Office Action 1: Non-Final Rejection (2024-11-13)
Rejection types: 102 - Lack of Novelty, 103 - Obviousness, 112 - Written Description/Enablement

CLAIMS:
CLAIMS
1\. An aircraft wing comprising a flap system that is mounted at a leading or trailing edge of the wing and comprises at least two movable flap devices spaced apart in a spanwise direction Y from each other, wherein each movable flap device is supported by at least one support station and is driven about a hinge axis by at least one drive system, wherein the flap system comprises a synchronizing torque element connected to the movable flap devices, wherein the synchronizing torque element is configured to transmit a torque between the movable flap devices to actuate a rotational movement of the movable flap devices, and wherein the synchronizing torque element is positioned between the movable flap devices and around an adjacent system or structural component that is also positioned between the movable flap devices.
2\. The aircraft wing according to claim 1, wherein the synchronizing torque element comprises an axis of rotation that is parallel and eccentric to a hinge line of the movable flap devices.
3\. The aircraft wing according to claim 1, wherein the synchronizing torque element is connected to at least one movable flap device using a universal joint.
4\. The aircraft wing according to claim 1, wherein the synchronizing torque element is supported by a plurality of side stay struts that are each connected to a corresponding flap master support.
5\. The aircraft wing according to claim 1, wherein the system component is an aircraft pylon containing a H2 storage tank.
6\. An aircraft comprising a wing according to claim 1.

OFFICE ACTION:
DETAILED ACTION
Notice of Pre-AIA  or AIA  Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Drawings
The drawings are objected to under 37 CFR 1.83(a).  The drawings must show every feature of the invention specified in the claims.  Therefore, the support station, drive system, synchronizing torque element, adjacent system, structural component, universal joint, side stay struts, flap master support, aircraft pylon, and H2 storage tanks, must be shown or the feature(s) canceled from the claim(s).  No new matter should be entered.
Corrected drawing sheets in compliance with 37 CFR 1.121(d) are required in reply to the Office action to avoid abandonment of the application. Any amended replacement drawing sheet should include all of the figures appearing on the immediate prior version of the sheet, even if only one figure is being amended. The figure or figure number of an amended drawing should not be labeled as “amended.” If a drawing figure is to be canceled, the appropriate figure must be removed from the replacement sheet, and where necessary, the remaining figures must be renumbered and appropriate changes made to the brief description of the several views of the drawings for consistency. Additional replacement sheets may be necessary to show the renumbering of the remaining figures. Each drawing sheet submitted after the filing date of an application must be labeled in the top margin as either “Replacement Sheet” or “New Sheet” pursuant to 37 CFR 1.121(d). If the changes are not accepted by the examiner, the applicant will be notified and informed of any required corrective action in the next Office action. The objection to the drawings will not be held in abeyance.
The drawings are objected to as failing to comply with 37 CFR 1.84(p)(5) because they include the following reference character(s) not mentioned in the description: none of the reference numbers are in the specification and there is not a detailed description of the figures in the specification.  It is unclear if the Applicant has support in the disclosure for the claimed invention.  Corrected drawing sheets in compliance with 37 CFR 1.121(d), or amendment to the specification to add the reference character(s) in the description in compliance with 37 CFR 1.121(b) are required in reply to the Office action to avoid abandonment of the application. Any amended replacement drawing sheet should include all of the figures appearing on the immediate prior version of the sheet, even if only one figure is being amended. Each drawing sheet submitted after the filing date of an application must be labeled in the top margin as either “Replacement Sheet” or “New Sheet” pursuant to 37 CFR 1.121(d). If the changes are not accepted by the examiner, the applicant will be notified and informed of any required corrective action in the next Office action. The objection to the drawings will not be held in abeyance.
Specification
The disclosure is objected to because of the following informalities:  none of the reference numbers from the figures are in the specification making it unclear what the reference numbers are referring to.  Further, there is not a description of the figures and it is unclear if the Applicant has support for the claimed subject matter.
