AU2019219790A1 - Device and method for improving the pitch control of a fixed-wing aircraft in stall/post-stall regime - Google Patents

Abstract

DEVICE AND METHOD FOR IMPROVING THE PITCH CONTROL OF A FIXED-WING AIRCRAFT IN STALL/POST-STALL REGIME 5 The present invention relates to a device and a method for improving the pitch control of a fixed-wing aircraft, particularly in a deep wing stall. The device in accordance with the present invention comprises at least a rotary-wing unit (14) for generating moment about the pitch axis of said aircraft to effectively actuate the pitch control of said aircraft, even in a deep stall. Computer simulation results suggested that an 10 airplane in the form of an engineless sailplane, equipped with the present invention will be able to perform multi-flip waterfall maneuver with ease. Simulations further revealed that the present invention would enable an airplane with thrust-to-weight ratio of greater than unity to have the capability to raise its angle of attack to a high angle while stationary on the ground in order to perform ultra-short takeoff based on 15 harrier maneuver. Most Illustrative Diagram: FIG. 5 Page 5/7 V =25 km-h-1 V =0 km-h-1 V =0 km-h-1 (a) (b) (c)

Description

DEVICE AND METHOD FOR IMPROVING THE PITCH CONTROL OF A FIXED-WING AIRCRAFT IN STALL/POST-STALL REGIME

FIELD OF INVENTION

The present invention relates to a device and a method for improving the pitch control of a fixed-wing aircraft, particularly in stall/post-stall regime.

BACKGROUND OF INVENTION

Stabilizing fins known as stabilizers are key components in maneuvering an 10 aircraft, and much of the effort to develop stabilizers and their control surfaces was done during the 1800s when aviation pioneers such as Sir George Cayley of Britain and Alphonse Penaud of France began experimenting with models as well as manned gliders [1]. A typical fixed-wing aircraft (or airplane) today still very much inherits the basic design concepts Cayley developed during the first half of the 19th century.

While the fin-based stabilizers perform as intended even in supersonic domain, the shortcoming becomes apparent in the stall/post-stall regime and extreme aerobatic flights which often operate in this regime. The shortcoming lies in the fact that a horizontal stabilizer and its elevator require sufficient air-flow in order to effectively actuate pitch control of an airplane. For that reason, most aerobatic fixed-wing aircraft 20 are designed such that the control surfaces are immersed in propeller wash so that they are able to perform many of the signature post-stall aerobatic maneuvers such as “prop-hang” (vertical hover), “flatspin”, “blender”, “harrier”, “tailslide”, “waterfall” and their derivatives [2-5], The aerobatic maneuvers most relevant to this invention are harrier and waterfall. The harrier maneuver is one in which the aircraft flies in its 25 post-stall regime in trim at high angles of attack near 45° with nose-up elevator input [5], This maneuver relies on lift from the wing and the vertical component of propeller thrust in order to sustain level trimmed flight [5], A distinctive characteristic of these aircraft is the thrust-to-weight ratio that exceeds unity [2] and therefore can potentially be used to create VTOL airplanes with ultra-agility.

However, the requirement of having sufficient airflow over a stabilizer imposes a fundamental limit on the scope and quality of post-stall maneuvers an

2019219790 21 Aug 2019 aerobatic airplane can perform. Take for example, the “waterfall” maneuver, where an airplane pivots 360° over its pitch axis with very little forward motion, and it can involve one or multiple flips [6,7], Aerobatic airplanes with conventional horizontal stabilizer will have difficulty executing the “waterfall” maneuver despite its 5 seemingly simple description, because in order to exert pitching moment around the pitch axis, strong propeller wash, and hence forward propulsion must be presence which inevitably pulls the aircraft forwards. It would be highly desirable to overcome such shortcoming, so that propeller thrust and pitch control of the airplane can be operated independently.

The present invention thus proposes the concept of “roto-stabilizer” as a possible replacement for the role of traditional stabilizers as an effective solution to overcome the limitations of conventional stabilizers in the post-stall regime.

SUMMARY OF INVENTION

An object of the present invention is to improve the pitch control of a fixedwing aircraft in the stall/post-stall regime.

