Control unit, force actuation element and method for adjusting an element of an aircraft

Abstract

Relates to a control unit (6, 7) for controlling the pitch of an aircraft (1), the control unit (6, 7) comprising at least one adjustable control element (6) and an adjustment element (7) for actively adjusting the at least one control element (6) by means of a continuous deformation of the at least one control element (6). The invention also relates to an aircraft (1) having such a control unit (6, 7) and to a method for controlling the pitch of an aircraft (1). There is also provided a force actuating element (8) that can be filled with a power transmitting fluid (10) to generate an actuating force. An aircraft (1) having such a force actuating element (8) and a method for adjusting an aerodynamic or hydrodynamic element (3, 51, 32, 34, 4, 5, 6) are also provided.

Description

Technical Field

[0001] The invention relates to a control unit for controlling the pitch of an aircraft, in particular an airplane, during flight, and to an aircraft comprising such a control unit and a method for controlling the pitch of an aircraft. The control unit comprises an adjustable control element, which can be adapted, for example, to realize a combined function of a horizontal stabilizer and an elevator. The invention also relates to a force-actuated element, which can be used in particular for adjusting an aerodynamic or hydrodynamic element of an aircraft or a watercraft. An aircraft having such a force-actuated element and a method for adjusting an aerodynamic or hydrodynamic element are also provided. Background Art

[0002] Angle of attack limitation and maneuverability enhancement

[0003] Much research has been done to prevent aviation accidents. Many electronic schemes have been developed to achieve this goal, both for flight path stabilization such as on Airbus aircraft and for speed stabilization such as on Boeing aircraft. Human factors, especially the human pilot's response to stress, have been studied intensively. After each accident, crew training is improved around the world to incorporate the lessons learned; flight simulator hardware and software are constantly improved to more realistically represent actual aircraft behavior, especially when approaching a stall.

[0004] The prior art horizontal tail arrangement has not been redesigned in decades. Furthermore, during this time, any measures to reduce the probability of pitch loss control have focused solely on control loops (e.g., stall warnings, stick shakers / pusher, electronic maneuvering augmentation / limiting) and flight training improvements. For example, in the Boeing 737MAX, the Maneuvering Characteristics Augmentation System (MCAS) system is used to correct for the reduction in nose-down tendency approaching the stall angle by introducing an additional nose-down moment, which changes the control input required by the pilot to maintain the pitch attitude. The system thereby changes the maneuvering characteristics of the aircraft.

[0005] Such a design based solely on the control loop could lead to fatal consequences in the event of a failure if the design is not robust enough to at least achieve fault protection.

[0006] The McDonnell Douglas / Boeing C-17 Globemaster transport aircraft entered service in 1991 and used this system in a sophisticated fail-safe version; it was equipped with six angle-of-attack vanes with comparison logic and was specifically designed as an angle-of-attack limiter to avoid the deep stall-lock conditions to which the aircraft was prone due to its high wing and T-tail configuration. Iloputaife described the system in detail in [1], calculating that the overall probability of an incorrect angle-of-attack limit was 10 -9 The vane sensor is considered to have higher accuracy than other measurement methods at high angles of attack. The inherent single point failure problem of the push rod and the related reliability problems are pointed out.

[0007] For small aircraft, Chevalier and Faulkner published a system in 1980 that included an anti-stall device for small general aviation aircraft and conducted full-scale wind tunnel tests [2]. When the aircraft reaches its stall angle of attack, a small spoiler is deployed on the lower surface of the horizontal tail, just forward of the elevator hinge, which significantly reduces the downforce acting on the horizontal tail, in effect acting as a lift plate. This in turn prevents the aircraft from pitching up further into a stall. The system works as intended; however, it still relies on angle of attack sensors and actuators for the spoilers, both of which are points of failure that could result in dangerous consequences (e.g., incorrect deployment of spoilers at low angles of attack, causing the aircraft to dive).

[0008] Wing shape morphing technology

[0009] In relation to aircraft, the term “morphing” is often used to describe an aircraft element that can significantly change its external shape, for example to adapt to changing mission conditions during flight. As early as 1920, Parker proposed an aircraft wing with variable camber [3].

[0010] Barbarino et al. tabulated a number of different techniques that had been published at the time for achieving wing shape deformation [4]. This served as a comprehensive guide to the state of the art, such as eccentuator-based techniques, flexspars, spanwise curved HECS wings, active flutter control, and actuator requirements, to name just a few. The need for actuator locking was noted, otherwise the actuator would have to carry the full load. Gern, Inman, and Kapania calculated the actuation power requirements for a smart wing with a morphing wing surface; their conclusions indicated that distributed actuation had great potential for system robustness in the event of any failure or damage [5]. Dimino et al. also studied the advantages of distributed actuation over axis mechanisms from a reliability and fail-safe perspective [6].

[0011] Jha and Kudva published their own classification method, in which they distinguished between robotic and organic designs and investigated the influence of wing geometric parameters on aircraft performance [7]. According to their definition, in robotic designs, the wing area remains constant and the entire wing or part of it rotates, while in organic designs, the wing parameters vary to a greater extent. The authors point out that "robotic designs produce variable sweep angles, variable pitch angles, and variable dihedral wings, while organic designs produce leading edge control surfaces, trailing edge control surfaces, folding wing tips, and spoiler mechanisms. In organic designs, internal mechanisms cause the overall dimensions of the wing to vary."

[0012] Therefore, wing shape deformation is a very broad research field.

[0013] External deformations usually require high actuation loads, which conflict with the stiffness requirements. Several solutions to this problem have been proposed, including steel bar actuated honeycomb trusses [8], [9],

[10] . Vos et al. introduced a mechanism for warping control of morphing wings

[11] . With an open trailing edge, the wing acts like a C-section with an opening. Screws displace the upper and lower leading edges in opposite directions, causing torsion. This approach conveniently places the open part of the wing at the trailing edge and effectively solves the problem of actuator locking, as the screw itself acts as a locking device. However, only linear torsion along the span can be achieved.

[0014] A variation of the above scheme was applied to helicopter rotors by Mistry and Gandhi

[12] . The opening and the actuation mechanism are located on the underside rather than on the trailing edge. In order to overcome the compromise between torsional stiffness and actuation force, the fixed D-shaped spar head is retained. The result is a controlled camber deformation along the span.

[0015] Smart spars also exploit the mechanical properties of the torsion box. Amprikidis and Cooper developed different methods to modify the shear center of the wing. By rotating the spar or moving the spar within the wing box, the shear center can be moved, resulting in a change in the coupling of bending and twisting behavior. The acting aerodynamic forces provide the torque required to twist the wing

[13] . This is a case of passive deformation; this deformation is particularly suitable for aeroelastic problems; swept wings have a robust bending-twist coupling and are therefore ideal for aeroelastic tailoring

[14] .

[0016] A similar approach is the variable section spar, in which the shear web is moved as needed to control the position of the shear centre

[15] . The authors explicitly point out that this is in fact an open section beam and is therefore more susceptible to flutter and divergence due to reduced torsional stiffness; on the other hand, the authors mention that "the ability to control the position of the shear centre can be used to suppress flutter and divergence, thereby extending the flight envelope of the vehicle". On a more general level, Cesnik and Brown studied active warping control of a morphing wing configuration

[16] . In addition to the mechanical description of the morphing wing, they also point out various problems with soft flexible wings, such as flutter.

[0017] Another concept of the same type is the variable stiffness spar, where a rotating spar is incorporated into the wing box

[17] ,

[18] . This spar was used on the FA-18 model and full-scale wind tunnel testing was performed, demonstrating increased maneuverability.

[0018] Instead of using an open beam, Raither et al. used electrothermal and electromechanical effects to change the properties of the shear web. This semi-passive approach again utilized aeroelastic coupling instead of actuators, achieving low weight and high efficiency without actuators

[19] .

[0019] In terms of active deformation, actuation force is a key issue to be addressed. Actuation based on eccentric mechanisms can generate the required force and has been applied in several cases, such as the DARPA Smart Wing project

[20] ,

[21] . However, the use of eccentric mechanisms requires continuous application of actuation force because no actuator locking is provided.

[0020] Piezoelectric elements generate strong forces and have been successfully applied in a variety of applications that require strong forces and high stiffness, such as flexible wing spars

[22] .

[0021] Molinari incorporated piezoelectric actuators into an aerodynamic shell combined with a fixed D-shaped wing spar, flexible ribs, and corrugated structures

[23] . Roll control was achieved through camber deformation, thus realizing a fully maneuverable model aircraft, demonstrating the applicability of such schemes in practical flight control problems. In addition to piezoelectric actuators, SMA is another strong candidate for new materials for aircraft deformation

[24] .

[0022] Previtali et al. designed a similar morphing wing based on flexible ribs. His results showed that the wing could achieve a high level of roll control capability with lower or equivalent weight compared to conventional wings

[25] .

[0023] The high roll control capability of this approach is supported by Pecora et al.

[26] . Regarding the effectiveness of "pure" torsional deformation in roll control (i.e. without a fixed D-spar), the authors estimated that only a slight change in the twist angle would be required to achieve the same effect as an aileron. The authors investigated linear, quadratic, and cubic wing torsions, with linear torsions being the most efficient. To achieve satisfactory roll control, wing deformations of 2.5° to 7.5° were required, which is much less than standard aileron deflections. Active torsional deformation can therefore be a powerful tool for minimizing deformation while achieving aerodynamic effects.

[0024] Flexible ribs do not necessarily need to consist only of non-buckling elements. Gandhi et al. demonstrated the applicability of buckling elements in flexible mechanisms to achieve selective deformation with distributed flexibility for camber deformation of an airfoil with a fixed D-shaped spar

[27] .

[0025] The relatively low strength of compliant structures is a problem when considering them for high-load applications. Bistable structures exploit controlled buckling and are able to withstand higher loads in more than one state without requiring very high actuation forces. Thus, elastic instabilities contribute to shape adaptability. Runkel investigated buckling-induced selective stiffness in chiral structures and successfully applied chiral structures to passively deformable aircraft wing ribs

[28] . Mattioni et al. described thermally induced active multistable structures and applied them to trusses adapted to sweep at noncritical actuation speeds

[29] . However, thermal actuation occurs slowly, making them unsuitable for primary flight control applications.

[0026] Pressure adaptive honeycomb cells have been applied to morphing trailing edge flaps

[30] . Other types of deformation aim to directly control the skin, thereby locally changing the camber. Popov et al. proposed a scheme for transition flow control by deformation of the upper skin

[31] . The FishBAC scheme proposed by Wood et al. consists of basic curved splines at the center of the airfoil, stringers and flexible skin

[32] .

[0027] Probably the single main challenge of any variant solution remains the skin. A certain flexibility of the skin is almost always required. The skin resists any angular deflection of the wing, i.e., with any wing twist, the skin also needs to twist

[33] . Many skin designs have been proposed, for example, skins with auxetic honeycomb cores, which still require some kind of flexible covering

[34] ,

[35] ,

[36] ,

[37] . In order to avoid polymer-based flexible skins, or to avoid the need to segment them, corrugated surfaces have been proposed and analyzed

[38] ,

[39] . Corrugated surfaces have been successfully used in a working solution by Molinari, for example

[23] .

[0028] Morphing techniques specific to large and heavy aircraft

[0029] Transport category aircraft present unique challenges for the application of morphing. Most importantly, large and heavy aircraft have significantly higher wing loadings than smaller aircraft. Additionally, there are multiple factors such as environmental and operational factors that may limit the use of certain technologies.

[0030] Vasista et al. summarized the highly multidisciplinary nature and challenges of implementing a morphing wing

[40] . Specific load requirements and deflection requirements and results from multiple research projects were listed. Issues such as fatigue and the difference between the target deformed shape and the actual deformed shape due to aerodynamic loads were highlighted. The average pressure acting on the wing of a large transport aircraft was estimated to be around 15 to 19 kPa, which is up to five times that of a small general aviation aircraft.

[0031] Passive morphing schemes such as those proposed by Runkel

[28] and Raither

[19] aim to change the response of the wing under load and are capable of withstanding high loads. The same applies to variable stiffness spar schemes, which have been used on the FA-18

[17] ,

[18] .

[0032] In their review of deformable skins, Thill et al.

[41] listed a number of requirements and possible problems in the operating environment. Among them, the chemical resistance of the skin is mentioned, and the fact that highly flexible skin solutions are frequently combined with the use of elastomers or rubbers. The glass transition temperature is often very different from the required operating temperature. This problem becomes obvious when considering that passenger aircraft are often de-iced with ethylene glycol mixtures and operate for several hours in ambient temperatures as low as -60°C.

[0033] Active morphing has so far achieved limited success on large and heavy aircraft. Shape morphing has been demonstrated on the trailing edge of NASA's Gulfstream research aircraft, which uses morphing flaps.

[42] A commercial application of active morphing on transport aircraft is Boeing's variable geometry V-shaped structure, where a morphing structure has been successfully integrated into the engine nacelle to achieve noise reduction.

[43]

[0034] Some small-scale attempts at morphing flight control for larger aircraft include flaps and rudders. Pecora et al. designed a morphing flap for a large civilian aircraft. The morphing flap was made entirely of aluminum and featured discrete elements and segmented skins, which was effectively equivalent to building several simple flaps in a row

[26] .

[0035] Amendola et al. designed an adaptive aileron for a regional aircraft; it followed a very similar segmentation approach

[44] . Their solution also used aluminum parts and involved finger-like ribs. The aileron included both active and passive segments.

