JP7621651B2 - Aircraft Propulsion and Torque Mitigation Technology - Google Patents

Description

The present disclosure relates generally to aircraft propulsion and torque mitigation techniques for aircraft, and specifically for vertical lift aircraft. More specifically, the present disclosure relates to techniques that provide rotational torque to rotate rotor blades of an aircraft, and that mitigate or even eliminate the need for counter-torque methods and devices such as tail rotors and counter-rotating blades.

There are various methods for rotating the rotor blades of a vertical lift aircraft, e.g., a helicopter. One approach utilizes a rotor shaft that connects to the rotor blades, where the rotor shaft is mechanically connected to the engine and to the aircraft fuselage. In that approach, when the engine is enabled, the torque generated to rotate the rotor blades also generates an opposing torque on the aircraft body in the opposite direction. If not countered, the opposing torque can cause the aircraft body to rotate about the axis of the rotor shaft, potentially causing the aircraft to spin uncontrollably. Therefore, counter torque mechanisms have been developed to mitigate or eliminate the opposing torque on the aircraft body, thereby facilitating controlled flight.

One method of torque cancellation is to attach a tail rotor to the aircraft, which generates a torque in the opposite direction to that produced by the rotation of the main rotor by an engine connected to the aircraft fuselage. The tail rotor and its associated assemblies often include linkages to the main engine, gearbox, tail boom drive shaft, tail rotor transmission, the tail rotor itself and associated structural members. Such components add significant weight, power consumption, complexity, initial manufacturing costs and ongoing maintenance costs to the aircraft. If the tail rotor (and/or associated components) has a mechanical failure, is damaged or is lost, the aircraft will often spin uncontrollably and crash. The use of a tail rotor also presents design, acoustic and operational challenges to the aircraft, including, but not limited to, increased weight, increased noise generation, increased vibration, and increased landing and parking space. A tail rotor that spins during operation also presents a safety hazard to personnel outside the aircraft.

Another method of torque cancellation on a vertical lift aircraft involves fitting the aircraft with rotor blades that rotate in counter-rotating directions. Such counter-rotating rotor blades may be coaxial with the main rotor blades or may be on a different axis than the main rotor blades. Yet another method of torque cancellation utilizes multiple rotating propellers on a vertical lift aircraft. As with tail rotors, such methods come at the cost of increased complexity, loss of efficiency, increased weight, and added maintenance and manufacturing costs to the aircraft. Failure of the counter-rotating mechanical elements or propellers can impair or eliminate the ability to autorotate, potentially leading to loss of control of the aircraft. Multiple rotating propeller systems are also often less efficient than single rotor vertical lift systems and may include electric engines powered by currently available battery technology, which may result in reduced flight time due to battery energy density limitations and added weight compared to using liquid fuel as the energy source.

In the past, there have been efforts to develop technology that would eliminate the need for counter-torque mechanisms on vertical lift aircraft. Such efforts have led to the development of two methods of eliminating counter-torque mechanisms: (1) engines mounted on the tips of the rotor blades themselves, as disclosed in U.S. Patent No. 2,761,635, issued Sep. 4, 1956 to S. Heller, Jr. et al., and (2) tubes or ducts along or within the rotor blades themselves, whereby thrust from the engines is delivered through such tubes or ducts to the outer tips of the rotor blades to rotate them. Both of these methods provide for the rotation of the rotor blades without creating torque on the body of the aircraft, by using the rotor blades themselves (i.e., by utilizing thrust devices on, within, directly connected to, and flush with the rotor blades) as a means of generating the torque required for flight. However, these methods are associated with various challenges, such as increased weight and instability for the aircraft's main rotor blades, increased parasitic drag, and excessive noise. The first method also impairs the aircraft's ability to autorotate. As a result, those two methods were not widely used.