Appropriate correction is required.
Claim Rejections - 35 USC § 112
The following is a quotation of the first paragraph of 35 U.S.C. 112(a):
(a) IN GENERAL.—The specification shall contain a written description of the invention, and of the manner and process of making and using it, in such full, clear, concise, and exact terms as to enable any person skilled in the art to which it pertains, or with which it is most nearly connected, to make and use the same, and shall set forth the best mode contemplated by the inventor or joint inventor of carrying out the invention.
The following is a quotation of the first paragraph of pre-AIA  35 U.S.C. 112:
The specification shall contain a written description of the invention, and of the manner and process of making and using it, in such full, clear, concise, and exact terms as to enable any person skilled in the art to which it pertains, or with which it is most nearly connected, to make and use the same, and shall set forth the best mode contemplated by the inventor of carrying out his invention.
The following is a quotation of 35 U.S.C. 112(b):
(b)  CONCLUSION.—The specification shall conclude with one or more claims particularly pointing out and distinctly claiming the subject matter which the inventor or a joint inventor regards as the invention.
The following is a quotation of 35 U.S.C. 112 (pre-AIA ), second paragraph:
The specification shall conclude with one or more claims particularly pointing out and distinctly claiming the subject matter which the applicant regards as his invention.
Claims 1-6 are rejected under 35 U.S.C. 112(a) or 35 U.S.C. 112 (pre-AIA ), first paragraph, as failing to comply with the written description requirement. The claim(s) contains subject matter which was not described in the specification in such a way as to reasonably convey to one skilled in the relevant art that the inventor or a joint inventor, or for applications subject to pre-AIA  35 U.S.C. 112, the inventor(s), at the time the application was filed, had possession of the claimed invention.
The specification seems to merely mention the claim language but does not go into detail about of how the invention is made.  It appears that the figures may show some of the claimed features but since the figures are not described in detail and none of the reference numbers are in the specification, it is unclear how the invention is made or used or even if there is support for the invention.  A detailed description of the figures and reference numbers may resolve the above rejection as long as no new matter is added.
Claim 5 is rejected under 35 U.S.C. 112(b) or 35 U.S.C. 112 (pre-AIA ), second paragraph, as being indefinite for failing to particularly point out and distinctly claim the subject matter which the inventor or a joint inventor (or for applications subject to pre-AIA  35 U.S.C. 112, the applicant), regards as the invention.
Claim 5 recites the limitation the system component in claim 5, line 1.  There is insufficient antecedent basis for this limitation in the claim.  This appears to be a typo and should be the structural component of claim 1.  For compact prosecution, the Examiner is interpreting “the system component” in claim 5, line 1, as -- the structural component --.
Claim Rejections - 35 USC § 102
In the event the determination of the status of the application as subject to AIA  35 U.S.C. 102 and 103 (or as subject to pre-AIA  35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis (i.e., changing from AIA  to pre-AIA ) for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of the appropriate paragraphs of 35 U.S.C. 102 that form the basis for the rejections under this section made in this Office action:
A person shall be entitled to a patent unless –
(a)(1) the claimed invention was patented, described in a printed publication, or in public use, on sale, or otherwise available to the public before the effective filing date of the claimed invention.
Claim(s) 1, 2, & 6 is/are rejected under 35 U.S.C. 102(a)(1) as being anticipated by Bellego et al (US Patent No. 4,789,119).