[0036] On the leading edge side, Rudenko et al. developed a droop nose solution involving a multi-link actuation mechanism. The morphing skin is made of GFRP laminates that provide high bending stiffness along the span direction and is embedded in an elastomeric base that allows cambermorphing

[45] .

[0037] Overall, the technical maturity level of active morphing remains low

[31] . As shown in this section, morphing options for transport aircraft remain quite limited and lack radically new designs. Friswell summarizes many of these issues in a publication entitled “Morphing Aircraft: an Improbable Dream?” He notes that for morphing of large aircraft, the main research focus is on auxiliary structures such as control surfaces, which still have a large impact. Actuator requirements and locking remain issues, and the flexible skins of the existing technology are also problematic in that they are not currently suitable as primary load-bearing components. He also notes that “One drawback of morphing aircraft research is that morphing tends to be added to existing aircraft that operate in the same manner. This retrofit approach directly eliminates some of the main advantages of morphing technology and limits any potential for versatility”

[46] .

[0038] Ajaj et al. conducted a comprehensive review of various morphing mechanisms and schemes, including elastic skin and morphing winglet flight control; they also studied the potential of the modified ones. Comparison of flexibility with mechanisms showed that flexibility is not useful for large and heavy aircraft. Their full conclusions are summarized in a publication titled "Morphing aircraft: The need for a new design philosophy"

[47] .

[0039] Background Technology - References

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[0044] [5] F. Gern, D. Inman, and R. Kapania, “Computation of Actuation Power Requirements for Smart Wings with Morphing Airfoils,” AIAA Journal, Vol. 43, No. 12, pp. 2481 et seq., December 2005.

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[0049]

[10] D. Ramrakhyani, G. Lesieutre, M. Frecker and S. Bharti, “Aircraft Structural Morphing Using Tendon-Actuated Compliant Cellular Trusses,” Journal of Aircraft, Vol. 42, No. 6, pp. 1615 ff., November-December 2005.

[0050]

[11] R. Vos, Z. Gürdal, and M. Abdalla, “Mechanism for Warp-Controlled Twist of a Morphing Wing,” Journal of Aircraft, Vol. 47, No. 2, pp. 450 ff., March-April 2010.

[0051]

[12] M. Mistry and F. Gandhi, “Design, fabrication, and benchtop testing of ahelicopter rotor blade section with warp-induced spanwise camber variation,” Journal of Intelligent Material Systems and Structures, Vol. 26 (No. 10), pp. 1272-1289, 2015.

[0052]

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[0053]

[14] S. Ehlers and T. Weisshaar, “Effect of Adaptive Material Properties on Static Aeroelastic Control,” in 33rd Structures, Structural Dynamics and Materials Conference, Dallas, Texas, April 13–15, 1992.

[0054]

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[0055]

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[0056]

[17] P. Chen, D. Sarhaddi, R. Jha, D. Liu, K. Griffin, and R. Yurkovich, “Variable Stiffness Spar Approach for Aircraft Maneuver Enhancement using ASTROS,” Journal of Aircraft, Vol. 37, No. 5, pp. 865 ff., September-October 2000.

[0057]

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[19] W. Raither, E. Furger, M. Zuendel and P. Ermanni, "Variable-stiffness skinconcept for camber-morphing airfoils," Journal of Intelligent Material Systems and Structures, London, United Kingdom, 2015.

[0059]

[20] J. Kudva, "Overview of the DARPA Smart Wing Project," Journal of Intelligent Material Systems and Structures, Volume 15, Pages 261-267, April 2004.

[0060]

[21] R. Musgrove, “The Eccentuator-a new Concept in Actuation,” NASA Angelley Res. Center Proc. of the 14th Aerospace Mech. Symp, Hampton, Virginia, 1980, N80 23501, pp. 57 et seq.

[0061]

[22] R. Barrett, “Active aeroelastic tailoring of an adaptive Flexsparstabilator,” Smart Materials and Structures, IOP Publishing, pp. 723ff, 1996.

[0062]

[23] G. Molinari, A. F. Arrieta, M. Guillaume, and P. Ermanni, “Aerostructural Performance of Distributed Compliance Morphing Wings: Wind Tunnel and Flight Testing,” AIAA Journal, Vol. 54, No. 12, December 2016.

[0063]

[24] J. Harrison, S. Wise, R. Bryant, R. Cano, and T. Gates, “Innovative materials for aircraft morphing,” 5th Annual Symposium on Smart Structures and Materials (SPIE), San Diego, California, June 16, 1998, pp. 240ff.

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[25] F. Previtali, G. Molinari, AFArrieta, M. Guillaume and P. Ermanni, "Design and experimental characterization of a morphing wing with enhanced corrugatedskin," Journal of Intelligent Material Systems and Structures, London, United Kingdom, 2016.

[0065]

[26] R. Pecora, F. Amoroso, and L. Lecce, “Effectiveness of Twist Morphing in Roll Control,” Journal of Aircraft, Vol. 49, No. 6, pp. 1666ff., November-December 2012.

[0066]

[27] F. Gandhi and P. Anusonti-Inthra, "Skin design studies for variablecamber morphing airfoils," Smart Materials and Structures 17 (2008) 015025 (page 8).

[0067]

[28] F. Runkel, U. Fasel, G. Molinari, AFArrieta and P. Ermanni, "Wing twisting by elastic instability: A purely passive approach," Composite Structures 206 (2018) 750-761.

[0068]

[29] F. Mattioni, Weaver, PKPM and M. Friswell, “The application of thermally induced multistable composites to morphing aircraft structures,” inSPIE Smart Structures and Materials+Nondestructive Evaluation and HealthMonitoring, San Diego, California, April 1, 2008, pp. 693012-1 et seq.

[0069]

[30] R. Vos and R. Barrett, "Mechanics of pressure-adaptive honeycomb and its application to wing morphing," Smart Materials and Structures, pp. 1-11, 2011, 20 094010.

[0070]

[31] A. Popov, T. Grigorie, and R. Botez, “Modeling and Testing of a MorphingWing in Open-Loop Architecture,” Journal of Aircraft, Vol. 47, No. 3, pp. 917 ff., May-June 2010.

[0071]

[32] B. Woods, M. Friswell, and N. Wereley, “Advanced Kinematic Tailoring for Morphing Aircraft Actuation,” AIAA Journal, Vol. 52, No. 4, pp. 788 ff., April 2014.

[0072]

[33] M.Detrick and G.Washington, "Modeling and Design of a Morphing Wing for Micro Unmanned Aerial Vehicles via Active Twist," in 48th AIAA / ASME / ASCE / AHS / ASC Structures, Struct.Dynamics, and Materials Conference, Honolulu, Hawaii, April 23-26, 2007, AIAA 2007-1788.

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[34] A. Dale, J. Cooper, and A. Mosquera, “Adaptive Camber-Morphing Wing using 0ν Honeycomb,” in 54th AIAA / ASME / ASCE / AHS / ASC Structures, Struct. Dynamics, and Materials Conference, Boston, Massachusetts, April 8–11, 2013, 10.2514 / 6.2013-1510.

[0074]

[35] K. Olympio and F. Gandhi, “Flexible Skins for Morphing Aircraft Using Cellular Honeycomb Cores,” Journal of Intelligent Material Systems and Structures, Vol. 21, pp. 1719 ff., November 2010.

[0075]

[36] DQS Wenjun, “Airfoil Design and Numerical Analysis for Morphing Wing Structure,” Advanced Materials Research, Vol. 228-229, ISSN No. 1662-8985, pp. 169-173, April 19, 2001.

[0076]

[37] G.McKnight and C.Henry, "Large strain variable stiffness composites for shear deformations with applications to morphing aircraft skins," in SPIESmart Structures and Materials+Nondestructive Evaluation and HealthMonitorng, San Diego, California, 2008, pp. 692919-1 et seq.

[0077]

[38] C. Thill, J. Etches, I. Bond, K. Potter and P. Weaver, "Corrugated composite structures for aircraft morphing skin applications," in 18th International Conference of Adaptive Structures and Technologies, Ottawa, Ontario, October 3-5, 2007.

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[39] I. Dayyani, A. Shaw, E. Saavedra Flores, and M. Friswell, “The mechanics of composite corrugated structures: A review with applications in morphing aircraft,” Composite Structures, vol. 133, pp. 358–380, December 1, 2015.

[0079]

[40] S. Vasista, L. Tong and K. Wong, “Realization of Morphing Wings: A Multidisciplinary Challenge,” Journal of Aircraft, Vol. 49, No. 1, pp. 11ff., January-February 2012.

[0080]

[41] C. Thill, J. Etches, I. Bond, K. Potter, and P. Weaver, “Morphing Skins,” The Aeronautical Journal, pp. 117 ff., March 2008.

[0081]

[42] S. Kota, J. Hetrick, R. Osborn, D. Paul, and E. Pendleton, “Design and application of compliant mechanisms for morphing A / C structures,” in Smart Structures and Materials / SPIE, San Diego, California, August 14, 2003, pp. 24–33.

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[43] J. Mabe, F. Calkins and G. Butler, "Boeing's Variable Geometry Chevron: Morphing Structure for Jet Noise Reduction," in 47th AIAA / ASME / ASCE / AHS / ASCStructures, Struct.Dynamics, and Materials Conference, Newport, Rhode Island, May 1-4, 2006, AIAA 2006-2142.

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[44] G.Amendola, I.Dimino, A.Concilio, F.Amoroso and R.Pecora, "PreliminaryDesign of an Adaptive Aileron for the Next Generation Regional Aircraft," Journal of Theoretical and Applied Mechanics, Volume 55, Issue 1, Pages 307-316, Warsaw, 2017.

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[45] A. Rudenko, A. Hannig, H. Monner and P. Horst, "Extremely deformablemorphing leading edge: Optimization, design and structural testing," Journal of Intelligent Material Systems and Structures, Volume 29, Issue 5, Pages 764-773, 2018.

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[46] M. Friswell, "Morphing Aircraft: An Improbable Dream?," in ASME2014 Conference on Smart Materials, Adaptive Structures and Intelligent Systems, Newport, Rhode Island, September 8-10, 2014, SMASIS2014-7754.

[0086]

[47] R. Ajaj, C. Beaverstock, and M. Friswell, “Morphing aircraft: The need for a new design philosophy,” Aerospace Science and Technology, vol. 49, pp. 154-166, February 2016. Summary of the invention

[0087] The object of the present invention is to provide an efficient and safe flight control system for the aerodynamics of an aircraft and to provide an actuation solution that can be used, in particular, for such a flight control system.

[0088] This object is achieved by a control unit according to claim 1 and a force-actuated element according to claim 15 .

[0089] A method for controlling the pitch of an aircraft, in particular a method for controlling the pitch of an aircraft by means of such a control unit, is claimed in claim 14. An aircraft comprising such a control unit is indicated in claim 12. Further embodiments of the control unit, the method and the aircraft are indicated in the dependent claims.

[0090] A method for adjusting an aerodynamic or hydrodynamic element of an aircraft or a watercraft, in particular by means of such a force-actuated element, is claimed in claim 21. An aircraft comprising such a force-actuated element is indicated in claim 19. Further embodiments of the force-actuated element, the method and the aircraft are indicated in the dependent claims.

[0091] Therefore, the present invention provides a control unit for controlling the pitch of an aircraft during flight, the control unit comprising:

[0092] at least one adjustable control element that allows the pitch of the aircraft to be controlled during flight; and

[0093] The adjustment mechanism is used to actively adjust at least one control element to control the pitch of the aircraft.

[0094] The adjustment mechanism is suitable for controlling the pitch of the aircraft by means of a continuous deformation of at least one control element.

[0095] By deforming at least one control element continuously, problems associated with failures of designs based on control loops can be eliminated or at least reduced, since designs based on aerodynamics work independently based on physical principles. Using a continuously deforming control element, the pitch of the aircraft can be controlled in a particularly effective, reliable and safe manner. Compared with pitch control based on two or more separate, adjacent surface elements adjusted relative to each other in their positions, a significantly lower curvature is obtained using a continuously deforming control element. Therefore, using such a control unit can avoid the generation of shock waves, which are usually generated at the hinge of two separate control panels that are angled relative to each other. Drag waves and shock waves are possible causes of inefficient flight control and increased fuel consumption. At least one continuously deformable control element of the control unit can also be a continuously deforming element. A continuously deforming element can particularly realize the functions of a combined stabilizer and elevator. The use of a continuously deformable element rather than two or more discrete panels in pitch axis control also has the advantage that a significant weight saving can be achieved, since the semi-monocoque design can be followed more closely at the empennage (i.e., the tail of the aircraft), and the horizontal tail root can be fixedly connected to the fuselage, as compared to the prior art which requires a large cutout in the empennage to allow the horizontal stabilizer to pivot. Furthermore, pitch control based on one or more continuously deformable control elements greatly eliminates single points of failure, such as screw crane failure, trim motor loss, or flight control computer loss. Thus, although continuously deformable elements have been used in certain other aspects of flight control, it has been found that in the context of the present invention, both the efficiency and safety of pitch control of an aircraft during flight can be improved if based on continuous deformation of the control element.

[0096] Therefore, the control unit, preferably the main flight control unit, comprises at least one adjustable control element and an adjustment mechanism. The adjustment mechanism can (but not necessarily) be arranged in, on or at the at least one control element. The adjustment mechanism is preferably capable of influencing at least one control element, such as the position and / or orientation and / or shape of at least one control element, so that the pitch of the aircraft is changed due to changes in the airflow and / or deflection in the direct area of ​​the at least one control element. The at least one control element generally has at least one surface or surface element, which is exposed to the airflow during flight or can be exposed to the airflow during flight to control the pitch of the aircraft.

