as both an airplane and a helicopter. Because a tilt rotor aircraft operates as both, it is desirable to have the feel of the tilt rotor aircraft change as it is converted from an airplane to a helicopter, and vice a versa, during flight. The way s that the 'feel' and resulting handling capabilities are accomplished currently in tilt rotor aircraft systems (such as the Bell XV1S and the V22 tilt rotor aircraft), is to use the heavier, more expensive variable force field actuator systems, as described above. It would be more desirable in tilt rotor aircraft applications, and for the aircraft industry as a whole, to have access to a less complicated, lighter and more reliable variable gradient cyclic force feel system, such as disclosed in the present invention.
SUMMARY OF THE INVENTION
In accordance with one aspect of the present invention there is provided a gradient actuator operationally coupled to lateral and longitudinal springs of a latitudinal and longitudinal control mechanism of a tilt rotor aircraft, the gradient actuator further coupled to a Nacelle position sensor via a controller for tracking the Nacelle position of the tilt rotor aircraft during operation so as to provide a force feel to a pilot of the aircraft according to the position of the Nacelle, wherein the gradient actuator is controlled by the controller to change the moment arm length of a lateral spring cartridge assembly with respect to a lateral actuator and a longitudinal spring cartridge assembly with respect to a longitudinal actuator, the change in moment arm length being based on input to the gradient actuator from the controller, the controller receiving a position signal from the Nacelle position sensor.
In accordance with one aspect of the present invention there is provided a gradient actuator operationally coupled to lateral and longitudinal springs of a latitudinal and longitudinal control mechanism of a tilt rotor aircraft, the gradient actuator further coupled to a Nacelle position sensor via a controller for tracking the Nacelle position of the tilt rotor aircraft during operation so as to provide a force feel to a pilot of the aircraft according to the position of the Nacelle, wherein the gradient actuator is controlled by the controller to change the moment arm length of a lateral spring cartridge assembly with respect to a lateral actuator and a longitudinal spring cartridge assembly with respect to a longitudinal actuator, the change in moment arm length being based on input to the gradient actuator from the controller, the controller receiving a position signal from the Nacelle position sensor.
BRIEF DESCRIPTION OF THE DRAWINGS
The foregoing and other objects, aspects and advantages are better understood from the following detailed description of a preferred embodiment of the invention with reference to the drawings, in which:
Figure 1 is a perspective view of a cyclic control stick mechanism for a helicopter incorporating the variable gradient control system of the present invention;
Figure 2 is a plan view of the position of the gradient to actuator, linkages and moments arms when the aircraft is in helicopter mode;
Figure 3 is a plan view of the position of the gradient actuator, linkages and moment arms when a tilt rotor aircraft is in airplane mode;
is Figure 4 is a superimposed plan view of the gradient actuator, linkages and moments arms in airplane mode over helicopter mode;
Figure 5 is a graphical illustration of gradient force on the y axis, plotted against a Nacelle angle on the x axis; and Figure 6 is a schematic illustration of interacting components 20 for one system configuration given the teachings of the present invention.
The foregoing and other objects, aspects and advantages are better understood from the following detailed description of a preferred embodiment of the invention with reference to the drawings, in which:
Figure 1 is a perspective view of a cyclic control stick mechanism for a helicopter incorporating the variable gradient control system of the present invention;
Figure 2 is a plan view of the position of the gradient to actuator, linkages and moments arms when the aircraft is in helicopter mode;
Figure 3 is a plan view of the position of the gradient actuator, linkages and moment arms when a tilt rotor aircraft is in airplane mode;
is Figure 4 is a superimposed plan view of the gradient actuator, linkages and moments arms in airplane mode over helicopter mode;
Figure 5 is a graphical illustration of gradient force on the y axis, plotted against a Nacelle angle on the x axis; and Figure 6 is a schematic illustration of interacting components 20 for one system configuration given the teachings of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
One preferred embodiment of a system in accordance with the present invention is practiced with the lateral and longitudinalitudinal pilot-actuated control system for a tilt rotor s aircraft. The operation of a standard force trim system in a tilt rotor aircraft will - first be generally discussed in order to provide a frame of reference to compare the benefits of the invention. It should be understood from the teachings of the invention that it can be applied generally to all manual control to stick system across the entire aircraft industry. For purposes of this description, the terms 'control stick' or 'stick' are meant to generically apply to manual control systems (e.g, control sticks, cyclic, steering mechanisms, etc.) commonly found in the aircraft (helicopter, airplane and tilt rotor aircraft) industry.
is If a pilot took the control stick in a conventional force trim helicopter system and moved.it one inch, he would encounter about one pound of force. If the pilot moved the same control stick 2 inches, he would encounter about 2 pounds of force. The same force would be encountered in either left or right movement of the stick.