Regarding claims 1 & 6
Bellego teaches an aircraft (See figure 1, ref # 1) comprising a wing, (See figure 1, ref # 2D & 2G) an aircraft wing (See figure 1, ref # 2D & 2G) comprising a flap system (See figure 1, ref # 4D, 4G, 5D, & 5G) that is mounted at a leading or trailing edge of the wing (See figure 1, ref # 2D & 2G) and comprises at least two movable flap devices (See figure 1, ref # 4D, 4G, 5D, & 5G) spaced apart in a spanwise direction Y from each other, wherein each movable flap device (See figure 1, ref # 4D, 4G, 5D, & 5G) is supported by at least one support station (See figures 2 & 3, ref # 9.4, 7.4, & 7.5) and is driven about a hinge axis (See figures 2 & 3, ref # 8.4) by at least one drive system, (See figures 2 & 3, ref # 11.4 & 11.5) wherein the flap system (See figure 1, ref # 4D, 4G, 5D, & 5G) comprises a synchronizing torque element (See figures 2 & 3, ref # 16) connected to the movable flap devices, (See figure 1, ref # 4D, 4G, 5D, & 5G) wherein the synchronizing torque element (See figures 2 & 3, ref # 16) is configured to transmit a torque between the movable flap devices (See figure 1, ref # 4D, 4G, 5D, & 5G) to actuate a rotational movement of the movable flap devices, (See figure 1, ref # 4D, 4G, 5D, & 5G) and wherein the synchronizing torque element (See figures 2 & 3, ref # 16) is positioned between the movable flap devices (See figure 1, ref # 4D, 4G, 5D, & 5G) and around an adjacent system or structural component (See figure 1, ref # fuselage) that is also positioned between the movable flap devices.  (See figure 1, ref # 4D, 4G, 5D, & 5G)
Regarding claim 2
Bellego teaches wherein the synchronizing torque element (See figures 2 & 3, ref # 16) comprises an axis of rotation that is parallel and eccentric to a hinge line of the movable flap devices.  (See figure 1, ref # 4D, 4G, 5D, & 5G)
Claim(s) 1-3 & 6 is/are rejected under 35 U.S.C. 102(a)(1) as being anticipated by Boehringer et al. (US Patent No. 4,441,675).
Regarding claims 1 & 6
Boehringer teaches an aircraft (See column 1, lines 5-10) comprising a wing, (See column 1, lines 5-10 & 50-54) an aircraft wing (See column 1, lines 5-10 & 50-54) comprising a flap system (See figure 3, ref # 8 & 9) that is mounted at a leading or trailing edge of the wing (See column 3, lines 18-21) and comprises at least two movable flap devices (See figure 3, ref # 8 & 9) spaced apart in a spanwise direction Y from each other, wherein each movable flap device (See figure 3, ref # 8 & 9) is supported by at least one support station (See figure 3, ref # 109) and is driven about a hinge axis by at least one drive system, (See figure 3, ref # 26) wherein the flap system (See figure 3, ref # 8 & 9) comprises a synchronizing torque element (See figure 3, ref # 107) connected to the movable flap devices, (See figure 3, ref # 8 & 9) wherein the synchronizing torque element (See figure 3, ref # 107) is configured to transmit a torque between the movable flap devices (See figure 3, ref # 8 & 9) to actuate a rotational movement of the movable flap devices, (See figure 3, ref # 8 & 9) and wherein the synchronizing torque element (See figure 3, ref # 107) is positioned between the movable flap devices (See figure 3, ref # 8 & 9) and around an adjacent system or structural component (See figures 3 & 4) that is also positioned between the movable flap devices.  (See figure 3, ref # 8 & 9)
Regarding claim 2
Boehringer teaches wherein the synchronizing torque element (See figure 3, ref # 107) comprises an axis of rotation that is parallel and eccentric to a hinge line of the movable flap devices.  (See figure 3, ref # 8 & 9)
Regarding claim 3
Boehringer teaches wherein the synchronizing torque element (See figure 3, ref # 107) is connected to at least one movable flap device (See figure 3, ref # 8 & 9) using a universal joint.  (See column 5, lines 50-68 & column 6, lines 1-2)
Claim Rejections - 35 USC § 103
The following is a quotation of 35 U.S.C. 103 which forms the basis for all obviousness rejections set forth in this Office action:
A patent for a claimed invention may not be obtained, notwithstanding that the claimed invention is not identically disclosed as set forth in section 102, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
The factual inquiries for establishing a background for determining obviousness under 35 U.S.C. 103 are summarized as follows:
1. Determining the scope and contents of the prior art.
2. Ascertaining the differences between the prior art and the claims at issue.
3. Resolving the level of ordinary skill in the pertinent art.
4. Considering objective evidence present in the application indicating obviousness or nonobviousness.
Claim(s) 3 is/are rejected under 35 U.S.C. 103 as being unpatentable over Bellego et al (US Patent No. 4,789,119) as applied to claim 1 above, and further in view of Boehringer et al. (US Patent No. 4,441,675).
