GB2560181B

GB2560181B - Swiveling tandem rotorcraft

Info

Publication number: GB2560181B
Application number: GB1703355.6A
Authority: GB
Inventors: Michael Tapper Paul
Original assignee: Individual
Current assignee: Individual
Priority date: 2017-03-02
Filing date: 2017-03-02
Publication date: 2019-11-13
Anticipated expiration: 2037-03-02
Also published as: GB2560181A; GB201703355D0

Description

SWIVELING TANDEM ROTORCRAFT

BACKGROUND OF THE INVENTION

Field Of Invention

The present invention relates to the design of tandem rotorcraft. A tandem rotorcraft could be a helicopter or an autogyro. The invention is independent of the size and purpose of the tandem rotorcraft.

Prior Art

The design and construction of aircraft is an extensively studied field with well defined terminology. The physics involved in aircraft and more particularly rotorcraft flight can be complex and counter-intuitive but is also well understood and analyzed in the literature.

Detailed discussion of aircraft design concepts well known by those skilled in the art is beyond the scope of this section. This section seeks only to highlight those concepts relevant to the current invention in order to put the current invention into a relevant context, to establish the extent and meaning of the terminology used in this document and to motivate the invention.

Aircraft Types and Terminology

Most current aircraft can be classified as either fixed-wing or rotary-wing. A fixed-wing aircraft generates lift using wings which are fixed to the main structure of the aircraft whereas a rotary-wing aircraft generates lift using one or more wings or rotor blades which are rotating with respect to the main structure of the aircraft.

On a rotary-wing aircraft, groups of one or more wings or rotor blades sweep out surfaces as they rotate. These surfaces can be approximated as disks or cones with respect to the main body of the aircraft. Each disk or cone relates to one or more wings or rotor blades and these wings or rotor blades together with the attached mechanics with which they rotate will be referred to herein as a rotor. A rotor may be comprised of a single structure such as a propeller or it may be a more complex mechanical assembly including for example multiple rotor blades as separate components all attached to a central hub. A rotor is normally mounted on a rotor shaft which is not counted as part of the rotor. A rotary-wing aircraft may have just one rotor but typically has several. One common arrangement, referred to herein as a “conventional helicopter” is a main rotor and a smaller tail rotor, the tail rotor being mounted on an extended boom with the tail rotor disk at right angles to that of the main rotor. Other arrangements include tandem rotors, coaxial rotors and quadcopters. A helicopter is a is a rotary-wing aircraft where motive power is supplied to the rotor or rotors whereas an autogyro is a rotary-wing aircraft which does not use a powered rotor to provide lift. Rotor rotation on an autogyro is sustained instead by aerodynamic forces. An autogyro can be powered using thrust from a propeller. If one or more propellers are used to provide predominantly thrust as the aircraft’s main source of power with lift being generated by unpowered rotors, then the aircraft is considered to be an autogyro and any propeller producing thrust is not generally considered to be a rotor. This is partly a matter of convention. Another example would be a conventional helicopter tail rotor which also provides thrust rather than lift, however this thrust is for stabilisation rather than motive power and the tail rotor is considered to be a rotor. Some helicopters use rotating blades or weights which are smaller than the rotor blades as part of a stabilization scheme. One such mechanism is called a flybar. A flybar or other rotating system used for stabilization in this way is also not counted as a separate rotor.

Many types of helicopter are capable of controlled descending flight and landing without power being supplied to any of their rotors or propellers. This is called autorotation. In autorotation, the main rotor or rotors work in a similar manner to rotors on an autogyro.

Rotorcraft is another word for rotary-wing aircraft. The term rotorcraft includes autogyros, helicopters in powered flight and helicopters in autorotation.

Tandem Rotorcraft A tandem rotorcraft is a rotorcraft with two main rotors. The main rotors are normally identified as being the largest in size and as being those providing most of the lift. A tandem rotorcraft could include further rotors or propellers, for example for providing thrust rather than lift. This thrust could be for motive power or additional stabilization. The term “tandem rotorcraft” includes but is not necessarily limited to tandem autogyros and tandem helicopters. As discussed previously, systems such as flybars are not counted as separate rotors. In most cases it should be clear to those skilled in the art and considering the above discussion whether or not an aircraft is a tandem rotorcraft, perhaps with additional systems, or a different type of aircraft.

Rotorcraft Control

Rotorcraft can be controlled in flight by a variety of means. This includes but is not limited to applying control to individual rotors and any other controllable aerodynamic surfaces. The combined and coordinated control of each of the individual rotors and other controllable surfaces provides a means to control the aircraft as a whole. Depending on the stability of the aircraft design, mechanical or electronic stabilisation systems or a combination of both may be required in order to make the aircraft easier to control by either a human pilot or by some higher level piloting system. Methods for applying flight controls to individual rotors are discussed below.

Control of a Fixed Pitch Rotor

Some rotors have rotor blades which have a fixed pitch. This means the chords of blade elements within the rotor are broadly fixed with respect to the hub to which they are attached. Such fixed pitch rotor blades may flex and twist to some degree when under load but the rotor would still generally be classed as what will be referred to herein as a fixed pitch rotor. With a fixed pitch rotor, the power supplied to the rotor is one means of control. Another means of control with some rotorcraft designs is varying the angle of the entire rotor with respect to the main body of the aircraft. This could be done by swivelling the rotor about an axis or using gimbals to rotate the rotor using several axes of rotation. By using either or both of these techniques a variety of rotorcraft designs are possible.