The device in accordance with the present invention comprises at least a rotary-wing unit for generating pitching moment about the pitch axis (lateral axis) of said aircraft, even in post-stall state. The device plays the role of a horizontal stabilizer 20 in controlling pitch of said aircraft. The device's moment arm should preferably be substantially equal to or longer than the moment arm of a horizontal stabilizer for which it is substituting. The primary advantage of the present invention over conventional stabilizer-elevator approach is its effective and authoritative pitch control ability during a deep stall, including when the airspeed of the aircraft is zero 25 with the airplane’s landing gears on the ground. Computer simulation results suggested that an airplane, in the form of an engineless sailplane, equipped with the present invention will be able to perform multi-flip waterfall maneuver with ease. Simulations further revealed that the present invention would enable an airplane to have the capability to raise its angle of attack to a high angle while stationary on the 30 ground in order to perform ultra-short takeoff based on the harrier maneuver. When the nose angle is raised further to 90°, vertical takeoff and landing (VTOL) is

2019219790 21 Aug 2019 possible. Additionally, an aircraft equipped with the present invention will retain control margins through a wide range of center of gravity (CG) positions. An obvious benefit of this is that large fuel loads do not need to be regularly distributed during flight to maintain CG limits, leading to design simplification, and flight safety 5 improvement. We termed the present invention as “roto-stabilizer” and we believed the concept of roto-stabilizer will play an important role in the future of aviation.

Reference will now be made to the drawings wherein like numerals refer to like elements throughout. The proceeding disclosure is provided by way of example and not by way of limitation.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of an airplane in the form of an engineless sailplane having a horizontal stabilizer and a vertical stabilizer in tail-aft configuration.

FIG. 2(a) is a perspective view showing the exemplary sailplane as in FIG. 1 except 15 its horizontal stabilizer is completely absent and is being substituted by the device in accordance with the present invention.

FIG. 2(b) is a close-up view of an embodiment of the device in the present invention.

FIG. 3 is a perspective view of an engineless sailplane employing the device in accordance with the present invention wherein the rotary-wing units are present in 20 opposing pairs because fixed-pitch rotary wings are used. Inset shows a control scheme termed as asymmetric-V proposed in the present invention.

FIG. 4 shows an embodiment of the device in the present invention being employed in a canard configuration on a fixed-wing aircraft.

FIG. 5 shows a series of side views of a typical ultra-short takeoff sequence. Initially, 25 the forward airspeed, V = 0 km-h¹.

FIG. 6 shows a side view of the canard airplane as in FIG. 4, and the possible locations of the device in the present invention can be represented by a two-dimensional diagram comprising a virtual circle wherein the center of the circle coincides with the CG of the aircraft of interest and the radius of the circle is substantially the device’s 30 moment arm.

2019219790 21 Aug 2019

FIG. 7 is a perspective view of an exemplary aircraft of “above CG” configuration incorporating the device in accordance with the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

The concept of roto-stabilizer proposed in the present invention generally involves the use of one or more rotary-wing units as primary substitution for the stabilizer of interest. The concept is broadly applicable to both horizontal and vertical stabilizers, but the primary focus in the present invention is on horizontal stabilizer given its distinct advantages.

The present invention (roto-stabilizer) relates to a device for improving the pitch control of an airplane, particularly in the stall/post-stall regime, which is useful for pitch-related aerobatic applications or ultra-short takeoff and landing (ultraSTOL). In embodiments, the device in accordance with the present invention comprises at least a rotary-wing unit for generating moment about the pitch axis of the airplane to effectively actuate the pitch control, even in a deep wing stall. The device's moment arm (17) should preferably be substantially equal to or longer than the moment arm (13) of the horizontal stabilizer (10) for which it is substituting. Three different general exemplary embodiments of the present invention are presented herewith, two of which are widely known in aviation, namely tail-aft and canard configurations. The other configuration is one in which the rotary-wing unit(s) is/are disposed substantially above the CG of said aircraft. Such configuration does not work properly with conventional surface-based elevator control, and is therefore very much unique to the present invention leading to a new airframe design, which may be an attractive candidate for the emerging field of personal aviation and urban 25 air mobility.