Regarding claim 3
Bellego teaches wherein the synchronizing torque element (See figures 2 & 3, ref # 16) is connected to at least one movable flap device.  (See figure 1, ref # 4D, 4G, 5D, & 5G)
Bellego is silent about a universal joint.
However, Boehringer teaches wherein the synchronizing torque element (See figure 3, ref # 107) is connected to at least one movable flap device (See figure 3, ref # 8 & 9) using a universal joint.  (See column 5, lines 50-68 & column 6, lines 1-2)
Therefore it would have been obvious to one of ordinary skill in the art at the time the invention was filed to have a synchronizing torque element is connected to at least one movable flap device using a universal joint as taught by Boehringer in the aircraft of Bellego, since universal joints are well-known joints.
Claim(s) 5 is/are rejected under 35 U.S.C. 103 as being unpatentable over Bellego et al (US Patent No. 4,789,119) as applied to claim 1 above, and further in view of Allain et al. (Pub No. US 2022/0371743 A1).
Regarding claim 5
Bellego teaches wherein the system component (See figure 1, ref # fuselage or pylon) is an aircraft pylon.  (See figure 1, ref # pylon)
Bellego is silent about a H2 storage tank.
However, Allain teaches wherein the system component is an aircraft pylon (See paragraph 0030, 0033 & figure 1, ref # 150) containing a H2 storage tank.  (See paragraphs 0036, 0045 & figure 2, ref # 160/184)
Therefore it would have been obvious to one of ordinary skill in the art at the time the invention was filed to have an aircraft pylon containing a H2 storage tank as taught by Allain in the aircraft of Bellego, so as to provide power to the fuel cell in order to power the electric motors.  (See paragraphs 0004-0005)
Conclusion
The prior art made of record and not relied upon is considered pertinent to applicant's disclosure.  The reference Barba (US Patent No. 6,224,022 B1) discloses an aircraft comprising a wing; an aircraft wing comprising a flap system that is mounted at a leading or trailing edge of the wing and comprises at least two movable flap devices spaced apart in a spanwise direction Y from each other, wherein each movable flap device is supported by at least one support station and is driven about a hinge axis by at least one drive system, wherein the flap system comprises a synchronizing torque element connected to the movable flap devices, wherein the synchronizing torque element is configured to transmit a torque between the movable flap devices to actuate a rotational movement of the movable flap devices, and wherein the synchronizing torque element is positioned between the movable flap devices and around an adjacent system or structural component that is also positioned between the movable flap devices; and wherein the synchronizing torque element comprises an axis of rotation that is parallel and eccentric to a hinge line of the movable flap devices.  The reference Vincekovic et al. (Pub No. US 2024/0308681 A1) discloses an aircraft comprising a wing; an aircraft wing comprising a flap system; wherein the system component is an aircraft pylon containing a H2 storage tank.  The reference Miller et al. (Pub No. US 2019/0023378 A1) discloses an aircraft comprising a wing; an aircraft wing comprising a flap system that is mounted at a leading or trailing edge of the wing and comprises at least two movable flap devices spaced apart in a spanwise direction Y from each other, wherein each movable flap device is supported by at least one support station and is driven about a hinge axis by at least one drive system, wherein the flap system comprises a synchronizing torque element connected to the movable flap devices, wherein the synchronizing torque element is configured to transmit a torque between the movable flap devices to actuate a rotational movement of the movable flap devices, and wherein the synchronizing torque element is positioned between the movable flap devices and around an adjacent system or structural component that is also positioned between the movable flap devices.