Control of a Variable Pitch Rotor

Many rotorcraft have rotors in which the pitch of the individual rotor blades varies in a controlled way as the blade rotates around the rotor disk or cone. This movement is also known as feathering. The pitch, as a function of the circular position of the blade, can be varied using an arm suitably attached to the blade and linked to a swashplate. An alternative mechanism of varying the pitch could also be used. Any such pitch control mechanism could be directly controlled via inputs from the pilot or piloting system or it could be indirectly controlled via a mechanical or electronic stabilisation system, or by some mixture of any of these things.

The control of blade pitch as a function of blade position around the rotor disk or cone can be separated into two components: collective pitch and cyclic pitch. Collective pitch is the mean pitch for the blade or blades calculated across all possible blade positions at a given time. This is a single value which can be varied over time by the control system if it is adjustable for a given rotor design. Cyclic pitch is the pitch deviation from the collective pitch and is a function of blade position at a given time. Although this could be a complicated function in general, for most helicopters it is determined from two independent input variables. These two input variables could be expressed in many ways, one of which is a roll input and a pitch input. The roll input relates to a demand for the rotorcraft to rotate about an axis running from the front of the rotorcraft to the rear of the rotorcraft and the pitch input relates to a demand for the rotorcraft to rotate around an axis running from the port side of the rotorcraft to the starboard side of the rotorcraft.

For a conventional helicopter with a single main rotor and a single tail rotor, the main rotor would be controllable in terms of a single collective pitch control input and two cyclic pitch control inputs. The tail rotor has only a collective pitch control input. Conventional tandem helicopters have collective and cyclic pitch control on both main rotors giving a total of six independent rotor control inputs. It will be understood that we are discussing only the principal controls of general classes of aircraft and that for particular aircraft designs further controls may be provided as necessary to provide fine tuning, trimming, speed related adjustment or other such adjustments without materially affecting the aircraft type or basic controls at the level being considered here.

Rotor Blade Flapping and Dragging

Many rotor designs allow rotor blades to hinge upwards and downwards as they rotate. If the blades are hinged near the rotor head then in a normal upright hover the blades all tend approximately to follow the surface of a cone as they rotate. The upwards and downwards movement is known as flapping and the tendency for independently flapping blades to sweep out a cone-like shape when supporting a load is called coning. As cyclic pitch is applied the shape of this cone is distorted in various directions so as to alter the direction of the resultant force exerted by the rotor on the main body of the rotorcraft.

If such hinges are not provided the rotor blades move around a surface which is to a first approximation a disk. A degree of coning may still take place due to the flexibility of the rotor blades. This lesser effect will not be referred to as coning hereinafter and the shape swept out by the blades in this situation will be referred to as a disk. A rotor with two opposing blades may be designed so that rather than each blade being hinged separately, one blade is able to tilt upwards provided the opposing blade tilts downwards in a seesaw arrangement, pivoting about the centre of the rotor. This allows the rotor disk to tilt in any direction but it must maintain broadly a disk rather than a cone shape. This is also known as a semi-ridged rotor. Rotors designed in this way are more suitable for aerobatic and inverted flight than those with independently hinged blades where the blades will form a cone. This is because if a rotorcraft with independently hinged blades were to hover inverted the tips of the blades and the wide part of cone would be towards the aircraft fuselage relative to the rotor head. Hence significant additional clearance would be needed between the rotor and the aircraft fuselage for inverted flight in comparison to that needed for upright flight. This in turn would move the centre of gravity further away from the rotor disk which would tend to produce less desirable aerobatic characteristics.

We note that blades can also be allowed another degree of freedom, referred to as dragging, where blades are allowed to pivot from the blade root and rotor head so that the rotor blade tips can move forwards and backwards in their normal direction of rotation. It is assumed hereinafter that such a degree of freedom may be provided in all the rotor types discussed. Rotors with blades which can both flap and drag via hinges and independently from any other blade will be referred to as fully articulated rotors.

It is also assumed hereinafter that all mechanisms providing freedom of movement for rotor blades may be equipped with springs, dampers, friction devices, end-stops and other appropriate mechanics.

Angular Momentum

The moment of inertia of a rotor blade increases in proportion to the square of its length as well as in proportion to its mass and rotational velocity. When controlling rotors with blades which are relatively long, as well as having significant mass and speed such as those on a conventional helicopter, gyroscopic effects are highly significant. When considering whether the angular momentum is relatively high or low its magnitude should be considered in comparison to the various moments of inertia of the aircraft body. As with a gyroscope, rotating such a rotor, when operational and other than about its axis of operation, requires or generates torque components which are perpendicular to those which would be required or generated if it were not operational. Although counter-intuitive, such effects are widely known and understood and will not be discussed in mathematical detail here. It can be observed or deduced however, that control principals and methods relevant to rotors with relatively long blades and relatively high angular momentum can be significantly different from the those principals and methods relevant to smaller diameter rotors, perhaps based on propellers. In the latter case it may be feasible to control a rotorcraft by pointing such rotors with motors on the fuselage whilst in the former case it may not be.