20 If the pilot is flying a standard airplane and moved the control stick therein either forward or backward one inch, he would encounter about 2 pounds per inch of force. If the control stick were moved 2 inches, he would encounter 3-4 pounds per square inch of force. In a conventional airplane, the pilot is encountering 25 about 2 pounds of force for every inch of controller movement. In a helicopter, however, the force per square inch relationship is much more linear.
When a pilot flies in a airplane, he desires a much stiffer feel over the control-stick controller than in a helicopter. The same high tension force is not be desirable in a helicopter where s faster movement and mechanical response is desired. A tilt rotor aircraft requires the. ability to do both in order to have variable force. Rather than solely using an electric motor to artificially place a higher torque on the controller, resulting in a higher-tensioned feel, the present invention uses a variable gradient actuator in combination with a three-dimensional phasing linkage to cause a moment arm on which the typical spring cartridge mechanism is attached to change its distance with respect to the Latitudinal/Longitudinal adjustment mechanism for the aircraft and stick. Based on the simple engineering principle 'moment is equal is to force times length a moment of one foot pound is equal to one foot moment arm times one pound of force', instead of changing the force on the stick electronically through motors, a mechanical variance in the moment arm relationship to the directional hardware and/or pilot control mechanism is changed.
Referring to Figure 1, a perspective view of the pilot operated section of the latitudinal and longitudinal control system 10 .in an aircraft incorporating the force feel adjustment improvement of the present invention is illustrated. The system 10 includes two spring cartridges, a lateral spring cartridge 5 and a longitudinal spring cartridge 6. The lateral spring cartridge 5 is tied to the control sticks 2, and controls the lateral motion of the sticks 2 (side to side motion). Through an additional linkage, the sticks 2 are tied to the longitudinal spring cartridge 6, which provides longitudinal motion (forward and back motion) to the sticks 2. From the perspective view of the Figure, it can be seen that lateral spring cartridge 5 is attached to a bell crank 7 that is linked at 9 and 10 to the sticks 2 for lateral motion. The spring cartridge's 5 opposite end, or what is referred to as ground end, is attached to a lateral trim actuator 15. The lateral trim actuator 15 may allow a pilot to reset to zero, or neutral force, to the sticks 2 position by using a beep switch 3 located on the sticks 2. Reset causes the bell crank 7 on the actuator 15 to move, changing the systems zero point. As the bell crank 7 moves, the sticks 2 move along with it such that a new zero point would be achieved. The same is true for longitudinal functions. From the is Figure, it can be seen that the longitudinal spring cartridge 6 is also associated with a longitudinal trim actuator 16. The longitudinal spring cartridge 6 and longitudinal trim actuator are linked with the sticks 2 and associated with the beep switch 3, as indicated above.
20 The actuators 15, 16 have a clutch (not shown), which controls, or clutches, the movement of each associated bell crank in and out. No spring force is required at all in some flight modes such as helicopter applications, which require very responsive stick action for precise aircraft motion. In such an 25 application, the bell cranks are actually declutched, allowing the sticks 2, 3 to move together against minimal force. Free motion, however, is completely unacceptable for a tilt rotor aircraft operating in airplane mode because a small cyclic reflection could cause very dramatic aircraft motion. The clutch for each actuator must therefore be disabled in airplane mode.
s For a control stick system to meet the requirements of the tilt rotor, it must.., have a control stick force that increases with increased speed when converting from helicopter to airplane (otherwise known as a variable gradient force). The invention, through the linkage arrangement described herein, changes the moment arm length of the lateral and longitudinal springs 5, 6 through a gradient actuator 17 and linkage arrangement 18, 19 to the lateral and longitudinal springs 5, 6, respectively.
Referring to Figure 2, what is illustrated is a plan view of the gradient actuator 17 and its associated linkages 18, 19 to the is lateral 5 and longitudinal 6 springs. The moment arm 21 for the roll, or lateral, axis (or lateral setting) is approximately 2.2 inches in Helicopter mode. The moment arm 22 for the pitch axis (or longitudinal setting) is approximately 1.8 inches.