NEXT OFFICE ACTION: DETAILED ACTION
Notice of Pre-AIA  or AIA  Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Allowable Subject Matter
Claims 1-3 & 5-6 are allowed.
The following is an examiner’s statement of reasons for allowance:
Regarding claim 1,
The prior art does not disclose or suggest the claimed “wherein the synchronizing torque element is supported by a plurality of side stay struts that are each connected to a corresponding flap master support” in combination with the remaining claim elements as set forth in claim 1.
Any comments considered necessary by applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee.  Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”
Conclusion

CITED REFERENCE SUMMARIES: Ref US20220371743A1: Claims:
A tank system for an aircraft, the tank system comprising: a tank comprising an inner enclosure comprising a closed inner skin configured to store dihydrogen and an outer enclosure comprising an outer skin that surrounds the inner enclosure and formed as a cylinder of revolution about a longitudinal axis, a chassis comprising a frame at one end of the tank, and attachment means which attach the outer skin of the outer enclosure to the frame and which comprise three attachment means of a first type distributed angularly about the outer enclosure at 3 o'clock, 12 o'clock and 9 o'clock as viewed along the longitudinal axis and a plurality of attachment means of a second type arranged between the attachment means of the first type, wherein each attachment means of the second type comprises a two link point connecting rod, one of the two link points providing attachment to the frame and another of the two link points providing attachment to the outer skin of the outer enclosure, and wherein the two link points of a given connecting rod are aligned parallel to the longitudinal axis.
a tank comprising an inner enclosure comprising a closed inner skin configured to store dihydrogen and an outer enclosure comprising an outer skin that surrounds the inner enclosure and formed as a cylinder of revolution about a longitudinal axis,
a chassis comprising a frame at one end of the tank, and
attachment means which attach the outer skin of the outer enclosure to the frame and which comprise three attachment means of a first type distributed angularly about the outer enclosure at 3 o'clock, 12 o'clock and 9 o'clock as viewed along the longitudinal axis and a plurality of attachment means of a second type arranged between the attachment means of the first type, wherein each attachment means of the second type comprises a two link point connecting rod, one of the two link points providing attachment to the frame and another of the two link points providing attachment to the outer skin of the outer enclosure, and wherein the two link points of a given connecting rod are aligned parallel to the longitudinal axis.
The tank system according to claim 1 , wherein the attachment means of the first type at 12 o'clock comprises a three link point connecting rod, two of the three link points attaching to the frame and a third link point of the three link points attaching to the outer skin of the outer enclosure.
The tank system according to claim 2 , wherein each link point between the connecting rod and the frame takes is formed as a male clevis of the connecting rod, a female clevis which is secured to the frame and into which the male clevis fits, and a pin passing through the female clevis and the male clevis, and wherein the axis of the pin is aligned in a direction that is generally radial with respect to the longitudinal axis.
wherein each link point between the connecting rod and the frame takes is formed as a male clevis of the connecting rod, a female clevis which is secured to the frame and into which the male clevis fits, and a pin passing through the female clevis and the male clevis, and
wherein the axis of the pin is aligned in a direction that is generally radial with respect to the longitudinal axis.
The tank system according to claim 3 , wherein the third link point between the connecting rod and the outer skin of the outer enclosure is formed as a ball link point embodied by a peg of the connecting rod, having an axis parallel to the longitudinal axis and which fits into a ball that is secured to the outer skin.
The tank system according to claim 2 , wherein the attachment means of the first type at 3 o'clock and at 9 o'clock are identical to the attachment means of the first type at 12 o'clock.