Torque reaction

Tandem rotorcraft with parallel rotor shafts and identical but counter-rotating rotors rotating at the same speed can benefit from having no net torque reaction when said rotors are driven by a motor or engine. This is one of the principal reasons a tail rotor is not generally required for tandem rotorcraft.

Advantages of Tandem Rotorcraft

Current tandem rotorcraft tend to have superior load carrying abilities compared to alternative helicopter designs of a similar scale and complexity. They also have the significant advantage of not requiring a tail rotor. The use of a tail rotor on a helicopter poses significant risks because of its proximity to the ground during take-off and landing, the degree to which it necessarily protrudes from the main body of the rotorcraft and consequently the ease with which it can be damaged combined with the serious nature of possible consequences of such damage.

Limitations of Current Tandem Rotorcraft

Disadvantages of current designs of tandem rotorcraft include relatively poor manoeuvrability in yaw and the need for a relatively complex rotor design to achieve acceptable manoeuvrability.

These factors may dictate that a rotor with independently flapping blades be employed which in turn is one of the reasons most current tandem rotorcraft are not capable of significant aerobatic flight or even inverted flight.

SUMMARY OF THE INVENTION

The current invention consists of a tandem rotorcraft having two main rotors attached to two main rotor shafts attached to a fuselage. Said fuselage includes means to control the pitch of said two main rotors as they rotate and means to allow said rotor shafts to swivel with respect to each other.

Said fuselage also includes two electric motors used to power each of said main rotors separately. Said swivelling is driven by controlling the pitch of said blades of said main rotors in order to achieve control of said rotorcraft in yaw and said main rotor shafts are synchronized by a control system to accurately control electrical drive signals sent to said motors in concert and to account for relevant feedback signals so that synchronization of said main rotor shafts is achieved without a mechanical connection. In conjunction with an appropriate control system said swivelling movement can be used to improve manoeuvrability with any rotor design and in particular to improve manoeuvrability with simpler rotor designs, hence allowing simpler rotor designs to be employed and the advantages of simpler rotor designs to be exploited along with the inherent advantages of tandem rotorcraft. Using a simpler rotor design could mean for example using a semi-rigid rotor rather than a fully articulated rotor with advantages being for example cost and aerobatic performance. Semi-rigid rotors can be constructed so as not to allow significant movement of the disk swept out in flight by the blades of the rotor in question with respect to the shaft to which it is attached. The feedback system used to synchronise said two motors could make use of feedback information from voltage measurements made on connections to the windings of said motors and measurements made of the position of said rotor shafts using position sensors.

Said tandem rotorcraft could include flight control means for stabilization and control which may use information from sources including all of or any combination of a first three axis gyroscope unit or equivalent system mounted on any part of the fuselage from which the rotation of the first of said main rotor shafts can be measured or a second three axis gyroscope unit or equivalent system mounted on any part of the fuselage from which the rotation of the second of said main rotor shafts can be measured or means to measure the rotational position of the section of the fuselage connected to the first of said main rotor shafts with respect to the section of the fuselage connected to the second of said main rotor shafts.

If said tandem rotorcraft has overlapping blades then said control system to accurately control electrical drive signals sent to said motors could be used to prevent said rotors from colliding. Further rotors for propulsion or stabilization may or may not be included without changing the nature of the invention.

Furthermore a method of the current invention consists of flying any tandem rotorcraft described above. For example a method of flying a rotorcraft consisting of providing a rotorcraft including the swivelling as described and controlling the pitch of said rotor blades in order to control said rotorcraft and controlling electrical drive signals so that synchronisation of the two main rotors is achieved without a mechanical connection.

BREIF DESCRIPTION OF THE DRAWINGS

Figure 1 illustrates the central mechanical elements of one embodiment of the current invention.

Figure 2 illustrates an example timing belt arrangement which could be used in the embodiment of figure 1.

Figure 3 illustrates an example swivelling mechanism which could be used in the embodiment of figure 1.

Figure 4 illustrates an example rotor head design which could be used in the embodiment of figure 1.

Figure 5 illustrates a possible way to describe the motions of the current invention for the purposes of discussion of control methods as well as a possible configuration for sensors to be used in a possible control scheme described below.

Figure 6 illustrates the central mechanical elements of figure 1 with the addition of example rotor blades and showing the two main body sections at an angle to each other as allowed for by the swivelling mechanism.

DETAILED DESCRIPTION

Although the first and second embodiments include mechanical methods to synchronise the rotors it will be understood that such mechanical methods fall outside the scope of the current invention which is synchronised electronically as described in the third embodiment.