Therefore, there remains a need in the field of vertical lift aircraft for other methods and apparatus that reduce or eliminate torque transfer by the engines to the body of the aircraft, thus mitigating or even eliminating the need for the use of counter torque mechanisms. The techniques of the present disclosure address that need and, in embodiments, provide an effective mechanism for reducing or eliminating torque transfer and its associated challenges while maintaining the ability of the aircraft to auto-rotate.

Features and advantages of the embodiments of the claimed subject matter will become apparent as we proceed to the following detailed description and by reference to the drawings, in which like numerals refer to like parts.

FIG. 1 is a schematic side view of an embodiment of the present disclosure, with some mechanical elements of a conventional helicopter omitted from the view for clarity, outlining how the rotational torque is separated from the fuselage and where the rotor blades are located below the rotary thrust support structure. FIG. 2 is a top view of FIG. 1 . FIG. 13 is a side schematic view of an embodiment of the present disclosure showing thrust delivery along a path to the outer tip of a thrust support structure. FIG. 4 is a top view of FIG. 3 . FIG. 2 is a schematic side view of an engine and thrust support structure mounted above a rotor blade, the engine delivering thrust longitudinally along the thrust support structure. FIG. 6 is a top view of FIG. 5 . FIG. 1 is a schematic side view showing an engine mounted perpendicular to and on top of a thrust support structure to generate thrust and to deliver such thrust to the outer tip of such thrust support structure; FIG. 8 is a top view of FIG. FIG. 1 is a schematic side view showing an engine with a gearbox and drive shaft mounted on top of and perpendicular to a thrust support structure, delivering power via the drive shaft to a propulsion means at the tip of such structure, which is above the rotor blades. FIG. 10 is a top view of FIG. FIG. 1 is a schematic side view showing an engine delivering thrust to a manifold, which delivers thrust through passages in a thrust delivery structure to an outer end of such structure, which then rotates to provide torque to a rotor shaft, with rotor blades resting on such thrust delivery structure. FIG. 12 is a top view of FIG. 2 is a schematic side view of an embodiment of the components shown in FIG. 1 arranged such that the rotor blades are above a thrust support structure. FIG. 14 is a top view of FIG. 13 . FIG. 2 is an enlarged cross-sectional view of the components of the present disclosure for the embodiment shown in FIG. 1 in which the thrust support structures are above the rotor blades. FIG. 14 is an enlarged cross-sectional view of the components of the present disclosure for the embodiment shown in FIG. 13 where the thrust support structure is below the rotor blades. FIG. 2 is a schematic cross-sectional side view showing one type of clutch in an engaged position. FIG. 2 is a schematic side view of one type of clutch in a disengaged position. FIG. 2 is a schematic end view showing one type of clutch in an engaged position. FIG. 2 is a schematic end view of one type of clutch in a disengaged position. FIG. 1 is a schematic side view showing an embodiment of the present disclosure in which an engine is fixed to the base of a rotor shaft, an engine drive shaft runs through the interior of the rotor shaft to a gearbox, and another drive shaft runs along a thrust support structure to a drive propulsion means at its tip. FIG. 21 is a top view of FIG. FIG. 2 is an exploded view of one type of orientation control device. 24 is a schematic side view of the same aircraft as shown in FIG. 1, except that this embodiment includes an azimuth control device as shown in FIG. 23. FIG. 25 is a top view of FIG. FIG. 1 is a schematic side view of an embodiment of the present disclosure illustrating a torque that is decoupled from the fuselage, thus eliminating the need for a counter-torque mechanism such as a tail rotor. 1 is a schematic side view of a conventional helicopter illustrating torque being transmitted to the fuselage, thus necessitating a counter-torque mechanism, in this case the tail rotor. FIG. 1 is a schematic side view illustrating an embodiment of the present disclosure including both a azimuth control device and a tail rudder for azimuth control. FIG. 1 is a schematic side view illustrating an embodiment of the present disclosure including only a tail rudder for azimuth control. FIG. 1 is a schematic side view illustrating an embodiment of the present disclosure in which there is no tail rudder and a azimuth control device provides azimuth control. 1 illustrates the airflow of a conventional helicopter under normal conditions. 1 illustrates conventional helicopter airflow with a vortex ring condition. 1 illustrates an embodiment of the present invention showing airflow under normal conditions. 1 illustrates an embodiment of the present invention showing a turbulent vortex ring condition. 