The tank system according to claim 2 , wherein each attachment means of the first type at 3 o'clock and at 9 o'clock comprises two pairs of connecting rods, wherein each connecting rod is a two link point connecting rod, one of the two link points providing attachment to the frame and another of the two link points providing attachment to the outer skin of the outer enclosure, wherein the two pairs of connecting rods of a given attachment means of the first type are arranged on either side of a median plane passing through the longitudinal axis, wherein the two pairs come closer together at their link points with the outer skin, and wherein the two pairs are further apart at their link points with the frame.
wherein each attachment means of the first type at 3 o'clock and at 9 o'clock comprises two pairs of connecting rods,
wherein each connecting rod is a two link point connecting rod, one of the two link points providing attachment to the frame and another of the two link points providing attachment to the outer skin of the outer enclosure,
wherein the two pairs of connecting rods of a given attachment means of the first type are arranged on either side of a median plane passing through the longitudinal axis,
wherein the two pairs come closer together at their link points with the outer skin, and
wherein the two pairs are further apart at their link points with the frame.
The tank system according to claim 1 , wherein each link point of each connecting rod is formed as a female clevis of the connecting rod, a male clevis which is secured respectively to the frame and to the outer skin of the outer enclosure and onto which the female clevis fits, and a pin passing through the female clevis and the male clevis, and wherein an axis of the pin is aligned in a direction that is generally tangential with respect to the longitudinal axis.
wherein each link point of each connecting rod is formed as a female clevis of the connecting rod, a male clevis which is secured respectively to the frame and to the outer skin of the outer enclosure and onto which the female clevis fits, and a pin passing through the female clevis and the male clevis, and
wherein an axis of the pin is aligned in a direction that is generally tangential with respect to the longitudinal axis.
The tank system according to claim 7 , wherein the outer skin comprises stiffeners that extend parallel to the longitudinal axis, and wherein each male clevis which is secured to the outer skin of the outer enclosure, and onto which a female clevis of the connecting rod fits, is one of the stiffeners.
wherein the outer skin comprises stiffeners that extend parallel to the longitudinal axis, and
wherein each male clevis which is secured to the outer skin of the outer enclosure, and onto which a female clevis of the connecting rod fits, is one of the stiffeners.
An aircraft having at least one tank system according to claim 1 .
Ref US20240308681A1: Claims:
An aircraft propulsion module comprising a hydrogen storage system, at least one electrochemical converter connected to the hydrogen storage system, wherein the at least one electrochemical converter is adapted to convert hydrogen supplied from the hydrogen storage system into electric energy, and at least one electric motor electrically connected to the at least one electrochemical converter, wherein the electric motor is adapted to generate thrust; wherein the propulsion module comprises at least one separation means adapted to separate at least one component of the propulsion module from the propulsion module.
a hydrogen storage system,
at least one electrochemical converter connected to the hydrogen storage system, wherein the at least one electrochemical converter is adapted to convert hydrogen supplied from the hydrogen storage system into electric energy, and
at least one electric motor electrically connected to the at least one electrochemical converter, wherein the electric motor is adapted to generate thrust;
the propulsion module comprises at least one separation means adapted to separate at least one component of the propulsion module from the propulsion module.
The aircraft propulsion module according to claim 1 , wherein at least one separable component is removable from the propulsion module after separation.
The aircraft propulsion module according to claim 1 , wherein at least one separable component is tethered to the propulsion module after removal.
The aircraft propulsion module according to claim 3 , wherein the separable component is tethered to one end of a tether, the other end of the tether being connected to a spool of the aircraft propulsion module.
The aircraft propulsion module according to claim 4 , wherein the end of the tether connected to the spool is remotely releasable from the spool.
The aircraft propulsion module according to claim 1 , wherein the at least one separable component comprises or is the at least one electrochemical converter, at least one electric energy storage unit, and/or the at least one electric motor.