There is a need for higher force on the sticks 2 when a tilt rotor aircraft must move into airplane mode. Referring to Figure 3, to accommodate this needed change in force, the gradient actuator 17 increases the length of the moment arms 21, 22 so that the springs 15, 16 can also accommodate the change. Through movement by the gradient actuator 17, the roll axis can be increased to about 3.4 inches and the pitch axis to about 3.44.
Referring to Figure 4, the invention in Airplane mode is shown superimposed over Helicopter mode. The respective positions of the moment arms, 21 and 22 for each mode are what accomplishes the feel and control advantages of the invention.
Referring to Figure 5, a graphical illustration shows what a s pilot may feel as the aircraft is transitioning from airplane mode to helicopter mode.. Two curves illustrate a force that decreases for longitudinal stick position (or pitch) from seven pounds per inch down to only about 2-3/4 pounds per inch as the angle of the tilt rotors, with respect to the horizon, is increased (or as the tilt rotor aircraft is otherwise moved from airplane mode into helicopter mode). The pilot will encounter a lateral feel that undergoes a similar change, from about 3-1/2 pounds in force in airplane mode down to about 1.8 pounds of force in helicopter mode.
The size of the change may be made dependent on what is desired by is the pilot. It is conceivable that the load on the cyclic stick could actually go down to 0 by having the moment arms move to a 0 moment arm length, if such a change were desired by the pilot. Such diverse operation would be coordinated by the controller, through the gradient actuator 17.
Referring to Figure 6, sensors 61, 62, located in each Nacelle 53,54 of the tilt rotor aircraft (not shown) send information to the controller 67, which then causes the gradient actuator 17 to make the necessary adjustments to the moment arms 21, 22. In addition to Nacelle placement, the controller 67 can also receive input based on air speed 63, which would be used to determine moment arm placement. Tailored pilot settings may also be input to the controller 67, manually 64 aad/or from memory 65. Presently, however, because of the complexity of measuring air speed and the possibility for controller failure or pilot miscalculation, it is most simple and reliable to tie moment arm adjustments directly to s the Nacelle angle. Such an arrangement could conceivably be made with minimal electronic control by slaving the gradient actuator 17 to sensors /transducers 61 or 62 located at either Nacelle.
Furthermore, the Nacelle angle can be sensed very reliably and redundantly with the placement of different independent sensors throughout the system.
Because both lateral and longitudinal system force feel values are set by the same parameter, a gradient actuator can be used to vary the force gradient of both systems. Great benefit is derived through the use of the lateral and longitudinal three-dimensional is phasing linkages driven by a single gradient actuator as described herein. This configuration results in a simple, light-weight system. It should be appreciated that if independent varying of lateral and longitudinal force feel values is desired, two separate actuators in response to signals from a control mechanism as described herein can also be used.
The variable gradient system of the present invention takes the conventional system and varies the length of the moment arm.
This is accomplished by using a single actuator (which is not independently back-drivable) to pivot the moment arms for the is lateral and longitude motions of the stick so that desired tensions are accomplished. The conventional fixed gradient force trim system described in the background, for example, can be converted into a variable gradient force adjustment system with the application of a three-dimensional phasing linkage arrangement. By s adding hinge points to the bell cranks of a conventional force trim system, the plane of the two dimensional linkage can be rotated via a gradient actuator linked to the hinge points. As the operating plane of the two dimensional linkage is rotated out of the original plane of operation, the effective moment arm of the bell cranks to with respect to the gear is reduced. The reduction in effective moment arm reduces the amount the spring cartridge is compressed or extended by the cyclic. The reduction continues until the linkage is rotated to about 90 from its original position. At about 90 the effective moment arm is zero and cyclic movement has minimal is effect on the spring cartridge. In effect, the three dimensional linkage system allows the cyclic force felt by the pilot to be continuously phased to zero force.
A pilot of a tilt rotor aircraft flying in helicopter mode can now move into airplane mode and realize a gradual increase of 20 controller force and stability. The present system is simple, relative to current systems, in that the redundancy required by most force trim systems be overcome, or otherwise eliminated. An aircraft using the present system would only lose the variable gradient force should the gradient actuator fail, leaving the 25 controller in one force position. Such a condition wouldn't change with speed of the aircraft, but would be a much more benign failure than having a system that goes completely limp. Furthermore, from the standpoint of cost, weight, reliability and simplicity, the present invention is a major improvement over current systems.