1 illustrates one embodiment of a rotating union consistent with the present disclosure. FIG. 2 is a cross-sectional side view of an embodiment consistent with the present disclosure in which one or more engines are integrated directly into a thrust support structure; FIG. 2 is a top view of an embodiment consistent with the present disclosure in which one or more engines are integrated directly into a thrust support structure. FIG. 2 is a cross-sectional side view of an embodiment consistent with the present disclosure in which one or more engines are mounted directly to a thrust support structure that is directly connected to a rotor shaft. FIG. 2 is a top view of an embodiment consistent with the present disclosure in which one or more engines are mounted directly to a thrust support structure that is directly connected to a rotor shaft. FIG. 1 is a schematic diagram of an aircraft consistent with the present disclosure utilizing either a motorized tail blade 77 or tail fan 78 assembly for directional (azimuth) control. FIG. 1 is a schematic side view illustrating an embodiment of the present disclosure using concentric inner and outer shafts to support and/or separate various aircraft components. 1 is a schematic side view illustrating one embodiment of an aircraft including a propulsion machine element capable of moving between vertical and horizontal orientations consistent with the present disclosure. 1 is a schematic side view of an example of an aircraft propulsion system including one or more damping elements consistent with the present disclosure. 1 is a schematic side view of an aircraft including a pulse jet engine consistent with the present disclosure; 1 illustrates an embodiment of an orientation control device consistent with the present disclosure. 1 is a schematic diagram of one embodiment of a propulsion system including variable rotor blades consistent with the present disclosure. FIG. 1 is a schematic diagram of one embodiment of a propulsion system consistent with the present disclosure in which various components are cooled by airflow. 1 illustrates a side and longitudinal cross-sectional view of one embodiment of a thrust support structure including a baffle consistent with the present disclosure; 1 illustrates an embodiment of a propulsion system in which the thrust support structure and rotor blades are connected to the rotor shaft without a clutch, consistent with this disclosure. 1 illustrates an embodiment of an aircraft propulsion system including a sliding rail system consistent with the present disclosure. FIG. 1 is a side view of an example of a propulsion system including a thrust support structure including one or more articulating nozzles. 1 illustrates one embodiment of a propulsion system including multiple sets of rotor blades driven by a single thrust support structure consistent with the present disclosure. 1 illustrates one example of the use of the techniques of the present disclosure in a composite helicopter environment. 13 illustrates another example of the use of the techniques of the present disclosure in a composite helicopter environment. 1 illustrates a top and side view of one embodiment of a propulsion system consistent with the present disclosure, the propulsion system not including a clutch and including a thrust support structure and rotor blades that are flush with one another; 1 depicts one embodiment of the present disclosure in which one or more fan units are driven by an engine to generate airflow that can be directed in a fixed or variably direction. 1 depicts another embodiment of the present disclosure in which one or more fan units are driven by an engine to generate airflow that can be directed in a fixed or variably direction. FIG. 2 is a side view of another example embodiment of the present disclosure in which separate shafts are utilized for different components of an aircraft propulsion system. FIG. 13 is a side view of another exemplary embodiment of the present disclosure, in which a fuel tank, a battery, and an engine are connected to a thrust support structure. 2 is a schematic cross-sectional side view of another embodiment of an aircraft including a propulsion system consistent with the present disclosure. 1 illustrates a top and side view of one embodiment of a propulsion system consistent with the present disclosure, the propulsion system including a thrust support structure configured to function as a rotor blade for an aircraft. 1 represents one embodiment of the present disclosure, in which a thrust support structure consistent with the present disclosure is configured to function as a rotor blade for an aircraft, and in which airflow generated by one or more engines is directed in a fixed or variable direction. 1 is a schematic cross-sectional side view of another embodiment of an aircraft including a propulsion system consistent with the present disclosure, the propulsion system including a thrust support structure configured to function as a rotor blade of the aircraft. 1 is a schematic cross-sectional side view of another embodiment of an aircraft including a propulsion system consistent with the present disclosure, the propulsion system including a thrust support structure configured to function as a rotor blade for the aircraft and including one or more articulating nozzles. FIG. 2 is a schematic cross-sectional side view of another embodiment of an aircraft including a propulsion system consistent with the present disclosure, in which airflow from one or more fuselage-mounted engines is directed into ducts within a pair of counter-rotating rotor blades. 2 is a schematic cross-sectional side view of another embodiment of an aircraft including a propulsion system consistent with the present disclosure, in which a portion of the airflow from one or more fuselage-mounted engines is directed toward the rear of the aircraft. 2 is a schematic cross-sectional side view of another embodiment of an aircraft consistent with the present disclosure.