The aircraft propulsion module according to claim 1 , wherein at least one separable component is in slidable connection with a support and held in position at the support by at least one axle leading through respective holes in the component and the support, the at least one axle is held in place by a fixation element attached the free end section of the axle, and at least one separation means is adapted to remove the at least one axle out of the respective holes against resistance from the respective fixation element.
at least one separable component is in slidable connection with a support and held in position at the support by at least one axle leading through respective holes in the component and the support,
the at least one axle is held in place by a fixation element attached the free end section of the axle, and
at least one separation means is adapted to remove the at least one axle out of the respective holes against resistance from the respective fixation element.
The aircraft propulsion module according to claim 7 , wherein the slidable connection is such that, when the at least one axle is removed, the separated component slides down the support pulled by its weight, and then falls out of a fairing of the propulsion module.
The aircraft propulsion module according to claim 7 , wherein at least one separation means comprises an electrically driven actuator and/or a mechanical linkage which, when operated, exerts a force on the at least one axle along its longitudinal axis.
The aircraft propulsion module according to claim 7 , wherein the at least one separable component comprises at least one elongated hole and/or the support comprises at least one elongated hole, through which, when the component is attached to the support, at least one connection line is laid, and wherein at least one of the holes comprises a cutting edge, in particular inclined cutting edge.
the at least one separable component comprises at least one elongated hole and/or the support comprises at least one elongated hole, through which, when the component is attached to the support, at least one connection line is laid, and wherein
at least one of the holes comprises a cutting edge, in particular inclined cutting edge.
The aircraft propulsion module according to claim 6 , wherein the at least one electrochemical converter, the at least one electric energy storage unit, and the at least one electric motor are located in a first part, in particular front part, of the propulsion module, wherein the hydrogen storage system is located in a second part, in particular rear part, of the propulsion module, and wherein the first part is separable from the second part by the at least one separation means.
The aircraft propulsion module according to claim 11 , wherein the first part and/or the second part comprise at least one elongated hole, through which, when the first part is attached to the second part, at least one connection line is laid, and wherein at least one of the holes comprises a cutting edge, in particular inclined cutting edge, designed such that, when the first part is separated from the second part by the at least one separation means, the cutting edge cuts the at least one connection line.
the first part and/or the second part comprise at least one elongated hole, through which, when the first part is attached to the second part, at least one connection line is laid, and wherein
at least one of the holes comprises a cutting edge, in particular inclined cutting edge, designed such that, when the first part is separated from the second part by the at least one separation means, the cutting edge cuts the at least one connection line.
An aircraft, comprising at least one aircraft propulsion module according to claim 1 .
A method for operating a propulsion module according to claim 1 , in which, during operation of the propulsion module, at least one separation means is actuated, by which actuation at least one component of the propulsion module is separated from the remaining propulsion module and then falls from the remaining propulsion module.
Ref US20190023378A1: Claims:
A control surface operating system for controlling aerodynamic control surfaces of an aircraft, the system comprising: a drive system operably connected to a drive shaft, the drive system including a continuously variable transmission; at least one actuator operably connected to the drive shaft and arranged to convert rotational movement of the drive shaft to move at least a portion of the aerodynamic control surface; and at least one control surface operably connected to the at least one actuator, wherein the at least one control surface is adjusted by the at least one actuator.
a drive system operably connected to a drive shaft, the drive system including a continuously variable transmission;
at least one actuator operably connected to the drive shaft and arranged to convert rotational movement of the drive shaft to move at least a portion of the aerodynamic control surface; and
at least one control surface operably connected to the at least one actuator, wherein the at least one control surface is adjusted by the at least one actuator.
The control surface operating system of claim 1 , wherein the at least one control surface is operably connected to two actuators.
The control surface operating system of claim 1 , wherein the drive system includes a power distribution unit arranged to drive the continuously variable transmission.
The control surface operating system of claim 1 , further comprising a controller arranged to control operation of the continuously variable transmission.
The control surface operating system of claim 4 , wherein the controller is part of the drive system.
The control surface operating system of claim 4 , wherein the controller controls operation of the continuously variable transmission based on a flight condition of the aircraft.
The control surface operating system of claim 1 , wherein the at least one control surface is one of a flap or slat on a wing or tail of the aircraft.