FIG. 1 shows the perspective view of an engineless sailplane having a horizontal stabilizer (10) and a vertical stabilizer (12) in tail-aft configuration. The designations of the elements belonging to the horizontal stabilizer (10) are preceded by the numeral ‘10’ and the same numbering methodology is used for other 30 assemblies having sub-elements. Elevator control surface (102) is attached to the horizontal stabilizer (10) for pitch control of the sailplane. The moment arm of the

2019219790 21 Aug 2019 horizontal stabilizer (10) is indicated by a dash line (13). It is noteworthy that the engineless sailplane relies on rising air or thermal to remain airborne.

FIG. 2(a) shows the same exemplary sailplane wherein its horizontal stabilizer (10) is completely absent and is being substituted by the device in accordance with 5 the present invention involving at least a rotary-wing unit (14). In other words, the rotary-wing unit(s) (14) is/are located behind the CG for said aircraft of tail-aft configuration. Each rotary-wing unit (14) further comprises a plurality of rotary wings (142) and a driving means (144) which drives said rotary wings (142) into rotation to generate thrust, as shown in FIG. 2(b). The driving means (144) in this 10 example is a brushless electric motor; another suitable candidate is internal combustion engines. Furthermore, the rotary wings (142) can be directly mounted on the driving means (144), or indirectly, for example, via transmission shaft or timing belt as in a typical tail rotor system of a model helicopter. The rotary wings (142) can be either variable pitch or fixed-pitch and each setup has its own pros and cons. Said 15 rotary wings (142) in this example are of variable pitch and practical angular travel range is usually ±10°. The rotary-wing unit (14) of variable pitch produces zero aerodynamic thrust when the pitch of the rotary-wings (142) is 0°. A common means to actuate the variable pitch of the rotary-wings (142) is digital servo especially if the said aircraft is relatively small with a wing span of about 2 meters. The rotating rotary 20 wings (142) produce an aerodynamic thrust either in upward or downward direction as indicated by arrows (16A), (16B), respectively. The moment arm of the device in the present invention is indicated by a dotted line (17). As a rule of thumb, the device's moment arm (17) should preferably be substantially equal or longer than the moment arm (13) of the horizontal stabilizer (10) for which it is substituting. The aerodynamic 25 thrust generated by the rotary-wing unit (14) in combination with its moment arm (17) produces a pitching moment that pitches the nose of the sailplane up or down depending on the angular pitch value of the rotary-wings (142). Since the generation of thrust does not require the aircraft to have forward airspeed, it is therefore possible for the aircraft to perform rapid multi-flip at an airspeed lower than the stall speed, 30 hence giving rise to the multi-flip waterfall maneuver - a feat that few, if not any, conventional motorless sailplanes, whether full-scaled or otherwise, could have accomplished. Additionally, powerful pitch control of the aircraft with no requirement for forward airspeed leads to important ultra-STOL or VTOL

2019219790 21 Aug 2019 applications relevant to urban air mobility sector but details of which will be disclosed and discussed in greater depth using the embodiment of canard configuration.

In conventional surface-based elevator (102) operation, its angular displacement can be actuated directly via pilot’s manual input, or alternatively, the 5 pilot’s control inputs can be fed into a gyro prior to being transmitted to the elevator (102) for auto-stabilization. Likewise, the device in the present invention can be controlled by either method. While simulation results suggested that it is possible to attain sufficient longitudinal stability by solely relying on the pilot’s direct input, it is preferably to include a gyro for enhanced stabilization and to reduce workload for the 10 pilot.

Level flight simulations revealed that the flying characteristics of a model sailplane equipped with the device in the present invention and a pitch gyro was hardly distinguishable from those of the original sailplane with conventional stabilizer. The simulated sailplane has a wingspan of 3.16 m. The rotary-wing unit 15 (14) was directly driven by a brushless motor at a nearly constant rotational speed of

6800 rpm and generating a maximum static thrust of 3.8 N on either direction (16A), (16B). The diameter of the rotary wings (142) was 254 mm. The final wing loading of the sailplane equipped with the device in accordance with the present invention is 32.94 g-dm’². The environmental wind was set to about 20 kmh¹ with orographic 20 updraft of approximately 1.7 ms¹.