[0097] Controlling the pitch of an aircraft during flight means that the pitch of the aircraft can be changed or at least maintained, for example under different flight conditions, by means of a corresponding adjustment of at least one control element.

[0098] An adjustment mechanism is a mechanism that allows active adjustment of at least one control element, for example, of its position and / or orientation and / or shape. In contrast to passive adjustment, active adjustment generally requires the generation of a certain energy flow in order to intentionally adjust the corresponding control element against the forces of the inflowing air and, in the case of deformation flight control, also against the resistance of the deformed mechanical structure. Thus, in particular for the pilot of the aircraft, it is preferably possible to change the pitch angle of the aircraft with the aid of the adjustment mechanism. In order to enable the generation of the energy flow, an energy source arranged in, on or at the aircraft is generally provided, such as a generator, a battery or an accumulator for providing electrical energy, or a pump for providing a flow of hydraulic fluid under pressure.

[0099] An aircraft is a vehicle that is able to fly by obtaining support from the air. The aircraft is preferably an airplane, but may also be, for example, a helicopter.

[0100] Pitch is a well-known term in aviation that refers to the rotation of an aircraft about its lateral axis, which usually extends along the span of the aircraft's main wing. The pitch angle is measured between the aircraft's longitudinal axis and the horizon.

[0101] Continuous deformation means that the respective control unit is deformed without any singular points in the deformation of its overall shape. In contrast to, for example, two hinged panels which are adjusted in their position relative to each other, in the case of a continuously deformable control element a large part of the element or even the entire element is deformed. If a large part of at least one control element is continuously deformed or is capable of being continuously deformed by the adjustment mechanism, the large part of at least one control element preferably amounts to at least 10%, more preferably at least 30%, even more preferably at least 50%, most preferably at least 70% of the total volume of the control element.

[0102] At least one control element (which can also be referred to as a continuously deformable element) can be particularly suitable for implementing the function of a combined stabilizer and elevator, in particular a combined horizontal stabilizer and elevator. In this case, the at least one control element is essentially a horizontal stabilizer, a part of which can be adjusted in its position relative to the rest of the stabilizer in order to control the pitch of the aircraft. The continuously deformable element is preferably arranged in the tail of the aircraft, i.e. on the fuselage at a distance behind the main wing. The stabilizer is an aerodynamic element that provides longitudinal (pitch) stability for the aircraft. For the purpose of controlling the pitch, only a part of the at least one control element or the entire control element can be deformed continuously. By combining the pitch stabilization function and the pitch control function by means of a deformable horizontal tail, i.e. a continuously deformable element, both functions can be implemented simultaneously in a more efficient and safer manner, with a smaller drag wave and the elimination or at least attenuation of the shock wave.

[0103] Essentially, pitch control is the most critical of all flight controls. Compared to roll control, even though the moment arms can be as large as those of the ailerons, for pitch control the actual wing area is much smaller, or even more, because only the outer deformable horizontal tail section at higher angles of attack when twisted is primarily responsible for any primary flight control effect, which reduces the control area even further.

[0104] In particular, in the case of large transport aircraft, which are always equipped with horizontal tails of the swept-wing type, the characteristic shape of the swept wing that deforms under external loads is a preferred candidate for deforming the horizontal tail in order to maintain a closed, continuous and load-bearing wing skin, keeping the actuation forces as low as possible while meeting such requirements.

[0105] The at least one control element is preferably adapted not only to perform a pitch control function but additionally to perform a trim function. In this case, the adjustment mechanism is preferably adapted to control the trim of the at least one control element by means of a further continuous deformation of the at least one control element. In some embodiments, the adjustment mechanism can be adapted to control the trim of the control element independently of the pitch of the aircraft.

[0106] The control element is, for example, in the form of a continuously deformable element combining the functions of a horizontal stabilizer and an elevator of the prior art, and generally has a leading edge and a trailing edge relative to the direction of the normal airflow during flight. In this case, the adjustment mechanism is preferably adapted to deform the leading edge and the trailing edge in the same direction, and the adjustment mechanism is advantageously adapted to deform the trailing edge to a greater extent than the leading edge. Although the deformation of the control element under simple "twist" is basically sufficient to control the pitch, this generally results in large tensions in the skin and complex actuation. In the case of deformation of both the leading edge and the trailing edge, it is easier to achieve the same effect with respect to pitch control, since this is in any case the characteristic shape of a swept wing as it deforms under external loads, and therefore, by definition, the wing can be actively deformed into this shape even if it has a closed continuous skin.

[0107] In a particularly preferred embodiment, the adjustment mechanism is adapted to deform at least one control element for controlling pitch such that a maximum deformation of the at least one control element during flight results in a stall of the airflow around one or more regions of the at least one control element, which are generally the regions furthest from the fuselage. In certain cases, by producing a controlled stall in certain regions of the control element (e.g. the horizontal tail wing tip), the downforce caused by the control element can be reduced and the tendency of the aircraft to pitch up can also be reduced. As a result of the stall in a portion of the control element, the pitch-related lift of the control element preferably varies non-linearly with increasing deformation of the control element.

[0108] Therefore, the adjustment mechanism is preferably suitable for deforming the at least one control element so that, with increasing deformation of the at least one control element, the lift caused by the at least one control element changes non-linearly. In this way, it is possible to avoid undesirable behavior above a certain angle of attack, i.e. the restoring stabilizing dive moment after an aerodynamic disturbance no longer increases with increasing angle of attack, but rather flattens or even reverses. This undesirable behavior can be corrected electronically, but if it does not work as expected, it may have unpredictable negative consequences. Due to the non-linear effect of the relationship between the deflection of the deformable horizontal tail and the lift it generates, this corrective intervention on stability can be achieved by purely aerodynamic means (and is therefore predictable and reliable).

[0109] In a preferred embodiment, the adjustment means are adapted to gradually deform the at least one control element in two directions, a first direction from the leading edge towards the trailing edge and a second direction from the fuselage towards the lateral outer edge of the at least one control element. The main direction of deformation of the at least one control element preferably extends somewhere between the main leading edge direction of the control element and the longitudinal axis of the aircraft. This is also the deformation observed in the characteristic shape assumed by a swept wing when deformed under external loads. The nomenclature of the leading edge and the trailing edge is based on the direction of the inflowing air in normal flight.

[0110] In order to produce a continuous deformation of the at least one control element, the adjustment mechanism preferably has at least one force-actuated element, the at least one force-actuated element comprising:

[0111] a first end portion;

[0112] a second end portion;

[0113] A first sheet connecting the first end portion and the second end portion; and

[0114] a second sheet connecting the first end portion and the second end portion;

[0115] wherein the first sheet and the second sheet together enclose an inner space of the force actuation element, the inner space being adapted to be filled with a power transmission fluid, so that a change in the amount and / or pressure of the power transmission fluid in the inner space causes the first end and the second end to move relative to each other, in particular away from each other, along a force actuation direction to adjust at least one control element,

[0116] And wherein movement of the first end and the second end relative to each other along the force actuation direction is affected by pressure of the power transmission fluid acting on the first sheet and / or the second sheet along a main pressurization direction, which is transverse to the force actuation direction.

[0117] The force-actuated element is preferably located in at least one control element. In a particularly preferred embodiment, the adjustment mechanism has an actuator structure with a plurality of compartments, in each of which the indicated force-actuated element is arranged. Thus, in this case, at least one control element can be continuously deformed by the actuator structure with a plurality of force-actuated elements. The actuator structure and the skin may even form at least one control element of the aircraft together with other elements. By providing a plurality of force-actuated elements, not only a continuous deformation across the control element can be achieved, but also a fault protection system can be particularly achieved, that is, if only one or several force-actuated elements fail, the control element is deflected by other force-actuated elements or other elements, and a certain degree of degraded flight control effect can still be achieved. Arranging more than one force-actuated element in different compartments makes the following embodiments possible: For example, in this embodiment, the adjustment mechanism is suitable for increasing the distance between the first end and the second end of the first force-actuated element, and at the same time reducing the distance between the first end and the second end of the second force-actuated element, so as to adjust at least one control element based on the combined actuation of the first force-actuated element and the second force-actuated element. Therefore, in this case, the adjustment mechanism includes at least such a first force-actuated element and a second force-actuated element, and is suitable for increasing the amount and / or pressure of the power transmission fluid in the internal space of the first force-actuated element, while reducing the amount and / or pressure of the power transmission fluid in the internal space of the second force-actuated element, so as to adjust the control element in two directions with the help of both the first force-actuated element and the second force-actuated element.

[0118] In particular, in the case of an adjustment mechanism having an indicated actuator structure with multiple compartments, the adjustment mechanism preferably comprises a plurality of such force-actuated elements, which are arranged in rows and / or columns and preferably in different compartments. The arrangement of rows and / or columns enables a simple but effective deformation to be achieved by a large number of force-actuated elements.

[0119] Further preferred embodiments, explanations and details regarding the force-actuated element are given further below in this section.

[0120] The invention also relates to an aircraft comprising a control unit as indicated above, the control unit comprising at least one adjustable control element and an adjustment mechanism for actively adjusting the at least one control element to control the pitch of the aircraft. The control element is preferably firmly connected to the fuselage of the aircraft.

[0121] Furthermore, the present invention provides a method for controlling the pitch of an aircraft, in particular a method for controlling an aircraft as indicated above and / or controlling the pitch of an aircraft by means of a control unit as indicated above, the method comprising at least the step of controlling the pitch of the aircraft by means of actively adjusting at least one control element of the aircraft, wherein the pitch is controlled by means of continuously deforming the control element.

[0122] The invention also relates to a force-actuated element, in particular a force-actuated element for adjusting an aerodynamic element or a fluid dynamic element of an aircraft or a ship, the force-actuated element comprising:

[0123] a first end portion;

[0124] a second end portion;

[0125] A first sheet connecting the first end portion and the second end portion; and

[0126] a second sheet connecting the first end portion and the second end portion;

[0127] In which, the first sheet and the second sheet jointly enclose the internal space of the force-actuating element, and the internal space is suitable for being filled with a power transmission fluid, so that changes in the amount and / or pressure of the power transmission fluid in the internal space cause the first end and the second end to move relative to each other along the force-actuating direction.

[0128] Movement of the first and second ends relative to each other in the force actuation direction is affected by pressure of the power transmission fluid acting on the first and / or second sheets in a main pressurization direction which is transverse to the force actuation direction.

[0129] The force-actuated element can in particular be part of an adjustment mechanism of a control unit as further indicated above, and / or part of an aircraft as further indicated above.

[0130] Thus, by means of the indicated force actuation element, the actuation force is influenced by the pressure of the power transmission fluid, which pressure acts in a direction transverse to the force actuation direction. By generating the actuation force in this way, i.e., by using a fluid pressure acting transversely to the force actuation direction, a particularly efficient force application and force transmission can be achieved. The well-known force amplification technology is effectively combined with the pneumatic / hydraulic principle, so that a relatively high actuation force is generated at both ends of the first sheet with minimal actuation pressure. Firstly, because pressure can be applied over a relatively large area by means of the fluid, a large force caused by the pressure can be generated, which acts transversely to the first sheet, forcing the first sheet into a straightened shape.

[0131] In the context of this document, a sheet is considered to be a two-dimensionally extending component whose thickness is relatively small compared to its length and width.

[0132] The first end and the second end are preferably directly connected by the first sheet and / or the second sheet, which means that the first sheet and / or the second sheet are directly attached to the first end and the second end, or the first sheet and / or the second sheet itself forms the first end and the second end. However, in some embodiments, it is also conceivable that the first sheet and / or the second sheet only indirectly connects the first end and the second end, i.e. indirectly connects the first end and the second end via another element, in particular via another sheet.

[0133] Secondly, due to the force amplification, as the first sheet gradually straightens, the pressure required to produce a certain continuous lateral actuation force at both ends of the first sheet decreases. In theory, for a completely straight first sheet, the pressure decreases to zero, at which point the first sheet simply becomes a straight sheet under the compressive (buckling) load.

[0134] Related to the final deformation, when the lateral actuating force generated at both ends of the first sheet increases as the first sheet straightens, the required pressure initially increases as the first sheet gradually straightens until the required pressure reaches a certain peak value. After reaching the peak value, even if the first sheet further straightens, the required pressure still decreases. For a completely straight first sheet, theoretically, the pressure is reduced to zero again.

[0135] The introduction of the actuation force by means of a distributed pressure can avoid or at least significantly reduce mechanical structural problems that occur when large point forces are introduced laterally to a surface. The possibility of generating a relatively low actuation pressure to generate a relatively high actuation force makes it possible to reduce the size of the corresponding pressure-generating equipment that needs to be provided in or on the aircraft or vessel. The reduction in the size of the pressure-generating equipment not only reduces weight, but also allows safer operation of the aircraft or vessel. The pressure-generating equipment can be, but does not necessarily have to be, part of the indicated adjustment mechanism.