The control surface operating system of claim 1 , wherein the continuously variable transmission includes at least two conical gear sets to enable variable gear ratios of the continuously variable transmission.
The control surface operating system of claim 8 , wherein a belt is driven by the at least two conical gear sets.
The control surface operating system of claim 1 , wherein the drive shaft comprises at least one torque tube.
The control surface operating system of claim 10 , wherein the drive shaft comprises multiple torque tubes that operably connect a plurality of actuators, wherein the plurality of actuators are arranged to control operation of at least two control surfaces.
The control surface operating system of claim 1 , wherein the at least one actuator is a linear actuator.
The control surface operating system of claim 1 , wherein the at least one actuator is a rotary actuator.
An aircraft comprising: at least one aerodynamic control surface; and a control surface operating system for controlling aerodynamic control surfaces of the aircraft, the control surface operating system comprising: a drive system operably connected to a drive shaft, the drive system including a continuously variable transmission and at least one actuator operably connected to the drive shaft and the at least one aerodynamic control surface, the at least one actuator arranged to convert rotational movement of the drive shaft to move at least a portion of the at least one aerodynamic control surface, such that the at least one control surface is adjusted by the at least one actuator.
at least one aerodynamic control surface; and
a control surface operating system for controlling aerodynamic control surfaces of the aircraft, the control surface operating system comprising:
a drive system operably connected to a drive shaft, the drive system including a continuously variable transmission and at least one actuator operably connected to the drive shaft and the at least one aerodynamic control surface, the at least one actuator arranged to convert rotational movement of the drive shaft to move at least a portion of the at least one aerodynamic control surface, such that the at least one control surface is adjusted by the at least one actuator.
The aircraft of claim 14 , wherein the drive system includes at least one of a power distribution unit arranged to drive the continuously variable transmission and a controller arranged to control operation of the continuously variable transmission, the controller configured to control operation of the continuously variable transmission based on a flight condition of the aircraft.
The aircraft of claim 14 , wherein the at least one control surface is one of a flap or slat on a wing or tail of the aircraft.
The aircraft of claim 14 , wherein the drive shaft comprises multiple torque tubes that operably connect a plurality of actuators, wherein the plurality of actuators are arranged to control operation of at least two control surfaces of the aircraft.
A method of controlling operation of aerodynamic control surfaces of an aircraft, the method comprising: driving a drive shaft with a continuously variable transmission, the drive shaft operably coupled to at least one actuator; and converting rotational movement of the drive shaft using the at least one actuator to move at least a portion of the aerodynamic control surface to adjust the aerodynamic control surface.
driving a drive shaft with a continuously variable transmission, the drive shaft operably coupled to at least one actuator; and
converting rotational movement of the drive shaft using the at least one actuator to move at least a portion of the aerodynamic control surface to adjust the aerodynamic control surface.
The method of claim 18 , further comprising controlling operation of the continuously variable transmission based on a flight condition of the aircraft.
Ref US4789119: Claims:
1. In a system for controlling the high-lift flaps of an aircraft comprising two fixed wings, symmetrical to each other with respect to the median horizontal plane of said aircraft and each provided with at least one engine for propulsion and a high-lift flap, each of said flaps being subjected to the blast of at least one engine, said control system comprising:a source of hydraulic fluid under pressure; a unit for distributing hydraulic fluid connected to said source; means for actuating said high-lift flaps controlled by said hydraulic fluid distributing unit; a device for controlling said hydraulic fluid distributing unit; a voluntary control member, at a pilot's disposal and controlling, via said control device, said hydraulic fluid distributing unit and said actuating means, said high-lift flaps between a retracted neutral position and at least one extended lift-augmenting position, and vice-versa,said actuating means are reversible hydraulic jacks, arranged so that each flap is actuated by a jack and each of the two chambers of one jack is connected to the corresponding chamber of the other jack, as well as to an orifice in sa
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