In accordance with the present invention, the rotary wings (142) can be fixedpitch as well. The pitch control can work with one rotary-wing unit (14) using a combination of aerodynamic thrust and natural pitching due to gravity, however, it is preferably that the rotary-wing units (14) are present in opposing pairs, as shown in 25 FIG. 3. This is because each unit (14) can only exert a thrust in one direction. Additionally, in the present invention, a control scheme is proposed to eradicate a well-known shortcoming relating to jitter issue of brushless electric motors and it is termed as “asymmetric-V”. Said control scheme is applied to the rotary-wing units (14A), (14B) so that power is supplied to the motors (144) driving the fixed-pitch 30 rotary wings (142) to ensure that they are rotating at a substantially minimal speed whenever in use.

2019219790 21 Aug 2019

Inset in FIG. 3 shows an example of a general plot of the outputs to the motors driving the rotary wings (142) as a function of pitch input. For example, if said motors were pre-determined to be running smoothly without jitter at an output of 5%, then a 5% output at zero pitch input will ensure that the motors (144) are rotating at a 5 minimal jitter-free rotational speed. To pitch up the aircraft, power to the rotary-wing unit (14A) is increased while the power to the other rotary-wing unit (14B) remains constant at 5%, and the reverse is true for pitching down the airplane.

FIG. 4 shows an exemplary embodiment of the device in accordance with the present invention being employed in a canard configuration on an aircraft. The device 10 is generally disposed towards the front end of the airframe. In this example, the device comprises one rotary-wing unit (14) mounted on a nose cone (18) which is connected to fuselage (19). The nose cone (18) can optionally be made to be rotatable about an axis ‘A’ parallel with the pitch axis of the airplane as shown in the inset of FIG. 4. In other words, the rotary-wing unit (14) is located in front of the CG for said aircraft of 15 canard configuration. The airplane in this example has a wing span of 2000 mm and a fuselage length of about 1145 mm. The wing loading was 73.27 g-dm’² which resulted in a nominal stall speed of approximately 43.6 km-h¹. The airplane has two primary propulsors (20), each with a propeller (22) attached, and together they generate a total static thrust that exceeds the weight of the airplane, as typically found 20 in 3D aerobatic airplanes. The primary propulsors (20) are attached to the wings (23). Differential thrust from the main propellers (22) was used to actuate yaw, especially in the post-stall regime. In addition to having the device in the present invention, the aircraft was equipped with a conventional canard (10) fitted with trailing-edge flaps (102) which act as elevators. The device in the present invention enables the nose of 25 the aircraft to be raised to establish a high angle of attack, thereby making use of the harrier maneuver to realize the ultra-STOL or even VTOL if the angle of attack were to be further increased to 90°. The ailerons (232) immersed in strong propeller wash will ensure authoritative roll control during harrier maneuver. The canard aircraft is of tri-gear configuration with nose gear (26) and main gears (28). The aircraft is also 30 equipped with rudder (122).

FIG. 5 is a series of side views of a typical ultra-short takeoff sequence. Initially, the aircraft is stationary with its nose gear (26) and main gears (28) on the ground [FIG. 5(a)], At the start of the takeoff sequence, the device in the present

2019219790 21 Aug 2019 invention raises the angle of attack of the aircraft to a high angle of approximately 45° [FIG. 5(b)], With the nose raised, the aircraft initiated the harrier maneuver involving thrusts from the primary propulsors (20). From simulations, the required ground run distance was only about 3 m, or approximately 3 body lengths.

Furthermore, the takeoff airspeed was only about 25 km-h¹ which was significantly lower than its stall speed (43.6 km-h¹) and this is a hallmark of the harrier maneuver [FIG. 5(c)], The horizontal acceleration in this case was about 3 m-s’² or 0.31G. In full scaled aircraft application, such acceleration is expected to be comfortable for the pilot or passengers and hence a viable takeoff method. Once the aircraft has gathered 10 sufficient airspeed, it may revert to conventional canard elevator (102) for pitch control and the present invention may be rotated via the nose cone (18) so that its thrust vector is directed to the front to augment the propulsion. The short landing attitude is essentially a descending harrier pass, i.e., very much a reverse process to that as shown in FIG. 5. Once the aircraft’s main landing gears (28) are on the ground, 15 the nose gear (26) is gradually lowered to the ground by the device in the present invention. The lowering of the nose angle can take place either in tandem with the deceleration of the air-craft, or after it has come to a halt. The distance from the point the main gears touched the ground to a complete stop for the simulated model is about 5.8 m. If desired, shorter ground roll distance could be achieved with higher angle of 20 attack during harrier descent prior to touch-down. At angle of attack of 90°, it is essentially a vertical landing.