[0136] If the force-actuated element is used in an aircraft or a vessel, the movement of the first end and the second end relative to each other along the force-actuated direction is preferably used to adjust the aerodynamic element or the fluid dynamic element. The aerodynamic element or the fluid dynamic element can be in particular a wing or a stabilizer, for example, a horizontal or vertical stabilizer. If it is a wing, the force-actuated element can be adapted to deform the wing or at least a part of the wing so that the aircraft or the vessel produces a rolling motion. Therefore, the aerodynamic element or the fluid dynamic element has the function of an aileron in this case. If the aerodynamic element or the fluid dynamic element is a horizontal stabilizer, the force-actuated element can be adapted to deform the horizontal stabilizer so that a pitching motion of the aircraft or the vessel is produced. Therefore, the aerodynamic element or the fluid dynamic element has the combined function of a stabilizer and an elevator in this case. If the aerodynamic element or the fluid dynamic element is a vertical stabilizer, the force-actuated element can be adapted to deform the stabilizer so that the aircraft or the vessel produces a yaw motion. Therefore, the aerodynamic element or the fluid dynamic element has the combined function of a vertical stabilizer and a rudder in this case. Alternatively or additionally, the aerodynamic element or fluid dynamic element can have the function of a trim element. In this case, the force actuation element is preferably adapted to deform the aerodynamic element or fluid dynamic element so that the trim function is achieved by the aerodynamic element or fluid dynamic element.

[0137] Aerodynamic elements or fluid dynamic elements are generally understood to be elements that are specifically designed to have certain properties relative to the gas or liquid medium flowing around them. The shape of the element can be optimized, for example, to reduce air or water resistance and / or generate upward or downward lift.

[0138] Adjustment of the aerodynamic or fluid dynamic element preferably, but not necessarily, means deformation of the aerodynamic or fluid dynamic element. Thus, the force-actuated element is preferably used to deform the aerodynamic or fluid dynamic element, in particular to deform it continuously. In other embodiments, adjustment of the aerodynamic or fluid dynamic element can also mean a change in the position and / or orientation of the aerodynamic or fluid dynamic element.

[0139] However, the use of force-actuated elements is generally not limited to aircraft or ships. Instead, the indicated force-actuated elements can also be used in a variety of other devices and / or applications. Many other such devices and / or applications can be envisioned, in which the application of force can be used to actuate the element, particularly in situations where relatively low actuation pressures are required and a low actuator stroke is acceptable.

[0140] The first sheet is preferably attached to the first and second ends of the force actuation element.The second sheet is preferably attached to the first and second ends of the force actuation element.

[0141] In a particularly preferred embodiment, the first sheet and the second sheet are in each case made entirely in one piece. In this case, both the first sheet and the second sheet are preferably continuous sheets. At least the first sheet, advantageously also the second sheet, preferably has a certain flexibility and elasticity, so that when the force-actuating element is pressed by the power transmission fluid, the sheet bends.

[0142] In other embodiments, the first sheet and the second sheet can also each include multiple elements, or each is part of a unit including multiple elements. In the latter case, the first end and the second end are connected via another element by the first blade and the second blade, respectively. The unit can, for example, include multiple rods used as other elements, which are, for example, connected to each other by hinges.

[0143] The force generated at the end, ie the actuation force, generally acts along a straight line which passes through the first end and the second end and thus defines the force actuation direction.

[0144] The inner space of the force-actuated element is preferably completely or partially bounded by the first sheet and the second sheet. In order to be able to generate a pressure acting on the first sheet and the second sheet by means of a power transmission fluid, the inner space is preferably sealed to the outside. The seal is preferably against the power transmission fluid, i.e., prevents the power transmission fluid from escaping from the inner space, and more preferably also against the surrounding air, i.e., prevents air from entering the inner space.

[0145] An increase in the amount and / or pressure of the power transmission fluid in the interior space generally results in an increase in pressure, thereby generating a movement of the first end and the second end away from each other. On the other hand, a decrease in the amount and / or pressure of the power transmission fluid in the interior space generally results in a decrease in pressure, thereby generating a movement of the first end and the second end towards each other. Therefore, the force actuation element is preferably adapted to generate actuation forces in two opposite directions.

[0146] It is stated that the main pressing direction is transverse to the force actuation direction, but this does not mean that the two directions must be at an angle of exactly 90° to each other. Instead, it means that the force actuation direction deviates significantly from the pressing direction, in the sense that the pressure exerted by the power transmission fluid on the surface of the first and second sheets in the direction of the local normal of the respective sheets is converted into a force extending in the longitudinal direction or extending in the first and / or second sheets.

[0147] The power transmission fluid is preferably a liquid. The liquid can in particular be water or a fuel, in particular kerosene, which can be used to power an aircraft or a ship. If a fuel is used as the power transmission fluid for the actuator, the corresponding fuel can be used for other means, for example, partially and only for emergency purposes to power the aircraft or the ship. However, the total fuel requirement for such actuation is very low compared to the amount of onboard fuel required to power the aircraft or the ship. By using fuel as the power transmission fluid, it is possible to reduce the overall system weight due to the ability to combine several system functions within the aircraft or the ship. In other embodiments, a power transmission fluid in the form of a gas (e.g. air or powder) is also conceivable.

[0148] In order to avoid or at least reduce feedback forces and / or torques acting on the first sheet and / or the second sheet from counter bearings at the first end and the second end, the force-actuated element preferably includes a first rotary bearing and / or a second rotary bearing at both ends, the first rotary bearing connecting the first sheet with the second sheet at the first end, and the second rotary bearing connecting the first sheet with the second sheet at the second end.

[0149] The invention also relates to an aircraft comprising an aerodynamic element, in particular a wing, a combined stabilizer and elevator, an elevator, an aileron or a stabilizer, and at least one force-actuated element as indicated above, for adjusting the aerodynamic element, preferably deforming the aerodynamic element, more preferably continuously deforming the aerodynamic element. In certain preferred embodiments, at least one force-actuated element is filled with fuel of the aircraft, in order to use the fuel as a power transmission fluid.

[0150] Furthermore, the present invention provides a method for adjusting an aerodynamic element or a fluid dynamic element of an aircraft, preferably an aircraft or a ship as described above, in particular adjusting an aerodynamic element or a fluid dynamic element by means of a force actuation element as indicated above, the method comprising at least the following steps: changing the amount and / or pressure of a fluid transport fluid in an interior space enclosed by a first sheet and a second sheet to generate a pressure acting on the first sheet and / or the second sheet along a main pressurization direction, and thereby causing the first end and the second end connected to each other by the first sheet and the second sheet to move relative to each other along a force actuation direction transverse to the main pressurization direction to adjust the aerodynamic element or the fluid dynamic element. BRIEF DESCRIPTION OF THE DRAWINGS

[0151] The following describes preferred embodiments of the present invention with reference to the accompanying drawings, which are for the purpose of illustrating preferred embodiments of the present invention and are not intended to be limiting.

[0152] Figure 1shows a perspective view of a wing of an aircraft according to the prior art;

[0153] Figure 2 shows a side view of the rear of an aircraft according to the prior art;

[0154] Figure 3 schematically shows a perspective view of the tail of an aircraft according to the prior art, wherein the elevator is attached by means of a hinge;

[0155] Figure 4 A schematic perspective view of the tail section of an aircraft according to the invention is shown with a continuously deformable deformation element which is suitable in each case for realizing the combined function of a horizontal stabilizer and an elevator;

[0156] Figure 5 Shows Figure 3 The horizontal tail downforce evolution and the local angle of attack of the horizontal tail of the aircraft of the prior art are shown;

[0157] Figure 6 Shows Figure 4 The horizontal tail downforce evolution and the local angle of attack of the horizontal tail of the aircraft of the present invention are shown;

[0158] Figure 7 Shows Figure 3 The pitch moment coefficient of the horizontal tail of the prior art aircraft shown;

[0159] Figure 8 Shows Figure 4 The pitch moment coefficient of the horizontal tail of the aircraft of the present invention is shown;

[0160] Fig. 9 Schematically shows a perspective view of the tail of another aircraft according to the present invention, which has a continuously deformable element including an elevator attached by a hinge to realize the functions of both a horizontal stabilizer and an elevator;

[0161] Fig.10 Schematically showing an element of an aircraft or a watercraft, which element can be partially filled with a power transmission fluid to deform;

[0162] Fig.11 Schematically shows an element of an aircraft or a watercraft according to another preferred embodiment, the element being deformable by means of a power transmission fluid;

[0163] Fig.12 Schematically shows a side view of a force-actuated element for adjusting an aerodynamic element or a fluid dynamic element of an aircraft or a ship according to a first embodiment of the invention;

[0164] Fig.13Shows Fig.12 A perspective view of a force actuating element;

[0165] Fig.14 schematically shows a side view of a fluid-filled force-actuated element for adjusting an aerodynamic element or a hydrodynamic element of an aircraft or a watercraft according to a second invention embodiment;

[0166] Fig.15 Shows Fig.14 A perspective view of a force actuating element;

[0167] Fig.16 schematically shows a side view of a fluid-filled force-actuated element for adjusting an aerodynamic element or a hydrodynamic element of an aircraft or a watercraft according to a third invention embodiment;

[0168] Fig.17 Shows Fig.16 A perspective view of a force actuating element;

[0169] Fig.18 Schematically shows an element of an aircraft or a watercraft according to a preferred embodiment, which element is continuously deformable by means of a power transmission fluid;

[0170] Fig.19 Schematically shown for explanation purposes is an element of an aircraft or a watercraft, which element is continuously deformable by means of a plurality of hydraulic pistons;

[0171] Fig. 20 schematically illustrates a force actuated element under compressive load, the element being pressurized (over-pressurized) by a gas power transmission fluid;

[0172] Fig.21 schematically illustrates a force actuated element under compressive load, the element being pressurized (over-pressurized) by a hydrodynamic transmission fluid;

[0173] Fig. 22 Three possible variations of the design of the force-actuated element of the present invention are schematically shown;

[0174] Fig.23 A perspective view showing a first version of a force actuated element of the present invention;

[0175] Fig.24 A perspective view showing a second version of the force-actuated element of the present invention;

[0176] Fig.25 a perspective view showing a third version of the force actuation element of the invention; and

[0177] Fig.26 A perspective view of a load bearing first sheet including fins / ribs is shown. DETAILED DESCRIPTION

[0178] In the attached drawings, several different embodiments of the invention are shown: a control unit of an aircraft or at least a part thereof, and method steps for controlling the pitch of an aircraft, as well as a force-actuated element of an aircraft or at least a part thereof, and method steps for adjusting an aerodynamic element or a hydrodynamic element of an aircraft or a watercraft. The same reference numerals are used to designate elements of the different embodiments having the same or similar function.

[0179] Figure 1 The outer part of a wing 3 of an aircraft according to the prior art is shown, comprising a structure with large discrete control surface panels which are connected to each other by hinges and can be actuated by respective actuators to control the flight path of the aircraft. Figure 1 As can be seen in the figure, the wing 3 includes an inner flap 31, an inner aileron 32, an outer flap 33 and an outer aileron 34 at its trailing edge. Due to the large discrete control surface panel, the possible size of the wing torsion box is reduced. For large aircraft and / or extended wingspan, the problem of aileron reversal may occur at high speeds. Therefore, a second inner aileron 32 is usually provided, which further increases the weight and complexity of the wing 3. When the shape of the wing 3 is adjusted by means of flaps 31, 33 and / or ailerons 32, 34, in order to control the flight characteristics of the aircraft during flight, drag waves will be generated, especially in the area where different control surfaces are connected to each other by hinges, and shock waves will be generated at high speeds.

[0180] Figure 2 The rear part of a conventional aircraft 1 according to the prior art is shown. The aircraft 1 comprises a fuselage 2, to which wings 3 are attached. In the tail region of the aircraft 1, a vertical stabilizer 4 with a rudder 41 and a horizontal stabilizer 5 with an elevator 51 are connected to the fuselage 2. The horizontal stabilizer 5 is formed in each case by an aerodynamic element, which serves to control the pitch of the aircraft 1 during flight. The elevator 51 of the horizontal stabilizer 5 forms a control element, which can be adjusted relative to the rest of the horizontal stabilizer 5 in order to change the pitch of the aircraft 1 or to maintain the pitch under changing flight conditions.

[0181] In order to keep the elevator 51 in an upwardly deflected position relative to the rest of the horizontal stabilizer 5, a trim tab is usually attached to the trailing edge 5 of the elevator 51 when the angle of attack of the aircraft 1 is increased. Figure 2 The horizontal stabilizer 5 (not shown) can relieve the pressure exerted on the control wheel by the air flowing in along the airflow direction A. However, due to the existence of the separate control surface, the horizontal stabilizer 5 according to the prior art is prone to generate drag waves and shock waves. In particular, when deflecting, shock waves may occur at the hinge between the horizontal stabilizer 5 and the elevator 51.

[0182] As an alternative to trim tabs, the Bell X-1 introduced a trimmable horizontal stabilizer as early as 1946. For this purpose, the horizontal stabilizer 5 is not firmly connected to the fuselage 2, but can be pivoted relative to the fuselage 2 (for example, by means of an actuator system 54 based on a screw crane). Figure 2 ). This allows pitch trimming by adjusting the horizontal stabilizer 5, thereby changing the required downforce while maintaining a zero deflection angle between the elevator 51 and the horizontal stabilizer 5, avoiding shock waves at high speeds. However, in order for the horizontal stabilizer 5 to pivot, a large cutout 21 must be present on the fuselage 2.