Mechanical Structure

The central mechanical structure a first embodiment of the current invention is shown in figure 1. It will also be understood that figure 1 shows only the core components relevant to aspects of the current invention being discussed and that additional structural elements, control elements, landing gear, rotor blades, fuselage coverings and other items well known in the art have been omitted for clarity. Pairs of blade grips (101a, 101b, 101c, 101 d) are attached to each of two rotor heads (102a, 102b) in such a way as to allow said pairs of blade grips to rotate about an axis passing through the centre of said two blade grips. This rotation allows the pitch of attached rotor blades to be varied. Said rotation is also known in the art as feathering. The blades are omitted from figure 1 for clarity but would fit into the slots shown in said blade holders in a standard way. Control of rotor blade pitch is achieved by control arms attached to said blade grips which are connected to swashplates, said swashplates may be controlled by three servo motors per swashplate. The structure of the control arms, swash plates and attached servo motors is well known in the art and is omitted from figure 1 for clarity.

Said two rotor heads (102a, 102b) are rigidly attached to two rotor shafts (103a, 103b) which are attached to main body sections (104a, 104b) of the rotorcraft in such a way as to allow the shafts to rotate freely but otherwise be constrained in movement with respect to said main body sections. Said shafts are driven by motors (105a, 105b) via main drive gears (106a, 106b) and one way bearings (107a 107b). Said one way bearings allow said rotors to be driven in their normal direction of rotation by said motors whilst being free to rotate beyond their driven position in said normal direction. Said two rotor shafts are rigidly attached to toothed pulleys (108a, 108b). Said toothed pulleys are connected by a toothed timing belt (109). Said toothed timing belt runs between said toothed pulleys and incorporates a one hundred and eighty degree twist along this distance This is shown in more detail in figure 2 where said toothed pulleys, said toothed belt and guide rollers (201a, 201b, 201c, 201d, 202) only are shown. We note that guide roller (202) is mounted inside the central structure in figure 1 (110) and hence not visible in figure 1. Said one hundred and eighty degree twist means that toothed pulleys and the attached rotor shafts (103a, 103b) in figure 1 are constrained to move synchronously in opposite directions. Said synchronous movement allows four rotor blades (not shown) attached said blade grips to rotate in overlapping disks in an intermeshing manner without colliding with each other. The two main body sections (104a, 104b) are connected together via a structure (110) which allows the body sections to rotate in relation to each other. Figure 3 shows a cross section and breakdown of structure 110 of figure 1. The tubing on the left hand side of the diagram (301) is rigidly connected to the larger diameter tubing in the centre of the diagram (302) or is part of the same single component. On the right hand side of the diagram another piece of smaller diameter tubing (303) is mounted in roller bearings (304a, 304b) whilst being constrained from sliding into or out of the larger diameter tubing via bushings (305a, 305b) attached to said tubing on the right (303). The tubing on the far left and far right of the diagram (301, 303) is rigidly connected to the main body sections of figure 1 (104a, 104b). It will be understood that this is just one of many mechanical arrangements which might be used to allow two parts of a fuselage to rotate with respect to one another, any of which could be used in other embodiments of the current invention.

Figure 4 shows a cross section of the rotor head used in the current embodiment, this is a cross section of components 101a, 101b and 102a of figure 1. Such a rotor head design is known in the art but described herein as an example of a design of rotor head suitable for use with the current embodiment of the current invention. It will be understood that many other designs could be used. Blade grips (101a, 101b) are each held in place with a pair of roller bearings (401a, 401b), a thrust bearing (402), a spacer (403), a washer (404) and a nut (405) attached to one end of a feathering shaft (406). Said roller bearings and said thrust bearings allow said blade grips to rotate.

Feathering shaft (406) is held within rotor hub (102a) by elastic dampers (407) and associated spacer (408). A further spacer (409) is used around said feathering shaft between said pairs of roller bearings. Said elastic dampers allow the blade grips to move up and down by a small amounts, in particular with the blade grip on one side being able to move in the opposite direction to the blade grip on the other side. Aside from unwanted mechanical play, the longitudinal axes of said blade grips are constrained to be aligned with the longitudinal axis of said feathering shaft, and hence aligned with each other. When fitted with suitable rotor blades this type of rotor sweeps out a disk which can tilt slightly in all directions with respect to the body via the flexing of said elastic dampers but which, neglecting any flexing of the rotor blades and mechanical play in the assembly, does not form a cone. In some designs said elastic dampers are replaced by rigid bushings. This variation is also suitable for use with the current embodiment of the current invention and results in the movement of the main body section to which the rotor disk is ultimately attached following the movement of the rotor disk more closely.

Advantages of the Current Invention

From the discussion of prior art above we note that a rotor with hinged blades sweeping out a steerable cone facilitates yaw control in current large scale tandem rotorcraft designs, but these designs are less suitable for aerobatic flight. They are also more complex mechanically due to the hinging requirements. The simpler tilting disk approach can also be used to achieve yaw control on a tandem rotorcraft however any degree of freedom afforded to the disk in the direction required to yaw the rotorcraft is also available to the disk in all other directions, including towards the fuselage. Hence providing tilting disks capable of powerful yawing moments comes with an increased risk of the rotor blades striking the aircraft fuselage.