While the invention has been described in detail above, it should be understood that it has been presented by way of example only, and not limitation. Thus, the breadth and scope of a preferred embodiment should not be limited by any embodiments described above, but should be defined only in accordance with the io following claims and their equivalents.
One preferred embodiment of a system in accordance with the present invention is practiced with the lateral and longitudinalitudinal pilot-actuated control system for a tilt rotor s aircraft. The operation of a standard force trim system in a tilt rotor aircraft will - first be generally discussed in order to provide a frame of reference to compare the benefits of the invention. It should be understood from the teachings of the invention that it can be applied generally to all manual control to stick system across the entire aircraft industry. For purposes of this description, the terms 'control stick' or 'stick' are meant to generically apply to manual control systems (e.g, control sticks, cyclic, steering mechanisms, etc.) commonly found in the aircraft (helicopter, airplane and tilt rotor aircraft) industry.
is If a pilot took the control stick in a conventional force trim helicopter system and moved.it one inch, he would encounter about one pound of force. If the pilot moved the same control stick 2 inches, he would encounter about 2 pounds of force. The same force would be encountered in either left or right movement of the stick.
20 If the pilot is flying a standard airplane and moved the control stick therein either forward or backward one inch, he would encounter about 2 pounds per inch of force. If the control stick were moved 2 inches, he would encounter 3-4 pounds per square inch of force. In a conventional airplane, the pilot is encountering 25 about 2 pounds of force for every inch of controller movement. In a helicopter, however, the force per square inch relationship is much more linear.
When a pilot flies in a airplane, he desires a much stiffer feel over the control-stick controller than in a helicopter. The same high tension force is not be desirable in a helicopter where s faster movement and mechanical response is desired. A tilt rotor aircraft requires the. ability to do both in order to have variable force. Rather than solely using an electric motor to artificially place a higher torque on the controller, resulting in a higher-tensioned feel, the present invention uses a variable gradient actuator in combination with a three-dimensional phasing linkage to cause a moment arm on which the typical spring cartridge mechanism is attached to change its distance with respect to the Latitudinal/Longitudinal adjustment mechanism for the aircraft and stick. Based on the simple engineering principle 'moment is equal is to force times length a moment of one foot pound is equal to one foot moment arm times one pound of force', instead of changing the force on the stick electronically through motors, a mechanical variance in the moment arm relationship to the directional hardware and/or pilot control mechanism is changed.
Referring to Figure 1, a perspective view of the pilot operated section of the latitudinal and longitudinal control system 10 .in an aircraft incorporating the force feel adjustment improvement of the present invention is illustrated. The system 10 includes two spring cartridges, a lateral spring cartridge 5 and a longitudinal spring cartridge 6. The lateral spring cartridge 5 is tied to the control sticks 2, and controls the lateral motion of the sticks 2 (side to side motion). Through an additional linkage, the sticks 2 are tied to the longitudinal spring cartridge 6, which provides longitudinal motion (forward and back motion) to the sticks 2. From the perspective view of the Figure, it can be seen that lateral spring cartridge 5 is attached to a bell crank 7 that is linked at 9 and 10 to the sticks 2 for lateral motion. The spring cartridge's 5 opposite end, or what is referred to as ground end, is attached to a lateral trim actuator 15. The lateral trim actuator 15 may allow a pilot to reset to zero, or neutral force, to the sticks 2 position by using a beep switch 3 located on the sticks 2. Reset causes the bell crank 7 on the actuator 15 to move, changing the systems zero point. As the bell crank 7 moves, the sticks 2 move along with it such that a new zero point would be achieved. The same is true for longitudinal functions. From the is Figure, it can be seen that the longitudinal spring cartridge 6 is also associated with a longitudinal trim actuator 16. The longitudinal spring cartridge 6 and longitudinal trim actuator are linked with the sticks 2 and associated with the beep switch 3, as indicated above.