The device in the present invention could also serve as a safety redundancy in the event the elevator surfaces (242) are malfunctioned. On the other hand, if the present invention were to suffer damage mid-flight, it is possible for the aircraft to 25 make a conventional runway landing using flight control surfaces, namely ailerons (232), elevators (102) and rudder (122).

A generalization can be drawn based on these embodiments disclosed thus far in regard to the configurational placement of the device in the present invention with respect to the CG of an aircraft while airborne. Referring now to FIG. 6 which shows 30 a sideview of the canard airplane, the possible locations of the device can be represented by a two-dimensional diagram comprising a virtual circle (30) wherein the center of the circle coincides with the CG of the aircraft of interest and the radius of the circle is substantially the device’s moment arm (17). The device can thus be

2019219790 21 Aug 2019 placed anywhere along the circumference (30) of the circle, and the thrust vectors (16A), (16B) are preferably substantially tangential to the circumference (30).

When the aircraft is stationary on the ground however, the pivot point would be shifted to coincide with the axle of the main gears (28), from which it can be 5 deduced that for efficient ultra-short takeoff, the thrust line (32) of the main propulsors (20) should preferably be no higher than the horizontal line (34) passing through the center of the axle (282) of the main gears (28), in order to minimize the effect of opposition of torques from the main propulsors (20) which results in the nose of the aircraft being pushed downwards. Note that if the aircraft is configured to 10 takeoff from the ground, the practical placement of the device will be reduced to an almost semi-circle as the remaining portion is hidden by the ground and therefore unavailable. Airplanes which are launched by hand or catapult will retain the full virtual circle. By analyzing the virtual circle diagram as depicted in FIG. 6, another possible embodiment is one wherein the device comprising at least one rotary-wing 15 unit (14) is located substantially above the CG as shown in FIG. 6 and it is termed as “above CG” configuration in the present invention.

FIG. 7 is a perspective view of an exemplary aircraft of “above CG” configuration incorporating the device in accordance with the present invention. This exemplary aircraft has thrust-to-weight ratio of greater than one, and it comprises at 20 least one main propulsor (20). While seemingly differing in physical appearance to the canard aircraft, they both share similar ultra-STOL or VTOL capability. Furthermore, from insights gained from the analysis of the virtual circle diagram in FIG. 6, the thrust line (32) of the main propulsors (20) are deliberately lowered. A distinct advantage of this “above CG” embodiment is that the thrust vector(s) of the 25 rotary-wing unit(s) (14) is/are substantially horizontal during cruising flight, and can therefore readily augment the propulsion produced by the main propulsors (20) without a need for additional complex rotational mechanism and this leads to enhanced reliability and weight-saving.

The foregoing description of the present invention has been presented for 30 purpose of illustration and description. Furthermore, the description is not intended to limit the invention to the form disclosed herein. Consequently, variations and modifications commensurate with the above teachings, and skill and knowledge of

2019219790 21 Aug 2019 the relevant art, are within the scope of the present invention. The embodiments described hereinabove are further intended to explain best modes known of practicing the invention and to enable other skilled in the art to utilize the invention in such or other embodiments and with various modifications required by the particular application(s) or use(s) of the present invention. It is intended that the appended claims be constmed to include alternative embodiments to the extent permitted by the prior art.

References:

1. Yoon, J.N. (2002) Origins of Control Surfaces.

http://www.aerospaceweb.org/question/history/q0103.shtml

2. Selig, M.S. (2010) Modeling Full-envelope Aerodynamics of Small UAVs in Real Time. Proceedings of the AIAA Atmospheric Flight Mechanics Conference, 7635.