[0183] The provision of a pivotable horizontal stabilizer 5 is both simple and robust; by means of a screw-crane based actuator system 54 the horizontal stabilizer 5 is slow and therefore not intended for primary flight control, but for trimming purposes. However, in the event of a malfunction, such as uncontrolled rotation of the screw-crane, the pilot needs to intervene immediately, a situation that can easily become extremely dangerous. Under this systemicity, so-called pitch trim loss of control has led to many fatal accidents in the past.

[0184] Figure 3 and Figure 4 The tail section of an aircraft with a conventional horizontal stabilizer 5 according to the prior art is shown ( Figure 3 ), the horizontal stabilizer 5 has an elevator 51 connected by a hinge. In contrast, the tail of the aircraft of the present invention ( Figure 4 ) has a continuously deformable element 6 which realizes the combined function of stabilizer and elevator. It is easy to see that by deflecting the continuously deformable element 6 in the upward direction in the region of the trailing edge 53, a pitch control effect similar to that achieved by means of the elevator 51 can be achieved. However, in contrast to Figure 3 Compared with the horizontal stabilizer 5 with discrete control surfaces, Figure 4 The continuous deformation element 6 is closer to the swept wing that continuously deforms under load. Since there is no deformation singularity in the absence of a hinge, the control element (i.e. Figure 4 The continuous deformation element 6 in the tail of the aircraft will generate fewer drag waves and shock waves. Due to the uninterrupted semi-monocoque fuselage design of the tail of the aircraft and the continuous deformation element 6 firmly connected to the fuselage 2 and without a separate elevator 51, Figure 3 This significantly reduces overall system weight compared to a conventional tail section. It also eliminates single points of failure such as screw crane failure, trim motor loss, or flight control computer loss.

[0185] Furthermore, the continuously morphing tail also includes a trim function that is hingeless and correspondingly eliminates or reduces shock wave vulnerability. Thus, from this perspective, the continuously morphing tail eliminates the need for a separate trim tab or trimmable horizontal stabilizer. For example, by designating an internal area of ​​the morphing tail dedicated to the trim function, or by integrating a simple offset in the control loop, any trim function that releases the control force can be incorporated.

[0186] Figure 5 and Figure 6 The following is a diagram showing an aircraft with a conventional horizontal tail according to the prior art. Figure 5 ) and for an aircraft having a continuously deformable horizontal tail according to the present invention ( Figure 6 ), the characteristic of the lift evolution with the deflection of the pitch control element (i.e. the elevator 51 or the continuously deformable element 6). In each case, α a , α b , α c and α d represent the span of the elevator at a specific position y a ,y b ,y c and d The local angle of attack of the horizontal tail at , which depends on the conversion into an elevator deflection command δ e The flight control signal δ S (e.g. specified by the pilot control unit). C l Indicates the local lift coefficient generated at the corresponding position of the elevator 51 or the continuously deformable element 6.

[0187] When using a conventional horizontal tail, the elevator 51 deflects the same angle over its entire wingspan to achieve a total lift of the aircraft L T .from Figure 5 As can be seen from the bottom graph, the total lift L T According to the deflection δ of the elevator 51 e Changes linearly.

[0188] When the horizontal tail of the present invention is used, the total lift of the aircraft L T With small deflections of the deformable elevator, the Figure 6 bottom graph). However, as the deflection increases, the total lift L T The curve flattens out, and depending on the specific design, the total lift L TThe curve may be reversed due to the deformed elevator tip reaching its local stall angle of attack. Thus, with the continuous deforming element 6, the lift achieved by the corresponding control surface varies nonlinearly with increasing deformation of the control element (i.e., here the continuous deforming element 6). The nonlinear curve of the resulting lift is due to the airflow separation 11, which eventually evolves in the outer area of ​​the continuous deforming element 6 with increasing deformation. This partial stall region of the actuated deformed horizontal tail may change the downforce. As a result, the overall lift distribution changes significantly, introducing nonlinearities and thus changing the maneuvering characteristics of the aircraft. In an ideal case, if the stall region on the horizontal tail is designed to limit (cap) the downforce at a specific aircraft angle of attack, it is even possible to envision aerodynamic pitch limits close to the stall angle of a specific main wing of the aircraft. However, the occurrence of a stall on an airfoil does not mean that the lift disappears. Depending on the specific airfoil, a more or less sharp drop in lift may be observed initially, but when the angle of attack is further increased, the lift increases again with the drag, which is consistent with what happens on a simple flat plate.

[0189] It should also be noted that in this case, the stall angle of the entire aircraft (which is relevant to flight safety) is independent of any stall occurring at the horizontal stabilizer, since the stall occurring at the horizontal stabilizer is primarily dependent on the total washout angle achieved on the horizontal stabilizer by deformation; the two stall issues are independent of each other. If the angle of attack of the entire aircraft is increased, this actually reduces any stalling effect on the deformed portions of the horizontal stabilizer, since this effectively partially rotates those outer stalling portions out of the stall condition. This correlation is therefore far from insignificant and requires fine tuning to achieve the desired self-limiting effect at the angle of attack.

[0190] Figure 7 and Figure 8 shows the pitching moment C m -α curves, which are a useful tool for investigating the static stability of an aircraft, as these curves depict the pitch-down moment as the angle of attack (here, the angle of attack of the entire aircraft) changes, i.e. the tendency to return to a neutral and stable pitch angle.

[0191] Ideally, the curve is a straight line with a negative inclination angle. In this case, the aircraft is considered statically stable, which means that the further the aircraft's angle of attack deviates from the stable value, the stronger the aircraft's tendency to return to that value. The steeper the slope of the curve, the greater the stability; a horizontal curve represents neutral stability, and a positive inclination indicates instability.

[0192] Figure 7The curve in shows a deviation from a straight negative slope at larger angles of attack (see arrow), which shows neutral stable behavior when approaching a stall, which is essentially similar to the problems experienced by the Boeing 737MAX aircraft, which ultimately led to the introduction of the MCAS system.

[0193] Conventional flaps such as elevators theoretically change the lift curve of the airfoil linearly when deflected. Therefore the basic stability problem of the aircraft approaching a stall remains unchanged, and Figure 7 The curves in all show the same basic shape as the deflection of the standard elevator increases. In general, a standard elevator does not drastically change the basic behavior of the airplane; the bank is approximately the same.

[0194] Figure 8 The curves for the case of the deformable elevator are shown in Figure 2. As expected, the curves for the unactuated horizontal tail are similar to Figure 7 However, once the morphing elevator is deflected, it can be clearly observed that the lift generated by the morphing elevator evolves in a nonlinear manner.

[0195] Relatively small deflections from the unactuated state cause significant changes in the moments, indicating that flight control is robust, especially for smaller deflections caused by the full morphing tailplane area deflection rather than just the smaller standard elevator area, which is particularly important considering that a morphing tailplane should ideally include a trim function. Some limited further deflections do not qualitatively change the results, indicating a more linear lift evolution up to that point, with the basic unstable behavior of the aircraft not being altered.

[0196] With further deflection up to full deflection, nonlinear effects on the curve due to stall begin to appear at the horizontal tail tip. Beyond a certain angle of attack, the curve flattens significantly, indicating an overall reduction in static stability. However, in addition to the flattening of the curve, the statically unstable region moves to higher angles of attack or disappears completely (see arrows). This behavior is very different from the behavior observed for prior art elevators.

[0197] This stabilizing effect is indeed particularly important when the nose is fully pitched up, because this full deflection of the elevator is usually only needed at slow speeds and close to the stall angle of attack of the entire aircraft. At this time, a large angle of attack is reached and the kinetic energy of the aircraft is low, so the consequences of a stall are more dangerous than at high speeds. During this slow flight close to the stall, any unstable behavior will quickly aggravate the stall and even lead to an unrecoverable stall.

[0198] At higher speeds, the elevator deflection required to achieve a certain pitch angle is much smaller, and although instabilities are undesirable in themselves, the aircraft is less likely to reach such a high angle of attack region at high speeds where these instabilities will become a problem. In this case, in addition to the possibility of overload due to the higher aerodynamic forces caused by larger elevator deflections, the aircraft is more likely to start a steep climb at a large pitch angle before eventually slowing down and increasing its angle of attack. Due to the higher speed, the elevator is more effective, which is combined with the higher kinetic energy state of the aircraft.

[0199] Fig. 9 The model aircraft tail section shown also represents a possible embodiment of an aircraft with the control unit of the invention in combination with a prior art elevator. Installing both types of pitch control elements simultaneously has the advantage that two independent systems are available. This also means redundancy, which is important for certification. There is no need to switch immediately specifically to a completely new system with independent deformation pitch control, but the well-known aspects of the separate elevator panels are still there, which also makes certification easier, for example, during the transition period.

[0200] In the following, different possible adjustment mechanisms for actuating primary flight control elements (eg, pitch control elements) will be discussed, including adjustment mechanisms employing force actuation elements with the present invention. Requirements for primary flight controls and various actuation media, including their advantages and disadvantages, are also considered.

[0201] Requirements for Transformer Flight Control

[0202] The application of morphing techniques to aircraft flight control must take into account several requirements, all of which ideally must be met simultaneously. Although for some requirements less than ideal compromises can be accepted in order to perfectly match other criteria, some are indeed mandatory.

[0203] Each morphing technology generally relies on a specific "actuation medium" that defines the usefulness of the resulting technology for a specific purpose, in this case active aircraft primary flight control. Most morphing technologies to date rely in one form or another on the following available actuation media:

[0204] Internal mechanical: bistable structure, shear center adjustment

[0205] (Electro) thermal type: shape memory alloy, material property adjustment

[0206] Direct electric type: piezoelectric element

[0207] Direct (electromechanical) mechanical: servos and levers

[0208] Hydraulic and pneumatic: pistons, servos and levers

[0209] In order for the technology to be useful in flight control, fast actuation rates need to be achieved. This applies particularly to primary flight controls, such as elevators, ailerons, and rudder, for which short propagation times are required to achieve fast flight responses. The time required to receive the appropriate signal to achieve full control deflection varies by aircraft type; it is typically around 0.5 seconds to 1 second for transport category aircraft. For some secondary flight controls, such as flaps, slower actuation is acceptable. For primary flight controls, this rules out slower solutions such as thermally actuated mechanisms.

[0210] While staggered deflections are acceptable in secondary flight controls, this is not the case for primary control surfaces: the mechanism must enable the full range of motion of the control to be used, and any intermediate settings must be precisely achieved and controllable. The propagation must be the same in both directions, without any unusual friction / hysteresis behavior. Mechanisms that rely on bistable characteristics, although attractive from a high load-deformation point of view, are excluded as candidates for use in primary flight controls unless they are modified in some way. However, even with the staggered stable states of such structures, the accuracy required to achieve intermediate actuation states may not be achieved. Thermal actuation is also problematic from this perspective, as very precise control of temperature is required, which is difficult to achieve in the wildly varying environment of an operating aircraft.

[0211] For deformation mechanisms capable of withstanding high loads, it is expected that the actuation force or actuation power delivered by the actuator will also be very high. In order to reduce the risk of actuator burning, premature fatigue or simply to reduce the required control actuation power, it is not only very desirable but indeed necessary to accurately lock the actuator in its deflected state. This locking and unlocking must be a very fast process to achieve fast actuation rates, so it is recommended to combine the locking mechanism with the actuation medium supply control. This effectively rules out systems using piezoelectric elements, which continuously require high electrical actuation currents; similar considerations apply to direct control and electric servo systems unless these systems are mechanically modified to block any feedback. Hydraulic actuators easily meet this requirement by blocking the control valve, thereby holding the actuator in place.

[0212] Regardless of which actuating mechanism is used, the actuating medium needs to be supplied in the required amount and conditions. For example, the voltage required by the piezoelectric element must be generated and supplied, and the hydraulic fluid must be supplied in the appropriate amount and pressure. In order to avoid feasibility problems, any deformation mechanism actuator should rely on supplies such as motors and pumps that can meet these prerequisites without major modifications. In this example, this will specifically mean that such an actuator does not require any excessively high voltage or pressure.

[0213] In large transport aircraft where human lives are at stake, safety aspects are paramount. Therefore any reasonable measures must be taken to mitigate unnecessary risks and hazards. In the context of deformation mechanism actuation, this means avoiding any system combination that itself poses a risk. Using electrically controlled piezoelectric elements in a wing structure filled with flammable fuel can be risky, especially considering the wiring that needs to carry high voltages. Even minor wire wear or damage (not uncommon in the ever-present vibration and fatigue environment of aircraft operation) can lead to arcing, fuel fire in the wing tanks and aircraft loss.

[0214] In addition, transport aircraft must be able to withstand a certain degree of lightning damage. Lightning strikes can affect the piezoelectric elements, either damaging them and thus rendering them unusable, or actuating them in an undesirable and uncontrollable manner. Either way is unacceptable for flight control purposes. Similar considerations apply to any electrical control system.

[0215] Prior art semi-monocoque wing designs that use spars, ribs and stringers to support thin skins may not be fully supported. Some instabilities must be introduced and then stabilized again in a variable and controlled manner. Nevertheless, deviations from prior art wing designs should be kept to a reasonable minimum, and the resulting structure must adequately replace the functions of spars, ribs and stringers in a composite structure.

[0216] At least, any actuation concept should not conflict with other systems.Preferably, the system will provide several functions simultaneously.

[0217] Many parts and mechanisms of an aircraft fulfill several functions. For example, in some older models, such as the A300 or Concorde, the hydraulic actuators that actuate the landing gear also serve as the sidestays of the landing gear. Another typical dual function can be seen in the integrated fuel tank, in which the wing structure itself plus some sealants doubles as the fuel tank. Placing the fuel in the wing tank has several structural and payload benefits. Therefore, it is preferred to retain such an arrangement in the morphing wing design, allowing the fuel to be stored inside the morphing wing.