The present invention solves this problem by allowing a very simple rotor design to tilt freely in the direction required to yaw the rotorcraft without allowing the rotor to tilt towards the fuselage with the same degree of freedom. Furthermore, the required movement is achieved by a simple low frequency swivelling movement within the main body of the tandem rotorcraft as opposed to a rapid oscillatory movement taking place at the rotor frequency as occurs with tilting movements implemented in the rotor via a see-saw arrangement for example. Indeed the present invention can allow a manoeuvrable tandem rotorcraft to be made with very rigid rotor heads so that the rotor heads provide little or no freedom for the disks to tilt independently of the fuselage components to which they are attached, relying instead on fuselage components following the rotor disks directly. We note also that the two rotors of the current invention could easily be oriented in flight at an angle of for example ninety degrees to each other which would be far in excess of such angles achievable with current designs of tandem rotorcraft.

Further Characteristics of the Current Invention A tandem rotorcraft with identical counter-rotating rotors with fixed blades mounted on two parallel rotor shafts rotating at the same speed can have little or no overall angular momentum in flight. With a ridged body such a tandem rotorcraft would respond to disturbances in a similar way to a craft with no rotating components as overall angular momentum is low in both cases. In allowing said swivelling movement in the current invention we note that the situation changes significantly and we can no longer treat the tandem rotorcraft as an approximately rigid object with negligible overall angular momentum. For example an in-flight disturbance which tends to lift one rotor with respect to the other will result in a tendency for the rotorcraft of the current invention to swivel due to gyroscopic forces. Applying different degrees of collective pitch to the two main rotors to change the attitude of the aircraft, which is a method currently employed in current tandem rotor designs, will have the secondary effect of inducing a swivelling motion also. Hence a different approach must be taken to control the tandem rotorcraft of the current invention to account for the gyroscopic effects and degrees of freedom of the new design in all types of manoeuvre and with all types of disturbance. For example to change the attitude of the tandem rotorcraft with respect to its longitudinal axis a carefully controlled combination of cyclic and collective pitch calculated for each rotor individually can be used.

This requirement to use different control and stabilisation methods clearly differentiates in a fundamental way a tandem rotorcraft of the current invention from those already known in the art. A working example of the current invention cannot be obtained by gradual process of adjusting current tandem rotorcraft designs, rather it must be clearly conceived, understood and designed from a both a mechanical and control point of view at the outset.

It was previously noted that tandem rotorcraft, particularly those with parallel rotor shafts and identical counter-rotating blades can benefit from having no overall torque reaction. We note that zero overall torque reaction is also possible with the current invention when the rotor shafts are aligned but with the rotor shafts in other positions the rotorcraft will be subject to some torque reaction when the rotors are driven. Again, with an appropriate control system compensating for these effects this need not be a problem in flight.

Powering and Synchronization

Although the current invention is independent of the means used to power the tandem rotorcraft, if any, it is envisaged that its use with powered main rotors will be of particular utility. Likewise the current invention does not require the main rotors to be counter-rotating or intermeshing but this is also an embodiment of particular convenience for some applications.

The embodiment described above and shown in figure 1 uses a timing belt, rotated through one hundred and eighty degrees to connect and synchronize the main rotor shafts in a counter-rotating sense.

In a second embodiment, a drive shaft, sometimes known as a torque tube on smaller rotorcraft, is used in place of the drive belt. Such a drive shaft may be connected to the rotor shaft using bevel gears rather than the pulleys used with the belt drive. We note that in this case, with a pair of bevel gears used at each end of the drive shaft, the rotors are synchronized in a counter rotating manner.

Such an embodiment represents a way of employing a drive shaft using a minimal number of components in the same way as the first embodiment represents a way of employing a drive belt with a minimal number of components. In both cases additional gears and intermediate shafts could be employed. In the case of the belt, such additional gearing may be chosen in place of the use of the one hundred and eighty degree twist in the belt, which requires some components to separate the belt along its run.

We note that a drive shaft is conventionally used to transmit torque along a rigid structure.

Counteracting torques are required to allow the shaft to synchronize the rotors rather than rotate the entire assembly as is possible in this case due to the swivelling nature of the rotorcraft fuselage. In this second embodiment, these counteracting torques are provided by cyclic pitch applied to the main rotors as required to maintain the relative orientation of the fuselage components required to generate the selected flight path. This is an additional burden for the control system and must be weighed against the advantages of employing a drive shaft in selecting the best embodiment for the particular application. A third embodiment will now be described which allows the drive shaft, drive belt or other such mechanical means of synchronising the rotor shafts to be eliminated. In this third embodiment a motor, for example a brushless motor is used to drive each of the rotor shafts, possibly via gears as shown in figure 1 (106a, 106b) although a different gearing arrangement, a belt, or direct drive of the shaft with the motor could be used. Unlike the first embodiment described herein, a one way bearing is not used in this embodiment and the rotor shaft and the motor shaft always rotate in a synchronous manner, allowing for the gear ratio employed. Normally brushless motors are driven by electronic speed control units, also known as ESCs. In this embodiment, a novel means of controlling the motors is employed wherein the two ESCs normally used to drive two motors have been combined to form a single logical system designed to synchronize the movement of the motors electronically to a sufficient degree to prevent a collision of rotor blades. This will involve the coordinated but individual control of the motors at the lowest level. The individual motors will need to be fed different amounts of power at different times to maintain synchronisation. The difference in the total angle through which one motor has rotated with respect to the other is tracked and maintained so that it remains below a required value. A standard brushless motor electronic speed controller monitors the position of the motor magnets as they rotate past the statically mounted windings and uses this information to determine exactly when to apply power to the next motor winding or to “commute” the motor. This is often, but not always, done by monitoring the voltage on a connection to the motor coil which is not currently being powered. For example, in a motor with three connections, at any one time when the motor is running two connections may be used to supply power and the third may be monitored to check the position of the magnets with respect to the windings. This information is not usually available outside the ESC unit, nor is it normally used to determine the total rotational movement of the motor, and nor will it provide information about the absolute position any attached rotorcraft rotor. However if such information from both motors is combined with some lower frequency information relating to the absolute positions of the tandem rotorcraft rotors and used in real time to appropriately control said motors it is possible to maintain the synchronization of said rotors by purely electronic means. Such absolute position information can be obtained by for example using a suitable shaft encoder attached to the main shaft or main gear either directly or indirectly or by using one more magnets attached to the shaft or main gear along with one or more hall effect sensors mounted on the body such that the sensor or sensors can detect the passing of the magnet or magnets and hence infer the position of the shaft. The absolute position of the two rotor shafts is primarily useful when starting the rotorcraft if the rotation of the motors itself can be accurately monitored but may also be used as part of the control and monitoring system. Although the current embodiment has been described using the example of a brushless motor it will be understood that other motor technologies could be employed without departing from the spirit of the current embodiment.