Aspects of the present disclosure relate to systems, devices, and methods for reducing or eliminating torque transfer in aircraft, such as vertical lift aircraft. In embodiments, aspects of the present disclosure include methods and devices in which the transmission of torque generated by one or more engines 5 to the aircraft fuselage 4 is reduced or even eliminated, for example, by one or more support bearings 2. In some examples, the amount of torque generated by the one or more engines 5 does not exceed the capacity of the support bearings 2 from isolating the engines 5, rotor shaft 1, and their associated components from the fuselage 4. The reduction/elimination of torque transfer to the aircraft fuselage 4 reduces or eliminates the need for counter-torque mechanisms, such as tail rotors or counter-rotating blades. In particular, the techniques described herein maintain the ability of the aircraft to auto-rotate in the event of loss of rotational thrust from the engines 5.

The aircraft body may include or be coupled to one or more movable units, such as, but not limited to, the fuselage 4, cargo boxes, camera and surveillance packages, weapons, aircraft superstructure, and flight and control equipment. In one embodiment, the disclosed technique employs an engine 5 capable of providing thrust to a support structure 7. The thrust support structure 7 is not flush with the rotor blades 3, but is coaxial with and extends to either side of the rotor shaft 1. In response to thrust applied by the engine 5, the thrust support structure 7 rotates about the axis of the rotor shaft 1. Such rotation engages the clutch 6 (connected to push the support structure 7) to the rotor shaft 1, causing the rotor shaft 1 to rotate. The rotation of the rotor shaft 1 in turn rotates the rotor hub 11 and the attached rotor blades 3.

When the thrust support structure 7 is not receiving thrust from the engine 5, the clutch 6 separates the thrust support structure 7 from the rotor shaft 1. In that state, the rotor shaft 1, rotor hub 11 and rotor blades 3 are free to rotate. In an embodiment, the rotor shaft 1 is connected to the fuselage 4 only by one or more support bearings 2, which allow the rotor shaft 1 (and attached rotor blades 3) to rotate freely relative to the fuselage 4 at all times. In some embodiments, the techniques described herein include a tube/duct configured to supply airflow (e.g., generated by the engine 5) to the outer tip of the thrust support structure 7. Such a tube/duct can be included, coupled to, or supply airflow to the outside or interior of the thrust support structure 7, for example, via one or more channels or cavities on the inside or outside of the thrust support structure 7. In an embodiment, the end of the tube/duct is positioned at an outlet from the tip of the thrust support structure 7. Thus, airflow through the tube/duct can cause rotation of the thrust support structure 7 around the axis of the rotor shaft 1. Rotation of the thrust support structure 7 in turn engages the clutch 6 with the rotor shaft 1, thus causing rotation of the rotor hub 11 and associated rotor blades 3. As previously mentioned, when the engine 5 is not providing thrust to the thrust support structure 7, the clutch 6 disengages the thrust support structure 7 from the rotor shaft 1, allowing the rotor shaft 1 and associated rotor blades 3 to freely rotate automatically.