[0218] The complexity of the mechanism increases the weight and increases the risk of failure. Therefore, it is desirable to keep these mechanisms as simple as possible, with a small number of moving parts and without hinges that may block or break. Flexible mechanisms perform well in these respects, so it would be beneficial to exploit at least some of the beneficial properties of flexible mechanisms in high-load deformation mechanisms. Mechanisms that are kept simple also generally require simpler and cheaper maintenance. From a safety perspective, such mechanisms should be designed to be operable after failure, allowing other components of the mechanism to partially take over the function in the event of a catastrophic failure, or at least such mechanisms should be designed to be fail-safe so that failure of the entire system occurs in a controlled manner without causing additional harm.

[0219] To date, no known wing morphing technology has been developed that takes all of these requirements into account. All of this combined suggests that there are good reasons to use hydraulic or pneumatic actuation.

[0220] As the only actuation medium that simultaneously meets the most critical requirements, including fast actuation rates, precise and continuous controllability, and lockability, wing morphing systems for large aircraft will most likely need to be based on hydraulic technology. Traditionally, the main disadvantage of hydraulic actuators is their relatively high weight; in addition, their high pressure requirements and their characteristic of introducing very high point loads are particularly problematic.

[0221] However, all three of these problems can be overcome by using the force-actuated element of the present invention which is lightweight, simple to control with low pressures, forms an integral part of the wing structure, and transmits very high actuation forces in the form of line loads.

[0222] Requirements for deformation of high-load wings

[0223] Attempts to transition active morphing technologies developed at the small aircraft scale to actual high-load wings have so far been thwarted because these existing technologies often fail to fully address a major issue when considering large transport aircraft: the wing loadings on such aircraft are 25 times that of small UAVs. To scale to larger aircraft sizes, the wing structure and morphing mechanism need to maintain a similar order of magnitude stiffness-to-weight ratio at each scale.

[0224] The structural solutions that can meet this requirement are likely to show geometrical similarity with respect to load-bearing functions across all aircraft size classes. For good reasons, the semi-monocoque fuselage design approach with closed, continuous and load-bearing wing skins has proven to be the state of the art for several decades, achieving acceptable stiffness-to-weight ratios for all wing sizes. The structural solutions are geometrically similar for both large and small aircraft wings.

[0225] Any deviation from this (e.g., by including very flexible components in the wing skin) is likely to result in a reduction in the overall stiffness-to-weight ratio, since the load-bearing capacity of the wing skin is reduced. Therefore, a closed, continuous and load-bearing wing skin should be retained. Any recognized high internal forces and actuation capable of generating such high forces must be used while meeting the requirements of morphing flight control as described above. The force-actuated element of the present invention is such an actuation.

[0226] Deformation of high-loaded wings

[0227] For a given passive load distribution, the passive deformation results in a specific shape, which for the horizontal tail may be similar to Figure 4 The shape shown. Because this is a shape caused by external loads, it is indeed possible to achieve this shape while maintaining a closed, continuous wing skin. Conversely, it is possible to achieve the same shape with the same skin through internal actuation. It is therefore of interest to verify how this shape can be adjusted so that different shapes and in particular different overall wash angles can be achieved for the same passive loads if the displacement of the leading edge is gradually constrained starting from the wing root. For a typical wing planform, such as Figure 4 In the case of the continuous deforming element 6 in FIG. 5 , if the same upwardly displaced contour lines I are observed, the following occurs: as the amount of constraint of the leading edge 52 increases, the angle between these contour lines I and the longitudinal axis of the aircraft increases, and the achievable twist / washout also increases, including the aerodynamic effect of this twist washout on lift. In fact, as the twist caused by the increased leading edge constraint increases, the contour lines I turn in the span direction.

[0228] It should be noted that the contour line I, when projected onto the wing plan view, generally remains almost straight, with a very moderate curvature away from the wing root near the leading and trailing edges. From a deformation point of view, this means that almost no deformation occurs in that direction, and if a section is cut along this line, the chord line of the resulting airfoil section will remain almost unchanged. On the contrary, in a section cut at right angles to the contour line I, the chord line of the resulting airfoil changes in a manner equivalent to the camber deformation. It can therefore be concluded that for a swept wing to obtain a characteristic shape identical to the characteristic shape caused by passive deformation, a mechanism equivalent to the camber deformation is required along the axis located between the leading edge of the swept wing and the longitudinal axis of the aircraft.

[0229] Internal actuation of a high load wing with closed skin

[0230] Given the desired basic external shape and type of deformation of the wing set, the problem remains how to actuate the deforming portion of the wing (or flight control element). The forces generated by the closed wing skin are expected to be very large and exceed the capabilities of most existing actuators.

[0231] Gandhi

[27] and Runkel

[28] investigated existing passive deformation techniques that could be applied to high-load situations. In particular, Gandhi's solution showed the possibility of achieving high-load camber deformations by introducing bistable buckling elements as diagonals in the trapezoidal torsion box of the wing. In this way, two different and stable positions are achieved by allowing deformations similar to an oblique parallelogram or trapezoid. Runkel's approach is very similar, effectively replacing the rib / torsion box with a shear mesh made of chiral structures that stabilizes in two different positions (stabilized by the same type of oblique parallelogram motion) depending on whether an external load is applied.

[0232] The most important requirement for using morphing technology for active flight control purposes is its ability to achieve fast, continuous and precisely controllable movements. Gandhi's Runkel solutions are both stable in only two different positions. However, Gandhi's approach is actually the opposite of the very old camber morphing solution proposed by Parker in 1920 ([3]), in which the wing consists of a number of trapezoidal "torsion boxes" whose diagonal lengths can be controlled by pulling control wires. It is thus possible to precisely control the diagonals by adjusting their lengths, thereby achieving continuous and precisely controllable camber morphing movements. Here, the movement of such a "torsion box" can also be imagined as the movement of a tilted parallelogram or trapezoid. Such a torsion box thus becomes a control element.

[0233] Therefore, the next question to be answered is how best to introduce control forces to produce forces large enough to continuously deform the control element and ultimately the wing having a closed, continuous skin built around the control element.

[0234] Fig.10 , an embodiment of a possible adjustment mechanism 7 is shown, which is based on pressing the power transmission fluid 10 into several constrained internal spaces 85 to achieve deformation. Fig.10As can be seen in the figure, the adjustment mechanism 7 according to the present embodiment includes an actuator structure 71, which has two compartments 72 here. In each compartment 72, a carrier sheet 83 is arranged to extend laterally from one corner of the corresponding rectangular compartment 72 to the opposite corner, thereby defining an internal space 85 with an approximately triangular cross-section. The carrier sheet 83 extends from a first end 81 to a second end 82, and the ends 81 and 82 are connected to each other via two vertical walls of the actuator structure 71. One of these vertical walls forms a partition wall, which separates the two compartments 72 of the actuator structure 71 from each other. The partition wall may also be referred to as a second sheet 84, because its shape is generally similar to that of the first carrier sheet 83. Similar to the carrier sheet 83, the second sheet 84 also at least partially surrounds the corresponding internal space 85 of each compartment 72, and connects the first end 81 to the second end 82 via at least another plate-like element, which is formed by another wall of the actuator structure 71. In the example Fig.10 In the undeformed state shown in the left figure of , the carrier sheet 83 is made more flexible than the second sheet 84 and is convexly curved toward the liquid-filled interior space 85.

[0235] In each case, the force actuation element 8 is formed by a respective carrying first sheet 83, a second sheet 84 and a respective first end 81 and second end 82. The partition wall of the actuator structure 71 separating the two compartments 72 from each other thus forms the second sheet 84 of each of the two force actuation elements 8.

[0236] The gap between the convex side of the carrier sheet 83 and the actuator structure 71 (i.e., the interior space 85) becomes a closed volume that can be filled with the power transmission fluid 10 under pressure. If an incompressible fluid is used, the closed volume directly and precisely controls the geometry of the parallelogram (see Fig.10 Right side). Fluid-assisted force actuation can thus be achieved, thereby achieving deformation of the adjustment mechanism 7 and any control element attached to the adjustment mechanism 7.

[0237] In order to reduce the demand for the power transmission fluid 10, another second sheet (also called a stopper) can be introduced on the convex side of the carrier sheet 83 to enclose the power transmission fluid 10 in a smaller volume, such as Fig.11As shown in the embodiment of FIG. 1 . Since its only purpose is to contain and enclose the power transmission fluid 10, the stopper (or second sheet 84) is only under positive hydrostatic or pneumatic pressure, and its geometry is selected to reduce the volume of the power transmission fluid 10 to nearly zero when the load-bearing sheet 83 reaches its desired maximum compression state (i.e., achieves the minimum diagonal length). Due to its appearance, the force-actuating element 8 generated is named a crescent actuator (CSA), which has two versions: a crescent hydraulic actuator (CSHA) and a crescent pneumatic actuator (CSPA).

[0238] exist Fig.12 and Fig.13 In the simplified embodiment of the CSA shown, the curved sheets 83, 84 of the CSA are replaced by straight panels 91, 92 and 93, 94 joined by belt hinges 95, 96, 97. Actuation is accomplished by introducing an internal load in the internal space 98 along the main pressurization direction P, which can be achieved, for example, by rotating a camshaft. The first rear panel 93 and the second rear panel 94 are connected to each other via a rear belt hinge 96 and to the load-bearing first front panel 91, 92 via a connecting hinge 97, and the first rear panel 93 and the second rear panel 94 effectively act as a support or support for the lateral force required to push downward the front belt hinge 95 connecting the load-bearing first front panels 91, 92. Once the load-bearing first front panels 91, 92 form a straight line in the fully extended state, the required normal force is reduced to zero and the stopper or rear panel, respectively, is not loaded.

[0239] In order to achieve this function as a support, assuming that the crescent-shaped actuator (force actuation element 8) as a whole resists the compression force generated by, for example, a compression spring, the rear panels 93 and 94 are always subject to tension, while the front panels 91, 92 are always subject to compression. This implies that the load-bearing front panels 91, 92 must be dimensioned to withstand buckling and therefore require greater strength than the stopper panels 93, 94, which are only subject to tensile stress. The first and second ends of the force actuation element 8 are formed here by a connecting hinge 97.

[0240] Although the pressure acting in the main pressurization direction P can be generated substantially by, for example, a hydraulic piston or a mechanical extender driven by a screw jack, according to a preferred embodiment, a power transmission fluid is used, for example water, fuel or air. This allows a uniform force to be introduced into all the sheets, avoiding stress concentrations. In addition, a large part of the panel is thus in contact with the pressurized fluid, allowing a relatively low pressure (acting in the main pressurization direction P) to generate a large total lateral force (in the main actuation direction F).

[0241] In all embodiments, the total resultant force achieved by such a force-actuated element 8 (which can but need not form part of an adjustment mechanism 7 for adjusting, for example, a control element of an aircraft) is directed from the second end 82 to the first end 81 (or from the first end 81 to the second end 82), i.e. along the force-actuated direction F.

[0242] If forced Fig.12 and Fig.13 The planar plate force actuator element 8 is filled with a fluid 10, such as Fig.14 and Fig.15 As shown in the embodiment of FIG. 1 , the volume of the cavity enclosed by the panels 91 to 94 will expand (assuming that there is no exit of the fluid 10 from the brake on the other side). The fluid pressure acting on the panels 91 to 94 forces the panels 91 to 94 to separate, which in turn straightens the two load-bearing front panels 91, 92, thereby driving the ends of the force-actuating element 8 (i.e., the connecting hinge 97) apart with a higher force. The rear panels 93, 94 act as a support for the fluid 10 pressed on the load-bearing front panels 91, 92. Therefore, the same applies: the rear panels 93, 94 are only subjected to tensile stresses, and the load-bearing front panels 91, 92 must undergo buckling.

[0243] The rear panels 93, 94 must be rigid enough to withstand buckling under the pressure of the fluid 10 acting on them. For the load-bearing front panels 91, 92, their rigidity must also withstand the buckling caused by the fluid pressure, but in addition, a more complex problem must be solved: these panels are under compressive loads, making them prone to buckling. Since they are connected by hinges, the load-bearing front panels 91, 92 will buckle in the first mode, which is also the buckling motion caused by the fluid pressure, so the buckling tendency of the panels is amplified by the fluid pressure acting on the panels. Therefore, the dimensions of the load-bearing front panels 91, 92 must be designed to withstand first-order buckling and additional buckling-inducing forces acting in the normal direction of their surfaces.

[0244] Once the force-actuated element 8 reaches its fully extended state, the front panels 91, 92 that carry the load form a straight line, so that the force is directly transmitted through the force-actuated element 8, and in an ideal situation, no pressure is required to hold the front panels 91, 92 that carry the load in place. In addition, in the fully extended state of the force-actuated element 8, the stopper is actually not loaded.

[0245] If a liquid rather than a gas is used as the fluid 10, the relatively very low compressibility of the liquid will lead to a further desirable effect. When the fluid 10 is maintained at a constant volume (i.e., no liquid 10 is added to or removed from the cavity), and assuming that the panel is rigid, the geometric constraints dictate that the force actuator element 8 maintains its shape without any input of a holding force or actuation force. Any change in the extension of the actuator will result in a change in its volume, which is not possible when the volume of the internal space 98 is maintained constant by blocking the fluid supply.