20 The actuators 15, 16 have a clutch (not shown), which controls, or clutches, the movement of each associated bell crank in and out. No spring force is required at all in some flight modes such as helicopter applications, which require very responsive stick action for precise aircraft motion. In such an 25 application, the bell cranks are actually declutched, allowing the sticks 2, 3 to move together against minimal force. Free motion, however, is completely unacceptable for a tilt rotor aircraft operating in airplane mode because a small cyclic reflection could cause very dramatic aircraft motion. The clutch for each actuator must therefore be disabled in airplane mode.
s For a control stick system to meet the requirements of the tilt rotor, it must.., have a control stick force that increases with increased speed when converting from helicopter to airplane (otherwise known as a variable gradient force). The invention, through the linkage arrangement described herein, changes the moment arm length of the lateral and longitudinal springs 5, 6 through a gradient actuator 17 and linkage arrangement 18, 19 to the lateral and longitudinal springs 5, 6, respectively.
Referring to Figure 2, what is illustrated is a plan view of the gradient actuator 17 and its associated linkages 18, 19 to the is lateral 5 and longitudinal 6 springs. The moment arm 21 for the roll, or lateral, axis (or lateral setting) is approximately 2.2 inches in Helicopter mode. The moment arm 22 for the pitch axis (or longitudinal setting) is approximately 1.8 inches.
There is a need for higher force on the sticks 2 when a tilt rotor aircraft must move into airplane mode. Referring to Figure 3, to accommodate this needed change in force, the gradient actuator 17 increases the length of the moment arms 21, 22 so that the springs 15, 16 can also accommodate the change. Through movement by the gradient actuator 17, the roll axis can be increased to about 3.4 inches and the pitch axis to about 3.44.
Referring to Figure 4, the invention in Airplane mode is shown superimposed over Helicopter mode. The respective positions of the moment arms, 21 and 22 for each mode are what accomplishes the feel and control advantages of the invention.
Referring to Figure 5, a graphical illustration shows what a s pilot may feel as the aircraft is transitioning from airplane mode to helicopter mode.. Two curves illustrate a force that decreases for longitudinal stick position (or pitch) from seven pounds per inch down to only about 2-3/4 pounds per inch as the angle of the tilt rotors, with respect to the horizon, is increased (or as the tilt rotor aircraft is otherwise moved from airplane mode into helicopter mode). The pilot will encounter a lateral feel that undergoes a similar change, from about 3-1/2 pounds in force in airplane mode down to about 1.8 pounds of force in helicopter mode.
The size of the change may be made dependent on what is desired by is the pilot. It is conceivable that the load on the cyclic stick could actually go down to 0 by having the moment arms move to a 0 moment arm length, if such a change were desired by the pilot. Such diverse operation would be coordinated by the controller, through the gradient actuator 17.
Referring to Figure 6, sensors 61, 62, located in each Nacelle 53,54 of the tilt rotor aircraft (not shown) send information to the controller 67, which then causes the gradient actuator 17 to make the necessary adjustments to the moment arms 21, 22. In addition to Nacelle placement, the controller 67 can also receive input based on air speed 63, which would be used to determine moment arm placement. Tailored pilot settings may also be input to the controller 67, manually 64 aad/or from memory 65. Presently, however, because of the complexity of measuring air speed and the possibility for controller failure or pilot miscalculation, it is most simple and reliable to tie moment arm adjustments directly to s the Nacelle angle. Such an arrangement could conceivably be made with minimal electronic control by slaving the gradient actuator 17 to sensors /transducers 61 or 62 located at either Nacelle.
Furthermore, the Nacelle angle can be sensed very reliably and redundantly with the placement of different independent sensors throughout the system.
Because both lateral and longitudinal system force feel values are set by the same parameter, a gradient actuator can be used to vary the force gradient of both systems. Great benefit is derived through the use of the lateral and longitudinal three-dimensional is phasing linkages driven by a single gradient actuator as described herein. This configuration results in a simple, light-weight system. It should be appreciated that if independent varying of lateral and longitudinal force feel values is desired, two separate actuators in response to signals from a control mechanism as described herein can also be used.
The variable gradient system of the present invention takes the conventional system and varies the length of the moment arm.