The support bearings 2 generally function to isolate torque generated by the engine 5 and/or rotation of the aforementioned mechanical elements from the fuselage 4. In an embodiment, the torque generated by the rotation of the engine 5 and/or the aforementioned components does not exceed the ability of the support bearings 2 to isolate such torque from the fuselage 4. Because of such isolation, counter-torque devices are not required and can be eliminated from the aircraft design. More specifically, the techniques described herein can eliminate the need for counter-torque devices, such as tail rotors or counter-rotating wings, and all their associated costs, power consumption, weight, maintenance and safety issues.

In yet another embodiment of the present invention, the techniques described herein utilize engines 5 that are located at various positions relative to the thrust support structures 7. Additionally, the thrust support structures 7 can be located at various positions relative to other components of the aircraft. For example, the thrust support structures 7 can be coaxial with the rotor blades 3 and perpendicular to the rotor shaft 1, and can be located above, below, or above and below the rotor blades 3. It should be noted that there can be more than one support structure 7, either all on the same plane or on different planes that are coaxial with the rotor shaft 1. For example, in an embodiment, a first thrust support structure 7 can be located coaxially with and above the rotor blades 3, and a second thrust support structure 7 can be located coaxially with and below the rotor blades 3, with the first and second thrust support structures 7 each being coupled to one or more engines 5.

The techniques disclosed herein can provide numerous advantages, especially when compared to counter-torque rotation solutions, including, but not limited to: (1) elimination of anti-torque mechanisms such as tail rotors and counter rotors; (2) azimuth control can be provided by azimuth control device 12 or a similar device performing the same function such as azimuth control device 12 imposing a rotational force on the fuselage for directional control and stabilization, or azimuth control can be provided by tail rudder 8, or azimuth control can be provided by a combination of tail rudder 8 and azimuth control device 12, so the aircraft need not have a tail at all; (3) the risk of boom strike can be eliminated since the tail rotor and its extended tail boom are no longer required; (4) deliverable payload for power capacity and the like can be increased since the weight and drag of the tail rotor, boom, gearbox and associated structural components are eliminated; (5) tail rotor noise is eliminated since a tail rotor is not required; (6) disk tilt is eliminated since the tail rotor is eliminated, so the aircraft can take off, land and hover without the need to counterbalance tail rotor thrust by cyclic trim; (7) the tail rotor or a second set of rotor blades can be driven without the need to counterbalance tail rotor thrust by cyclic trim; (8) manufacturing and maintenance costs can be reduced; (9) hazards associated with a rotating tail rotor can be eliminated; (10) hazards associated with loss of control due to loss, failure or damage to the tail rotor and its structural and operating components can be reduced or even eliminated; (11) takeoff, landing and parking footprint can be reduced; (12) ability to utilize existing manufacturing and testing technology with currently available ground facilities, already trained pilots and pilot training procedures, federal, state and local regulatory controls, with the added benefits of increased fuselage capacity and increased safety; (13) easy adaptability of the technology for a variety of aircraft sizes from smaller drones to large general and military aircraft; (14) ability to utilize a wide variety of engines; (15) the enhanced autorotation capabilities of the present invention provide enormous safety advantages over many of the multiple rotor designs proposed by others. (16) Vortex ring relaxation by the thrust lines of the engine 5 creates a high-velocity airflow that fans out horizontally over the rotor blades 3, resulting in a low-pressure area over the rotor blades 3, further improving lift, etc. These advantages are listed by way of example, and other advantages will be apparent to those skilled in the art.