[0246] Unlike gases whose volume is directly related to pressure, even very high pressures cannot change the volume of liquids (assuming incompressibility). Therefore, assuming that panels 91 to 94 are rigid, the amount of force acting on the ends of the force-actuated element 8 will not change this geometry and extension state. Therefore, the force-actuated element 8 effectively behaves as a single solid with very high rigidity, rather than a thin shell tube. However, in reality, the actual actuator rigidity depends on the compressibility of the liquid, which is still very low compared to gases, but different liquids are different. Once a panel or hinge is loose (for example, the front panels 91, 92 of the load-bearing are buckled), the actuator will fail. Another advantage of using liquid instead of gas is that the liquid can support and change the buckling behavior of the front panels 91, 92 of the load-bearing, thereby directly affecting their size.

[0247] exist Fig.16 and Fig.17 In the embodiment of the invention, the rigid planar plates 91 to 94 have been replaced back with thin and flexible (e.g., compliant) sheets 83, 84, which mainly maintain the same mechanical behavior of the force-actuated element 8. The stopper (i.e., the second sheet 84 here) still acts as a support for the fluid 10 pressed against the load-bearing sheet (i.e., the first sheet 83 here). Therefore, the second sheet 84 is in tensile stress and the load-bearing first sheet is in compression. The thin-walled second sheet 84 easily adapts to the extended shape of any actuator while reliably maintaining the volume of the fluid 10; because the pressure of the fluid 10 acts on the concave area, the pure tensile stress generated does not cause a problem. For the load-bearing first sheet 83, the pressure of the fluid 10 acts on its convex side. Therefore, the pressure of the fluid 10 acts as a debuckling force (this is actually the actual operating principle of CSHA), which has an opposite effect on the buckling of the above-mentioned planar plates 91 to 94. If a thinner sheet is used to bend it into a curved shape, the load-bearing first sheet 83 itself is easier to buckle than the rigid planar plate / hinge version. However, in contrast to the buckling inducing effect of the planar panels 91 to 94, this effect is counteracted by the debuckling effect of the pressure acting on the convex side of the load-bearing first sheet 83. Fig.14 and Fig.15The use of two partially flexible thin walled sheets 83, 84 reduces complexity and weight compared to a flat panel version.

[0248] For the forces generated by the force actuating element 8, the actuating pressure is typically very low, which is a strong contrast and improvement over the hydraulic systems of the prior art which typically require very high fluid pressures; as previously mentioned, this actuating pressure reaches zero when the actuator is fully extended. The stop is therefore in tension (which disappears after full extension). The compressive load acting on the loaded first sheet 83 is particularly high when the stop is short (i.e. the enclosed volume of fluid 10 is low); it reaches a reasonable level after the stop length exceeds 105% of the length of the loaded sheet. This is at the expense of a higher enclosed fluid volume and actuator weight, unless this volume is used for other purposes.

[0249] The crescent-shaped hydraulic actuator is a force-actuated element 8 which is intended to be lightweight, to control volumes with a relatively low required working pressure (reaching zero in the fully extended state), and to be able to generate high internal forces at the required low strokes, which match the deformation of the highly loaded control element, deforming the wing internally with a closed, continuous skin. The length of this force-actuated element can be selected individually as required, and it generates a well-defined linear load at each end 81, 82, especially if these ends are formed in such a way that the forces are transmitted through cylindrical surfaces of the roller bearing type.

[0250] As part of the adjustment mechanism 7, optionally with an actuator structure 71, the force-actuated element 8 as shown in the drawings and described herein can be used to adjust (i.e., deform, in particular continuously deform) an aerodynamic element of an aircraft, such as a flight control element. This design can be used for all types of primary flight control functions, such as elevators, ailerons and rudders. In addition, the disclosed force-actuated element 8 can also be used to adjust (i.e., deform, in particular continuously deform) a fluid dynamic element of a ship, such as a control element. Therefore, all explanations above and below about aircraft also apply to ships.

[0251] Since these force-actuated elements 8 are embedded in the control element 6 and in particular inside the wing structure where fuel is usually stored, it is facilitated that the force-actuated elements 8 use fuel as a power transmission fluid, the fuel portion used in the CSHA then effectively becoming "flight control fuel". Since a final reserve fuel of at least 30 minutes of flight time is usually carried on board and is almost never used during the aircraft's service life (for example, this fuel weighs about one metric ton in a standard passenger aircraft of the A320), this amount of fuel can be used for flight control purposes, rather than just being carried in the fuel tank, thereby making efficient use of the resources on board the aircraft, taking into account its overall system architecture, and not making the entire system too heavy. Reaching this minimum fuel state and starting to consume this final reserve fuel will automatically require the declaration of an emergency flight situation, that is, a state that allows weakening flight control in addition to other certification margins (for example, when only a reduced number of CSHA force-actuated elements 8 can be controlled instead of all the CSHA force-actuated elements 8 as designed due to the reduction of the available "flight control fuel"). Due to the low required fluid pressure, a prior art fuel pump can be used for the fluid supply. This effectively eliminates the need for a separate hydraulic system to power the flight controls. In large transport category aircraft, the horizontal stabilizer usually contains a fuel tank for trim purposes. In all current transport category aircraft, the auxiliary power unit (APU) is located near the horizontal stabilizer. This means that in any case the fuel, as well as the fuel lines and pumps, are located at the tail of the aircraft to allow the fuel to also be used to deform the horizontal stabilizer for the purpose of controlling the pitch axis.

[0252] In addition to exploiting the fuel present in the wing, the hydraulic version of the CSHA has significant advantages over the pneumatic CSPA, which justifies the additional weight of an actuating liquid instead of an actuating gas. If a nearly incompressible liquid is used instead of a gas, such a closed volume has some of the properties of a solid. On the one hand, as long as no liquid is added or removed, this limits the movement of the supporting first sheet 83 to the sides, thereby firmly locking it in any desired position, including the neutral position, thus meeting the specific requirements of morphing flight control. With a gas under pressure, the same firm locking cannot be achieved, since the gas can be further compressed.

[0253] The force actuation element 8 in the form of a CSHA is embedded in Fig.18 In the stack of rectangular or square compartments 72 of the actuator structure 71 in the form of multiple tubes or sleeves shown, it is now possible not only to accurately and continuously control the outer curvature by appropriately adding or removing power transmission fluid 10 in each CSHA, but also to lock a certain shape into place by simply locking the fluid supply pipeline without the need for any additional actuation.

[0254] In order to achieve improved controllability of the deformation actuation in both directions, the system (ie the adjustment mechanism with the actuator structure 71) can be extended by introducing a second row of such force actuation elements 8 acting in the opposite direction, for example Fig.18 and Fig.19 As shown in the embodiments. Fig.18 and Fig.19 In both embodiments, the adjustment mechanism 7 comprises an actuator structure 71 having a plurality of rectangular compartments 72 arranged in rows and columns. In each of the compartments 72, a force actuation element 8 is provided. In each case, the adjustment mechanism 7 can be embedded in the continuous deformation element 6 of the aircraft to actively adjust (i.e., deform) the continuous deformation element 6 with a deformation type equivalent to camber deformation and thereby control the flight path of the aircraft.

[0255] exist Fig.18 In the embodiment, a CSHA type force actuation element 8 is used as an actuator, and Fig.19 In the embodiment of the embodiment of the present invention, a hydraulic piston is used as the force actuation element 8. In each case, the actuator structure 71 and a plurality of CSHA force actuation elements 8 or hydraulic pistons form an adjustment mechanism 7 for continuously deforming an aerodynamic element or a fluid dynamic element (e.g., a continuously deforming element 6 that realizes the combined function of a horizontal stabilizer and an elevator). In the undeformed state of the actuator structure 71, the force actuation elements 8 are all arranged parallel to each other. Another part of the adjustment mechanism can be formed by a pressure generating device (not shown in the drawings).

[0256] To deform an aerodynamic or fluid dynamic element, Fig.18 The CSHA is Fig.19 The CSHA is a standard hydraulic cylinder of the invention and acts in a very similar and powerful manner, so that the blocked fluid supply locks the specific extension state and deformation in place. The deformed actuator structure 71 acts in the same manner as the return spring together with the CSHA acting in the opposite direction. The main advantage of the CSHA over a standard hydraulic actuator is its lightweight design: it basically consists of two flexible sheets 83, 84 (fiber-reinforced composite materials are ideally suited for this purpose), which have a very large fluid action area (as opposed to the small area of ​​a standard piston), and a force amplification function. This results in a relatively very low actuating fluid pressure, which is equivalent to F=pxA, making it possible to use thin-walled, flexible and lightweight sheets 83, 84. In addition, the sheets 83, 84 produce linear loads rather than point loads, and can be trimmed to any length required. Compared with a standard piston hydraulic actuator, the ratio of the actuating force to weight is similar to that of a standard piston hydraulic actuator, while the ratio of the actuating force to fluid pressure is several orders of magnitude higher than that of a standard piston hydraulic actuator.

[0257] It should be noted that in all disclosed embodiments of the force-actuated element 8, the force-actuated direction F is transverse to the main pressure direction P, but Fig.19 Except for the conventional piston hydraulic actuator shown in Fig.19 In the conventional piston hydraulic actuator shown in , the force actuation direction is the same as the main pressurizing direction P.

[0258] In addition to actuator locking, another advantage of using a liquid power transmission fluid 10 rather than a gaseous power transmission fluid 10 becomes apparent when considering load carrying capacity. Fig. 20 and Fig.21 The fully extended and thus straightened condition would be an unstable state; although in this state the theoretically required pressure in the actuator is zero, at any pressure above zero the supported first sheet 83 may suddenly bend to the other side due to this positive gas pressure ( Fig. 20 Since this will increase the enclosed volume, the pressure will decrease again, making this new, more rounded state the stable state and rendering the force-actuated element 8 ineffective. In the case of actuation using an incompressible liquid as the power transmission fluid 10, the load-bearing first sheet 83 will not suddenly bend to the other side even in its straightened, fully extended state, because the liquid acts as a connecting element to the stoppered second sheet 84, effectively holding the load-bearing sheet in place ( Fig.21 on the right side of the ).

[0259] Furthermore, when comparing the buckling and load-bearing capacity in the fully extended state, it is apparent that the use of a liquid rather than a gas changes the buckling mode. With a gas, the load-bearing first sheet 83 is able to simply buckle in the first mode because the volume of the enclosed interior space 85 is able to expand, but with a liquid, the shape of the enclosed volume will be isovolumetric. Any buckling of the load-bearing first sheet 83 must account for this isovolumetric shape change, with the result that the buckling is at least in the second mode, significantly increasing the load-bearing capacity. In summary, a force-actuated element 8 filled with a liquid is generally more stable and more load-bearing than a force-actuated element filled with a gas, for reasons very similar to those of filling a metal tube with sand before bending it, with properties closer to those of a solid than to those of a tubular structure. Thus, the combination of the two sheets 83, 84 of a CSHA-type force-actuated element 8 with its liquid "core" causes it to behave like a sandwich structure.

[0260] However, the second mode / isovolumetric buckling also reduces the load-bearing capacity. This occurs even if the force-actuated element 8 is not fully extended by additional superposition of second-order buckling (i.e. the load-bearing first sheet 83 is intentionally buckled in the first mode). This occurs in particular in low extension states, where the required lateral force on the load-bearing first sheet 83 is high compared to the generated actuation force. Therefore, what may happen under high load conditions is that the CSHA first introduces the second mode buckling into the load-bearing first sheet 83 after the introduction of the liquid and thereafter switches back to the first mode buckling once the force-actuated element 8 has been extended somewhat (i.e. the lateral force required to generate the required actuation force is reduced).

[0261] This extension behavior is undesirable because it introduces a sharp curvature to the load-bearing first sheet 83, which can exacerbate material failure. Therefore, the load-bearing first sheet 83 should be adjusted to facilitate first mode buckling while resisting higher modes, essentially functioning as a hinge in its middle. This is functionally equivalent to Fig.14 and Fig.15 The front panel of the load bearing or the strap hinge 95 between the front panels of the load bearing are shown to be very similar. This can be achieved by adjusting the bending stiffness of the load bearing sheet along its width. It must be noted that the change in bending stiffness occurs gradually, with a limited and well-defined minimum value in the hinge area. In addition, fatigue issues also need to be addressed.

[0262] In the case of composite materials, this hinge function can be achieved by varying the twist along the width of the sheet, or by using a flexible matrix in the hinge area, which is essentially a flexible element. Another option is to geometrically vary the moment of inertia of the carrier sheet 83 along its width. For example, Fig.26 As shown, this can be accomplished by adding swept ribs or fins 90 to the carrier sheet 83; to avoid interfering with the fluid cavity, the swept ribs or fins 90 are preferably added to the concave side.

[0263] When the sheet 83 with this hinge function is bent, the resulting first mode buckling shape is closer to a V-shape than a circle. However, this does not hinder the basic function of the CSA, because the V-shaped load-bearing sheet will still adapt to the geometric sleeve size and provide the same function as a circular sheet.