3. International Miniature Aerobatic Club (2009) About IMAC.

http://www.mini-iac.com/

4. The World Air Sports Federation (2018) F3M Large Radio Control Aerobatics World Cup. http://www.fai.org/world-cups/f3m

5. Selig, M.S. (2014) Real-Time Flight Simulation of Highly Maneuverable

Unmanned Aerial Vehicles. Journal of Aircraft, 51,1705-1725. DOI:

10.2514/1.C032370

6. Wikipedia (2018) 3D Aerobatics.

https://en.wikipedia.org/wiki/3D_Aerobatics

7. Giant Scale News (2013) The keys to 3D.

https://www.giantscalenews.com/threads/the-3d-compiliation-facts-how-to- reference-more.6332/

Claims (8)

1. CLAIMS:

   1. A device for improving the pitch control of a fixed-wing aircraft in stall/post-stall regime comprising:

   at least a rotary-wing unit (14) for generating moment about the pitch axis of 5 said aircraft, each of the rotary-wing unit (14) further comprising a plurality of rotary wings (142), the device's moment arm (17) should preferably be substantially equal to or longer than the moment arm (13) of a horizontal stabilizer (10) for which it is substituting.
2. 2. The device as claimed in Claim 1, wherein said horizontal stabilizer(s) is/are 10 physically absent.
3. 3. The device as claimed in Claim 1, wherein said rotary wings (142) are variable pitch.
4. 4. The device as claimed in Claim 1, wherein said rotary wings (142) are fixed-pitch, and the rotary-wing units (14) are preferably present in opposing pairs.

   15 5. The device as claimed in Claim 4, wherein a proposed control scheme termed as asymmetric-V is applied to the rotary-wing units (14) so that power is supplied to the electric motors (144) driving the fixed-pitch rotary wings (142) to ensure that they are rotating smoothly without jitter at a substantially minimal pre-determined speed whenever in use.

   20 6. The device as claimed in Claim 1, wherein the rotary-wing unit(s) (14) is/are located behind the CG for said aircraft of tail-aft configuration, the axis/axes of rotation of said rotary-wings (142) is/are substantially parallel to the yaw axis of said aircraft.

   7. The device as claimed in Claim 1, wherein said rotary-wing unit(s) (14) is/are 25 located in front of the CG for said aircraft of canard configuration, and the axis/axes of rotation of said rotary wings (142) is/are substantially parallel to the yaw axis of said aircraft.

   8. The device as claimed in Claim 1, wherein said rotary-wing unit(s) (14) is/are located above the CG for said aircraft of “above CG” configuration, and the axis/axes

   2019219790 21 Aug 2019 of rotation of said rotary wings (142) is/are substantially parallel to the roll axis of said aircraft.

   9. A fixed-wing aircraft, comprising the device according to Claim 6.

   10. A fixed-wing aircraft, comprising the device according to Claim 7.
5. 5 11. A fixed-wing aircraft, comprising the device according to Claim 8.

   12. A method for improving the pitch control of a fixed-wing aircraft in stall/poststall regime comprising at least a rotary-wing unit (14) for generating moment about the pitch axis of said aircraft, each of the rotary-wing unit (14) further comprising a plurality of rotary wings (142); aerodynamic thrust generated by the rotary-wing unit
6. 10 (14) in combination with its moment arm (17) produces a pitching moment that pitches the nose of said aircraft up or down in order to control the pitch of said aircraft during flight and while in a deep stall, including when said aircraft is stationary on the ground with forward airspeed of zero to achieve an attitude of high angle of attack suitable for ultra-short takeoff or vertical takeoff; the ultra-short takeoff sequence 15 comprising:

   (a) said aircraft is stationary with its nose gear (26) and main gears (28) on the ground;

   (b) raising the angle of attack of the aircraft to a high angle; and (c) said aircraft initiating harrier maneuver involving thrust(s) from the

   20 primary propulsor(s) (20).
7. 13. A fixed-wing aircraft, implementing the method according to Claim 12.
8. 14. The device as claimed in Claim 1, wherein possible locations of said device can be represented by a two-dimensional diagram comprising a virtual circle (30) in which the center of the circle coincides with the CG of the aircraft of interest and the

   25 radius of the circle is substantially the device’s moment arm (17); said device can be placed anywhere along the circumference (30) of the circle, and the thrust vectors (16A), (16B) are preferably substantially tangential to the circumference (30).