Figure 1 is a representative side view of one embodiment of the present disclosure, showing components that allow the aircraft of the present invention to operate without a counter-torque device such as a tail rotor or counter-rotating rotor blades. Figure 2 is a top view of the same embodiment. As shown, in Figure 1, the rotor shaft 1 and all its appendages are supported by an assembly consisting of a support bearing 2, a bearing support 46, and a bearing support structure 43. The resulting assembly is attached to the fuselage 4 such that the only points of contact of the rotor shaft 1 and its appendages with the fuselage are the bearing support structure 43 and the support bearing 2. The torque exerted by the rotor shaft 1 and its appendages does not exceed the ability of the support bearing 2 to isolate the fuselage 4 from said torque. Since the torque is not transmitted to the fuselage, the requirement for a counter-torque device used in conventional helicopters can be eliminated.

Rotating Thrust Support Structure As illustrated in Figures 1, 2 and 15, torque generated by the engine 5 is delivered to the rotor shaft 1 by rotation of the thrust support structure 7. The thrust support structure 7 is a structural member that supports the engine 5 located at its outer end. In the illustrated embodiment, the thrust support structure 7 is coaxial and perpendicular to the rotor shaft 1 and lies in a different plane than the rotor blades 3 (e.g., above the rotor blades 3). However, the thrust support structure 7 and the rotor blades 3 may be in a common plane, as described below. In any case, the thrust support structure 7 rotates about the rotor shaft 1 when the engine 5 is operational.

Clutch FIG. 15 is a partial cross-sectional view of the components of the present disclosure for the embodiment of FIG. 1 where the thrust support structure 7 is above the rotor blades 3. As shown, the thrust support structure 7, with the engine 5 connected at its outer end, is coaxial with the rotor shaft 1 and is located above the rotor blades 3. The thrust support structure 7 is connected to a clutch 6 having an engaged and disengaged state. When the engine 5 is enabled, they cause the thrust support structure 7 to rotate and cause a rotational torque. Such torque causes the attached clutch 6 to engage and clamp tightly against the rotor shaft 1, which in turn rotates the rotor hub 11 and the attached rotor blades 3. The clutch 6 can be a one-way bearing or any suitable device that performs the same function. Numerous devices that perform this function in various ways are known to those skilled in the art.

One type of clutch that can be used is illustrated in more detail in Figures 17, 18, 19 and 20. Figures 17 and 19 illustrate the clutch 6 in an engaged state, positively gripping the rotor shaft 1. Figures 18 and 20 illustrate the clutch 6 in a disengaged state, where the engine 5 is not operational and the rotational speed of the thrust support structure 7 is less than the rotational speed of the rotor shaft 1. The clutch 6 and its associated thrust support structure 7 then automatically disengage from the rotor shaft 1, allowing the rotor shaft 1 to rotate freely, which allows the rotor blades 3 attached to the rotor shaft 1 to rotate automatically.

The operation of one example of the clutch 6 will now be described in connection with Figures 17, 18, 19 and 20. Figure 19 is a schematic diagram of the clutch 6 in an engaged position. The rotational movement of the clutch outer case 58 moves the clutch mobile bearing 52 along the incline to the position shown in the clutch bearing cavity 64. When the clutch mobile bearing 52 is in this position, the clutch outer case 58 is brought into contact with both the clutch outer case 58 and the clutch inner race 114. In that position, the clutch mobile bearing 52 securely grips the clutch inner race 114 (which is directly connected to the rotor shaft 1) to rotate the rotor shaft 1 in the same direction as the clutch outer case 58. Figure 17 is a cross-sectional schematic side view of the type of clutch 6 shown in Figure 19 in an engaged position. The thrust bearing assembly 57 is attached to the rotor shaft 1 by fasteners 54 that hold the clutch 6 in place in its position on the rotor shaft 1 and allow the clutch 6 to engage or disengage the rotor shaft 1. As shown in FIG. 17, the clutch movable bearing 52 can be securely engaged with the rotor shaft 1, so that the clutch 6 and the rotor shaft 1 can rotate together at the same time.