[0264] For the expected very high internal actuation loads required for active transport aircraft wing deformation, these line loads must be transferred from the CSHA force actuation element 8 to the actuator structure 71 and ultimately to the wing structure. Fig. 22In the embodiments a) and c) of the invention, rotary bearings 86 (also called roller bearings) are introduced to allow smooth rotational movement of the two CSHA ends 81, 82 within the corners of the actuator structure 71 (which has a matching negative shape). These essentially cylindrical bearings allow high loads to be transmitted along the line. The first sheet 83 for load bearing and the second sheet 84 for stop can be introduced into such roller bearings 86 so that their ends point towards the center of rotation of the bearing. However, only the sliding contact part of the roller bearing needs to be cylindrical; the rest can be of different shapes, suitable for providing a firm contact with the two sheets 83, 84.

[0265] As for the specific arrangement of the first sheet 83 for carrying and the second sheet 84 for stopping, there are several options to discuss, such as Fig. 22 Detailed illustration. It is desirable to transfer the load from the load-bearing sheet 83 directly to the bearing without introducing any bending moments. Option b) allows simple production from composite materials, which are placed around a single core and reinforced at the edges to create a rigid bending connection 87 which is circular and can therefore double as a rotary bearing. However, this is not ideal because the force vector along the force actuation direction F is never in line with the load-bearing sheet 83 and therefore there are high bending moments near the bearing. Both options b) and c) use a specially adapted cylinder which acts as both a rotary bearing 86 and a container for the first sheet 83 for bearing and the second sheet 84 for stopping. Option c) requires a smaller volume or power transmission fluid 10 and both sheets 83, 84 can be attached to a common groove in the cylinder, making option c) the preferred option.

[0266] At both ends 81, 82 of the CSA force actuation element 8, appropriate seals need to be incorporated to a) allow removal or addition of fluid, b) contain the fluid 10 in the force actuation element 8, c) not restrict movement of the carrier sheet 83 and the second sheet 84 relative to each other, and d) maintain a constant volume of the fluid 10, particularly in the case of a liquid-operated CSHA, where the fuel volume defines the extended state of the actuator. To this end, such seals need to be flexible while avoiding "expansion" effects under pressure; this can be achieved, for example, using materials with a flexible matrix and fiber reinforcement. Removal and addition of fluid needs to be rapid because (particularly for CSHAs) the fluid flow directly controls the operating speed of the actuator. Therefore, the flight control actuation speed is directly proportional to the volumetric flow rate; since relatively low pressures are expected, any fluid line attachments should include a suitably large cross-sectional area.

[0267] Figure 23 to Figure 25 A practical embodiment of a force-actuated element 8 is shown which achieves effective sealing of the inner space 85. Fig.23In its simplest form, a simple fluid-filled bag 88 with one bag opening can be introduced between the sheets 83 and 84 to meet all these requirements. A disadvantage of this design may be that some of the volume that could otherwise be used for the power transmission fluid 10 is now used by the walls of the bag 88 itself, i.e. the minimum containment volume in the fully retracted state cannot be so close to zero. On the other hand, in order to avoid excessive peak forces in the actuator sheets 83, 84, a certain length difference between the actuator sheets 83, 84 (a length difference of 105% produces reasonable results), a minimum fluid volume will still be retained. The resulting minimum volume can then be filled by the bag walls, thus reducing the control fluid demand accordingly and actually playing a positive role in the disadvantage of having the bag extend along the entire length of the actuator.

[0268] like Fig.23 The bag version shown has one additional aspect that any fluid control valve will require to quickly control and meter the inflow and outflow of actuating fluid 10 through bag opening 89 . Fig.24 A version with two attachments (i.e., two bag openings 89) is shown, which, although making the bag more complex, enables the addition and removal of fluid to be allocated individually to each attachment, so that the flow is actually only one-way. This simplifies the design of the fluid control valves, since the fluid control valves are only of the one-way type, and although two such valves are required, one valve is only connected to the fluid feed line, and the other valve is connected to the return line, which can simply lead to the fuel tank in the case where fuel is used as the fluid. The single and dual attachment systems work essentially independently of any particular bag design.

[0269] exist Fig.25 In the embodiment shown, instead of using a bag 88 extending along the length of the force-actuated element 8, only an actual sealing cover is applied, the two sheets 83, 84 of the force-actuated element 8 taking on an additional function by directly containing the fluid 10. The sealing cover connection will bring additional challenges, such as bonding problems, as well as general sealing problems and fluid permeability problems in the case of composite materials used for the carrier sheet 83 and the stopper 84.

[0270] The invention is of course not limited to the embodiments described and shown in the drawings. Numerous modifications are possible.

[0271] Reference numerals list

[0272] 1 Aircraft

[0273] 2 Body

[0274] 21 Incision

[0275] 3 Wing

[0276] 31 Inboard flaps

[0277] 32 Inboard aileron

[0278] 33 Outer flaps

[0279] 34 Outer ailerons

[0280] 4 Vertical stabilizer

[0281] 41 Rudder

[0282] 5 Horizontal stabilizer

[0283] 51 Elevator

[0284] 52 Front Edge

[0285] 53 trailing edge

[0286] 54 Actuator system based on screw crane

[0287] 6 Continuous deformation element

[0288] 7 Adjustment mechanism

[0289] 71 Actuation structure

[0290] 72 compartments

[0291] 8 Force-actuated elements

[0292] 81 First end

[0293] 82 Second end

[0294] 83 First Sheet

[0295] 84 Second sheet

[0296] 85 Interior Space

[0297] 86 Swivel bearing

[0298] 87 Rigid bending connection

[0299] 88 bags

[0300] 89 bags opening

[0301] 90 fins / ribs

[0302] 91 First front panel

[0303] 92 Second front panel

[0304] 93 First rear panel

[0305] 94 Second rear panel

[0306] 95 Front strap hinge

[0307] 96 rear strap hinge

[0308] 97 Connecting hinge

[0309] 98 Interior Space

[0310] 10 Fluid

[0311] 11 Airflow Separation

[0312] A Airflow direction

[0313] F Force actuation direction

[0314] P Main pressure direction

[0315] I Contour

Claims

1. A control unit (6, 7) for controlling the pitch of an aircraft (1) during flight, the control unit (6, 7) comprising: at least one adjustable control element (6) allowing the pitch of the aircraft (1) to be controlled during flight; as well as an adjustment mechanism (7) for actively adjusting the at least one control element (6) to control the pitch of the aircraft (1); It is characterized in that The adjustment mechanism (7) is suitable for controlling the pitch of the aircraft (1) by means of a continuous deformation of the at least one control element (6).

2. The control unit (6, 7) according to claim 1, wherein: The at least one control element (6) is adapted to additionally perform a trimming function, and wherein the adjustment mechanism (7) is preferably adapted to control the trimming of the at least one control element (6) by means of an additional deformation of the at least one control element (6).

3. A control unit (6, 7) according to claim 1 or 2, wherein: The control element (6) has a leading edge (52) and a trailing edge (53) relative to the usual airflow direction (A) during flight, and wherein the adjustment mechanism (7) is suitable for deforming the leading edge (52) and the trailing edge (53) of the control element (6) in the same direction.

4. The control unit (6, 7) according to one of the preceding claims, wherein The adjustment mechanism (7) is adapted to deform the at least one control element (6) such that a deformation up to a maximum of the at least one control element (6) during flight results in airflow loss around one or more regions of the at least one control element (6).

5. The control unit (6, 7) according to one of the preceding claims, wherein The adjustment mechanism (7) is adapted to deform the at least one control element (6) so that as the deformation of the at least one control element (6) increases, the lift force generated by the at least one control element (6) changes nonlinearly with the deformation.

6. The control unit (6, 7) according to one of the preceding claims, wherein The at least one control element (6) is suitable for additionally performing the function of a horizontal stabilizer (6) of the aircraft.

7. The control unit (6, 7) according to one of the preceding claims, wherein The adjustment mechanism (7) is suitable for gradually deforming the at least one control element (6) in two directions: a first direction from the leading edge (52) toward the trailing edge (53) and a direction from the fuselage (2) of the aircraft (1) toward the lateral outer edge of the at least one control element (6).

8. The control unit (6, 7) according to one of the preceding claims, wherein The adjustment mechanism has at least one force-actuated element (8), and the at least one force-actuated element (8) comprises: A first end portion (81); a second end portion (82); a first sheet material (83) connecting the first end portion (81) and the second end portion (82); and a second sheet (84) connecting the first end (81) and the second end (82); wherein the first sheet (83) and the second sheet (84) together enclose an inner space (85) of the force-actuated element (8), the inner space (85) being adapted to be filled with a power transmission fluid (10), such that a change in the amount and / or pressure of the power transmission fluid (10) in the inner space (85) causes the first end (81) and the second end (82) to move relative to each other along a force-actuated direction (F) to adjust the at least one control element (6), And wherein, the movement of the first end (81) and the second end (82) relative to each other along the force actuation direction (F) is affected by the pressure of the power transmission fluid (10) acting on the first sheet (83) and / or the second sheet (84) along the main pressurization direction (P), and the main pressurization direction (P) is transverse to the force actuation direction (F).

9. The control unit (6, 7) according to claim 8, wherein: The force actuation element (8) is located within the at least one control element (6).

10. A control unit (6, 7) according to claim 8 or 9, wherein: The adjustment mechanism (7) comprises a plurality of such force-actuated elements (8) arranged in rows and / or columns, and wherein the adjustment mechanism (7) preferably comprises an actuator structure (71) having compartments (72), each compartment being used to accommodate one force-actuated element (8).

11. The control unit (6, 7) according to one of claims 8 to 10, wherein: The adjustment mechanism (7) comprises at least a first force-actuated element (8) and a second force-actuated element (8), and wherein the adjustment mechanism (7) is suitable for increasing the amount and / or pressure of the power transmission fluid (10) in the internal space (85) of the first force-actuated element (8), and simultaneously reducing the amount and / or pressure of the power transmission fluid (10) in the internal space (85) of the second force-actuated element (8), so as to adjust the control element (6) by means of both the first force-actuated element (8) and the second force-actuated element (8).

12. An aircraft (1) comprising a control unit (6, 7) according to one of the preceding claims, the control unit (6, 7) having at least one adjustable control element (6) and an adjustment mechanism (7), the adjustment mechanism (7) being used to actively adjust the at least one control element (6) to control the pitch of the aircraft (1).

13. The aircraft according to claim 12, wherein: The control element (6) is firmly connected to the fuselage (2) of the aircraft (1).

14. A method for controlling the pitch of an aircraft (1), in particular a method for controlling the pitch of an aircraft (1) according to claim 12 or 13, said method comprising at least the step of controlling the pitch of the aircraft (1) by actively adjusting at least one control element (6) of the aircraft (1), It is characterized in that The pitch is controlled by continuously deforming the control element (6).

15. A force-actuated element (8), in particular a force-actuated element (8) for adjusting an aerodynamic element or a fluid dynamic element (3, 51, 32, 34, 4, 6) of an aircraft (1) or a ship, the force-actuated element (8) comprising: A first end portion (81); a second end portion (82); a first sheet material (83) connecting the first end portion (81) and the second end portion (82); and a second sheet (84) connecting the first end (81) and the second end (82); wherein the first sheet (83) and the second sheet (84) together enclose an inner space (85) of the force-actuated element (8), the inner space (85) being adapted to be filled with a power transmission fluid (10), such that the amount and / or pressure and / or a change in pressure of the power transmission fluid (10) in the inner space (85) causes the first end (81) and the second end (82) to move relative to each other along a force-actuated direction (F), It is characterized in that The movement of the first end (81) and the second end (82) relative to each other along the force actuation direction (F) is affected by the pressure of the power transmission fluid (10) acting on the first sheet (83) and / or the second sheet (84) along the main pressurization direction (P), and the main pressurization direction is transverse to the force actuation direction (F).

16. The force-actuated element (8) according to claim 15, wherein: The power transmission fluid (10) is a liquid.

17. The force-actuated element (8) according to claim 15 or 16, wherein: The first sheet (83) and the second sheet (84) are both continuous sheets, and the first sheet (83) and the second sheet (84) together substantially completely enclose the interior space (85).

18. The force-actuated element (8) according to any one of claims 14 to 16, further comprising a first rotary bearing (86) and / or a second rotary bearing (86), the first rotary bearing (86) connecting the first sheet (83) and the second sheet (84) at the first end (81), and the second rotary bearing (86) connecting the first sheet (83) and the second sheet (84) at the second end (82).

19. An aircraft (1) comprising an aerodynamic element (3, 51, 32, 34, 4, 5, 6), in particular a wing (3), a combined stabilizer and elevator (6), an elevator (51), an aileron (32, 34) or a stabilizer (4, 5); and At least one force-actuated element (8) according to one of claims 15 to 18 for adjusting the aerodynamic element (3, 51, 32, 34, 4, 5, 6).

20. An aircraft (1) according to claim 19, wherein: At least one force-actuating element (8) is filled with fuel of the aircraft (1) to use the fuel as a power transmission fluid (10).

21. A method for adjusting an aerodynamic or hydrodynamic element (3, 51, 32, 34, 4, 5, 6) of an aircraft (1) or a watercraft, in particular by means of a force-actuated element (8) according to one of claims 15 to 18, the method comprising at least the following steps: The amount and / or pressure of the fluid transport fluid (10) in the internal space (85) enclosed by the first sheet (83) and the second sheet (84) is changed to generate a pressure acting on the first sheet (83) and / or the second sheet (84) along a main pressurization direction (P), and thereby cause the first end (81) and the second end (82) connected to each other by the first sheet (83) and the second sheet (84) to move relative to each other along a force actuation direction (F) transverse to the main pressurization direction (P) to adjust the aerodynamic element or fluid dynamic element (3, 51, 32, 34, 4, 5, 6).

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