Stabilization and Control Arrangements

As mentioned above, the current invention will require a control system which is significantly different to any which might be applicable to other tandem rotorcraft known in the art and there are many ways to implement this. The choice of implementation will depend on the scale and purpose of the rotorcraft and there will be several ways to implement a control system even given a particular scale and purpose. A system is now described which is suitable for a rotorcraft of the current invention measuring approximately one meter in total length for piloting automatically or using remote control. This size has been chosen as an example because it is a convenient size for testing prototype rotorcraft and associated control systems. The system described is also suitable for use with a wide variety of, but not necessarily all, rotorcraft sizes.

Figure 5 shows the mechanical elements of the rotorcraft of figure 1 with the addition of two gyroscopic sensing units (501a, 501b) and a rotational position sensing unit (502). The control system currently being described takes inputs from said two gyroscopic sensing units and said rotational position sensing unit. Said rotational position sensing unit has many possible implementations but the figure illustrates a small potentiometer or rotary shaft encoder rigidly mounted on one section of the rotorcraft with its output shaft connected via a belt to the section of fuselage on the other side of the roller bearings so that rotational movement can be measured. This is a simple method to implement and illustrate but other methods could be used. Said gyroscopic sensing units could be implemented as a micro electro mechanical systems (MEMS) or some other technology. Each gyroscopic sensing unit provides outputs proportional to the rate of angular movement detected along each of three orthogonal axes. These output signals are filtered and can be processed further to derive estimates of the total rotational movement which has taken place along a given axis and also the rotational acceleration about a given axis. These estimates may be derived by integrating and differentiating the original outputs of the gyroscopic sensing units or by the use of other techniques such as Kalman Filtering. Estimates of absolute rotational position derived from outputs of rate sensitive gyroscopic sensors are subject to drift and require an accurate estimate of the initial position to work with. Since the relative orientation of the two rotor shafts is an important parameter in correctly controlling the rotorcraft it is useful to add a position sensor to directly measure this with the required accuracy, hence eliminating drift problems. We note that this is only one approach, and it is possible to gain sufficient information from the two gyroscopic sensors alone given a careful and thorough analysis of the way all movements on all axes relate to each other and to the possible movements of the rotorcraft considering its structure. It is also possible to gain sufficient information by using just one three axis gyroscopic sensor combined with said rotational position sensing unit. There are also various ways in which

information from linear accelerometers, possibly included on the same device as a MEMS gyroscopic sensor, can be combined with the other inputs to improve the accuracy of the measurements of the position, speed and acceleration of the rotorcraft.

In the following discussion we will consider the direction of forward flight to be as shown on Figure 5 (503) and take pitch, roll, and yaw to as having their conventional meanings with respect to this forward direction as also indicated by the central arrows (504). We consider the movements of the overall rotorcraft in pitch, roll and yaw and we will take these to be controlled in terms of pitch, roll and collective inputs to the front and back rotor disks, also as identified in the figure (505a, 505b). In this context, pitch refers to a rotation about a pitch axis rather than any pitch angle of any rotor blade. Applying a pitch, roll or collective input to a rotor disk shall mean moving the swashplate in an appropriate direction so that said rotor disk has a tendency to make said pitching, rolling or directional movement and hence applies a torque (in the case of pitch or roll) or force (in the case of collective) to the fuselage component to which it is attached such that said fuselage component would pitch, roll or be moved appropriately if it were not subject to any other forces or constraints. Using these six control inputs the tandem rotorcraft of the current invention can be controlled to generate the overall pitch, roll and yaw movements as well as swivelling movements between the two fuselage sections required to fly the rotorcraft. In addition to these rotational movements the rotors also create a resultant force which can be used to both manoeuvre and support the aircraft in flight. In this discussion it will be understood that this is not the only way of analysing the forces involved in the flight of a rotorcraft of the current invention but simply a convenient analysis based on the most conventional axes and terminology.