20, which is a schematic diagram of the clutch 6 in the disengaged position. With the absence or reduction of rotational momentum of the clutch outer case 58, the clutch outer case 58 begins to decelerate in its relationship to the rotor shaft 1. The clutch moving bearings 52, due to their momentum and centrifugal forces, move away from and loosen their grip on the clutch inner race 114 (which is again attached to the rotor shaft 1), thus disengaging the rotor shaft 1 from the clutch outer case 58 and allowing the rotor shaft 1 to rotate freely with respect to the clutch 6. Also, the clutch 6 is disengaged from the rotor shaft 1 when the speed of the rotor shaft 1 is greater than the rotational speed of the clutch 6. The disengaged state of the clutch 6 allows the rotor shaft 1 to rotate freely and to autorotate without any induced drag from the thrust support structure 7 and the power system. Also, during autorotation, the preserved inertia of the power system is transferred to the rotor shaft 1 at any time when the speed of the rotor shaft 1 is equal to or less than the rotational speed of the clutch 6. While this condition exists, the power system inertial transfer provides more time to reduce the overall pitch to the safe angle required for the aircraft to autorotate.

Figure 38 shows another embodiment of the clutch 6, which in this case is directly connected to the rotor hub 11. This configuration allows the rotor blades 3 to rotate independently of the rotor shaft 1 and all other drive system components. The engine 5 is integrated into a directly connected thrust support structure 7 to drive the rotor shaft 1. This configuration with respect to its location on the rotor shaft can be applied with the thrust support structure 7 located above, below and/or in the same plane as the rotor blades 3.

Fuel, Power and Data Delivery System Figures 1 and 2 show the fuel tank 33 and fuel pump 34 connecting to the lower fuel line 35. Figures 1 and 15 show the lower fuel line 35 connecting to the fuel line shaft 48 which is connected to the lower rotary union 16. The lower rotary union 16 is shown in more detail in Figure 35. In an embodiment, the lower rotary union 16 and the upper rotary union 38 may be the same device with different orientations. As shown in Figures 35 and 15, the lower rotary union 16 has a fuel input coming from the lower fuel line 35, a data input coming from a data line 68 and a power input coming from a power line 69. Figure 15 shows a data conduit 66 with data line 68 running from the lower rotary union 16 up through the rotor shaft 1 to the upper rotary union 38. Figure 35 shows how fuel, data and power are transferred through the lower rotary union 16 and the upper rotary union 38. In the embodiment of Figure 15, after leaving the upper rotating union 38, the data wiring 68 runs through suitable protective shielding, along or through the thrust support structure 7 to the engines 5 or such other device needing data. The power wiring 69 runs in a similar manner to the data wiring, except through a separate power conduit 72, from the bottom of the rotor shaft 1 to the top of the rotor shaft 1 to the engines 5 or other device needing power. Figure 35 also shows details of the lower rotating union 16 and upper rotating union 38, as they are the same device with different orientations.

Fuel can be pumped through the lower rotary union 16 to the rotor shaft internal fuel line 47 that rotates in unison with the rotor shaft 1. FIG. 15 shows how fuel can be pumped through the rotor shaft internal fuel line 47 that runs up through the inside of the rotor shaft 1 to the upper rotary union 38. In the embodiment of FIG. 35, the rotary union seals 49 prevent fuel leakage as the lower rotary union 16 and the upper rotary union 38 rotate with the rotor shaft 1. The lower rotary union 16 and the upper rotary union 38, which are attached to the rotor shaft 1 by the rotary union bearings 50, can rotate freely with respect to the fuel line shafts 48 at the lower and upper ends of the rotor shaft 1. Fuel is then delivered from the fuel line shaft 48 at the upper end of the rotor shaft 1 through the upper fuel line 37 to the engine 5. In FIG. 15, the upper fuel line 37 is fixed to or housed inside the thrust support structure 7. These methods of fuel transportation and delivery for helicopter aircraft perform similar functions as those disclosed in U.S. Patent No. 2,761,635 (Hiller). Other elements such as power, air pressure and data can also be delivered from the lower part of the rotor shaft 1 to the upper part of the rotor shaft 1 by similar rotary union or slip ring type devices. The various figures do not show all conventional helicopter operation and control devices that can be used, but show the rotating swash plate 40 shown in representative form in Figures 1 and 15 as an example of such a device.