This is accomplished by using a single actuator (which is not independently back-drivable) to pivot the moment arms for the is lateral and longitude motions of the stick so that desired tensions are accomplished. The conventional fixed gradient force trim system described in the background, for example, can be converted into a variable gradient force adjustment system with the application of a three-dimensional phasing linkage arrangement. By s adding hinge points to the bell cranks of a conventional force trim system, the plane of the two dimensional linkage can be rotated via a gradient actuator linked to the hinge points. As the operating plane of the two dimensional linkage is rotated out of the original plane of operation, the effective moment arm of the bell cranks to with respect to the gear is reduced. The reduction in effective moment arm reduces the amount the spring cartridge is compressed or extended by the cyclic. The reduction continues until the linkage is rotated to about 90 from its original position. At about 90 the effective moment arm is zero and cyclic movement has minimal is effect on the spring cartridge. In effect, the three dimensional linkage system allows the cyclic force felt by the pilot to be continuously phased to zero force.
A pilot of a tilt rotor aircraft flying in helicopter mode can now move into airplane mode and realize a gradual increase of 20 controller force and stability. The present system is simple, relative to current systems, in that the redundancy required by most force trim systems be overcome, or otherwise eliminated. An aircraft using the present system would only lose the variable gradient force should the gradient actuator fail, leaving the 25 controller in one force position. Such a condition wouldn't change with speed of the aircraft, but would be a much more benign failure than having a system that goes completely limp. Furthermore, from the standpoint of cost, weight, reliability and simplicity, the present invention is a major improvement over current systems.
While the invention has been described in detail above, it should be understood that it has been presented by way of example only, and not limitation. Thus, the breadth and scope of a preferred embodiment should not be limited by any embodiments described above, but should be defined only in accordance with the io following claims and their equivalents.
Claims (8)
1. A gradient actuator operationally coupled to lateral and longitudinal springs of a latitudinal and longitudinal control mechanism of a tilt rotor aircraft, the gradient actuator further coupled to a Nacelle position sensor via a controller for tracking the Nacelle position of the tilt rotor aircraft during operation so as to provide a force feel to a pilot of the aircraft according to the position of the Nacelle, wherein the gradient actuator is controlled by the controller to change the moment arm length of a lateral spring cartridge assembly with respect to a lateral actuator and a longitudinal spring cartridge assembly with respect to a longitudinal actuator, the change in moment arm length being based on input to the gradient actuator from the controller, the controller receiving a position signal from the Nacelle position sensor.
2. The actuator according to claim 1, wherein the gradient actuator is operationally coupled to the lateral spring cartridge assembly and operationally coupled to the longitudinal spring cartridge assembly.
3. The actuator according to claim 1, wherein the gradient actuator is configured to provide variable tension to a control stick of the aircraft via the change in moment arm length of the lateral spring cartridge assembly with respect to the lateral actuator and the change in moment arm length of the longitudinal spring cartridge assembly with respect to the longitudinal actuator.
4. The actuator according to claim 1, wherein the lateral and longitudinal spring cartridges are operationally coupled to the control stick to control tension against a lateral and longitudinal motion of the stick.
5. The actuator according to claim 1, wherein the controller is further configured to receive an operational parameter from at least one of the aircraft's onboard devices.
6. The actuator according to claim 5, wherein the operational parameter includes aircraft airspeed.
7. The actuator according to claim 1, wherein the operational parameter is manually input by a pilot.
8. The actuator according to claim 1, wherein the operational parameter is input to the controller from a memory.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US09/369,948 US6254037B1 (en) | 1999-08-06 | 1999-08-06 | Variable gradient control stick force feel adjustment system |
| US09/369,948 | 1999-08-06 | ||
| CA002380346A CA2380346C (en) | 1999-08-06 | 2000-08-04 | Variable gradient control stick force feel adjustment system |
Related Parent Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA002380346A Division CA2380346C (en) | 1999-08-06 | 2000-08-04 | Variable gradient control stick force feel adjustment system |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CA2591645A1 CA2591645A1 (en) | 2001-02-15 |
| CA2591645C true CA2591645C (en) | 2010-10-12 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA2591645A Expired - Lifetime CA2591645C (en) | 1999-08-06 | 2000-08-04 | Variable gradient control stick force feel adjustment system |
Country Status (1)
| Country | Link |
|---|---|
| CA (1) | CA2591645C (en) |
-
2000
- 2000-08-04 CA CA2591645A patent/CA2591645C/en not_active Expired - Lifetime
Also Published As
| Publication number | Publication date |
|---|---|
| CA2591645A1 (en) | 2001-02-15 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| EEER | Examination request | ||
| MKEX | Expiry |
Effective date: 20200804 |