We assume inputs to the control system from the pilot or piloting system will be expressed in terms of requested rates of pitch, roll and yaw plus an additional input for the overall collective pitch plus and additional input for the amount of motive power to be supplied, if the rotorcraft is powered. These would also be the basic control inputs for a conventional helicopter. We note that for the rotorcraft of the current invention it would be possible to take an alternative set of piloting inputs, for example pitch, roll, angle of front rotor with respect to back rotor, collective pitch and motive power, but we will consider control based on the standard set for simplicity. For basic control of the rotorcraft the problem can be broken down into stabilizing and controlling the rotorcraft in pitch, roll and yaw with a closed loop control system with power and average collective pitch controlled directly with open loop control systems.

Control about the roll axis is the most straightforward element of the closed loop system. The control system compares the requested rate of roll to the measured rate of roll and derives a control output which is used to correct the rate of roll of the rotorcraft. The measured rate of roll is the average of the two roll rates measured on the two gyroscopic sensing units. One way to generate said control output to correct the rate of roll of the rotorcraft is to feed the difference between the measured and requested roll rates into a proportional, integral and derivative type controller (PID controller). The control output is used to apply roll to both front and back rotors by the same amount in the appropriate direction.

Control about the pitch axis can be achieved in a similar manner to control about the yaw axis. In this case the measured rate of pitch is the average of the rate of pitch measured by each of the two gyroscopic sensing units and again a PID controller can be employed on this axis. The control signal derived is used to pitch both front and back rotors by the same amount in the appropriate direction. A technique known as differential collective pitch, whereby different amounts of collective pitch are applied to each rotor is commonly used on tandem rotorcraft but is not used in this example because it results in a tendency for the tandem rotorcraft of the current invention to swivel.

Before yaw can be controlled, a controller must first be introduced to control the angle of the two rotor shafts with respect to each other. In this case the angle is being measured directly by a dedicated sensor and a PID controller can again be used to control this angle with respect to a required angle. The difference in measured roll rates between the two gyroscopic sensors also gives information about the rate of change of this angle which can be used as an input to the control algorithm. This may result in a controller which differs from a simple PID controller and analysis and design of such a system can be performed with well known control theory. Having introduced a method of controlling and stabilising the angle between the two rotor shafts, control of the rotorcraft yaw rate can be added to this. The requirement to yaw the rotorcraft is converted into a requirement to adjust the angle between the rotor shafts. The degree to which this is viewed as two separate levels of control system is not significant as either viewpoint could be taken to generate systems which are identical mathematically and we could equivalently view this as a single yaw control system with additional states and feedback paths. The measured yaw rate of the rotorcraft can be calculated from the average of the yaw rates of the two gyroscopic sensors along with the difference in pitch rates from the two gyroscopic sensors. The greater the angle between the rotor shafts the greater the degree to which the yawing of the overall rotorcraft will be measured as pitching by the individual gyroscopic sensors. For small angles, considering the average yaw measured by the gyroscopic sensors may be a sufficient as an approximation. In adding feedback for yaw control two points should be considered: (1) the angle between the two rotor shafts does directly imply a yaw rate, but is rather one factor in the rate of change of the yaw rate (2) the direction of the rate of change of yaw produced by a given offset angle between the rotor shafts changes when the direction of the force exerted on the chassis by the rotor changes.

Hence for example when the rotorcraft is in an inverted hover adjustments to the angle between the two rotor shafts to maintain a given yaw rate as measured by the gyroscopic sensors must be made in the opposite sense to that made when in an upright hover. One way to detect the direction of the force exerted on the fuselage by the rotors is by using accelerometers on the fuselage components. Note that it is the overall direction of the force from the rotor which is important rather than whether or not the rotorcraft is inverted.

Although the control systems described above are sufficient to control a swivelling tandem rotorcraft in upright and inverted flight, further control refinements can improve the efficiency and aerobatic performance of the tandem rotorcraft of the current invention. One such control refinement is the use of opposing cyclic pitching inputs on the front and rear rotors to support rapid yawing movements. For example, when the rotorcraft is upright and rotates rapidly in yaw in either direction with the main shafts offset in the appropriate direction a cyclic pitch up input can be applied to the front rotor, and a cyclic pitch down input can be applied to the rear rotor. This will allow the rotor to more efficiently follow the desired movement without creating unwanted disturbances in the other parts of the control system. Likewise, although differential collective pitch was not used as the primary method to control the pitch of the rotorcraft as discussed above, some appropriate amount of differential cyclic pitch can be introduced as a further refinement to the flight controls.

General

While the invention has been particularly shown and described with reference to particular embodiments, it will be understood by those skilled in the art that various changes in form and detail may be made without departing from the spirit and scope of the invention. Such changes may include but are not limited to: (1) altering the direction considered to be the normal forward direction of flight, (2) adding springs, dampers, brakes, locking devices, movement limiting devices or other mechanisms to the swivelling part of the fuselage, (3) making the parts of the fuselage which swivel with respect to each other different in size, for example so that one is much bigger and the other is much smaller, (4) adding other substantial fuselage components to the design which move with one or other of the fuselage sections discussed or that are suspended or supported in some other way by the invention described herein, (5) using rotors which are overlapping or not overlapping or displaced vertically with respect to each other, (6) using rotors which are synchronized or unsynchronized in rotational movement, (7) mounting one or other of the rotor shafts so that it is inclined at an angle to the other such that the rotor shafts cannot assume a parallel orientation regardless of the position of the swivelling mechanism, (8) using any number of motors, including one or zero, to provide motive power, (9) using any rotor blade length or any number of rotor blades or dissimilar main rotors

Claims

CLAIMS I claim: 1 A tandem rotorcraft including: (a) two main rotors attached to, (b) two main rotor shafts, attached to, (c) a fuselage including (d) means to control the pitch of the rotor blades of said two main rotors as they rotate and (e) means to allow said main rotor shafts to swivel with respect to each other and (f) two electric motors used to power each of said main rotors separately where said swivelling is driven by controlling the pitch of said blades of said main rotors in order to achieve control of said rotorcraft in yaw and said main rotor shafts are synchronized by a control system to accurately control electrical drive signals sent to said motors in concert and to account for relevant feedback signals so that synchronization of said main rotor shafts is achieved without a mechanical connection.
2 The tandem rotorcraft of claim 1 wherein each of said main rotors is constructed so as not to allow significant movement of the disk swept out in flight by the blades of the rotor in question with respect to the shaft to which said rotor in question is attached.
3 The tandem rotorcraft of claims 1 or 2 wherein said control system makes use of feedback information from one or more of the following sources: (a) voltage measurements made on connections to the windings of said motors and (b) measurements made of the position of said rotor shafts using position sensors.
4. The tandem rotorcraft of any of the above claims including flight control means to stabilize and control said tandem rotorcraft, said flight control means using information derived from sources including but not limited to any one, two or three of the following: (a) a first three-axis gyroscope unit or equivalent system mounted on any part of the fuselage from which the rotation of the first of said main rotor shafts can be measured (b) a second three-axis gyroscope unit or equivalent system mounted on any part of the fuselage from which the rotation of the second of said main rotor shafts can be measured, (c) means to measure the rotational position of the section of the fuselage connected to the first of said main rotor shafts with respect to the section of fuselage connected to the second of said main rotor shafts.
5. The tandem rotorcraft of any of the above claims wherein the disks swept out by the movement of rotor blades making up said two main rotors overlap and said rotor blades are prevented from colliding by virtue of the said main rotors rotating in opposite directions in a synchronized manner so that blades from said two main rotors intermesh.
6. The tandem rotorcraft of any of the above claims combined with any number of further rotors for the purposes of propulsion or stabilisation.
7 A method of flying a rotorcraft consisting of (a) providing two main rotors attached to, (b) two main rotor shafts, attached to, (c) a fuselage including (d) means to control the pitch of the rotor blades of said two main rotors as they rotate and (e) means to allow said main rotor shafts to swivel with respect to each other and / 4- \ 4-T-r TZ"V ZX I ZX Z&4-4-4 Z4 4-V-4 Z-V-4- Z-V4-Z1 1 1 ΖΊ Z-l z~l 4-Z-v 4-^ Z-VTT 7 z~l»* Z-l K Z"vT n z4 4-V-4 z\ 4 4-% 4*Z"%4- Ζ"%4*·ΖΊ ΖΊ z^va Z% 4-Z% 4- Z-l I XT and driving said swivelling by controlling the pitch of said blades of said main rotors in order to achieve control of said rotorcraft in yaw and synchronising said main rotor shafts using a control system to accurately control electrical drive signals sent to said motors in concert and to account for relevant feedback signals so that synchronization of said main rotor shafts is achieved without a mechanical connection.
8 The method of claim 7 wherein each of said main rotors is constructed so as not to allow significant movement of the disk swept out in flight by the blades of the rotor in question with respect to the shaft to which said rotor in question is attached.
9 The method of claims 7 or 8 wherein said control system makes use of feedback information from one or more of the following sources: (a) voltage measurements made on connections to the windings of said motors and (b) measurements made of the position of said rotor shafts using position sensors.
10. The method of claims 7, 8 or 9 including flight control methods to stabilize and control said tandem rotorcraft, said flight control methods using information derived from sources including but not limited to any one, two or three of the following: (a) a first three-axis gyroscope unit or equivalent system mounted on any part of the fuselage from which the rotation of the first of said main rotor shafts can be measured (b) a second three-axis gyroscope unit or equivalent system mounted on any part of the fuselage from which the rotation of the second of said main rotor shafts can be measured, (c) means to measure the rotational position of the section of the fuselage connected to the first of said main rotor shafts with respect to the section of fuselage connected to the second of said main rotor shafts.
11. The method of claims 7, 8, 9 or 10 wherein the disks swept out by the movement of rotor blades making up said two main rotors overlap and said rotor blades are prevented from colliding by virtue of the said main rotors rotating in opposite directions in a synchronized manner so that blades from said two main rotors intermesh.
12. The method of claims 7, 8, 9, 10 or 11 where any number of further rotors are provided for the purposes of propulsion or stabilisation.