Figure 35 depicts one example of a rotating union that performs the functions of delivering fuel, current, and data to and from the rotor shaft 1. The rotating union provides a simple way to connect components that may be rotating at different speeds relative to one another while ensuring continuity of those connections. The primary structure of the rotating union consists of the rotating union housing 63, which contains the rotating union bearings 50 that allow the rotating union housing 63 to rotate freely about the rotating union shaft 55.

The lower fuel line 35 is connected to the rotary union shaft 55 by a threaded connection that allows fuel to be transferred to the body of the rotary union shaft 55. The rotary shaft internal fuel line 47 is threaded into the rotary union housing 63 and presses up against the rotary union seal 49 to allow fuel transfer through the rotary shaft 1. The data wires 68 and power wires 69 pass through machined passages in the rotary union shaft 55. These wires connect to contacts 70 that allow continuity of these connections to brushes 67 contained within the bearing support 46. The wires continue from the brushes 67 through passages in the rotary union housing 63 and into conduits (not shown) that run along the interior of the rotor shaft 1.

In an embodiment, the primary structure of the rotary union includes a rotary union housing 63 containing a rotary union bearing 50 that allows the rotary union housing 63 to rotate freely with respect to the rotary union shaft 55. The lower fuel line 35 is connected to the rotary union shaft 55 by any suitable means (e.g., a threaded connection) to allow fuel to be transferred to the body of the rotary union shaft 55. For example, the rotary shaft internal fuel line 47 may be threaded into the rotary union housing 63 and press up against the rotary union seal 49 to allow fuel transfer through the rotary union shaft 1. Data lines 68 and power lines 69 pass through machined passages in the rotary union shaft 55. Such wiring connects to contacts 70 that allow continuity of these connections to brushes 67 contained within the bearing support 46. The wiring continues from the brushes 67 through passages in the rotary union housing 63 and into conduits (not shown) that run along the interior of the rotor shaft 1.

The azimuth control device 12 is shown in representative form in Figs. 1 and 15 and in an exploded view in Fig. 23 to explain how such an assembly can be used for azimuth control of an aircraft. In this embodiment, the azimuth control device 12 is an electric reversible motor, the shaft of which is the rotor shaft 1. When azimuth control is required, the motor is momentarily energized to generate a rotational force on the fuselage. Due to the mass rotational speed of the rotor shaft 1 and its auxiliary components, the net effect of the energized stator 18 causes a reaction force against the rotating magnet 17 attached to the spacer 19 through the azimuth control device spacer hole 44, which is attached to the rotor shaft 1 by a key and keyway 23. The rotor shaft 1, which is also the shaft of the azimuth control device 12, is separated from the motor body support plate 20 by the azimuth control motor bearing 21. The azimuth control motor bearing 21 is held by a shaft clamp 22 attached to the rotor shaft 1. All of the motor body support plate 20 and stator 18 are coupled together with fasteners 54 (e.g., bolts) through the azimuth control stator retaining holes 51, thereby attaching the stator 18 and motor body support plate 20 assembly to the fuselage 4 by the fuselage structural members 56. When the stator motor windings 73 are energized, they rotate the attached stator 18 and connected components and the fuselage 4. Since the azimuth control 12 is essentially a reversible motor, the effect created by energizing the azimuth control 12 can cause the aircraft to yaw clockwise or counterclockwise. It should be noted that increasing the distance of the stator motor windings 73 and the magnets on the magnet rotor 17 from the center of the rotor shaft 1 provides more torque and efficiency for the azimuth control 12. Axial flux motors of this design are similar to those utilized in electric and hybrid vehicles.

The present disclosure also provides other mechanisms for controlling azimuth, some of which are illustrated in Figures 28, 29, and 30. Figure 28 is a representative side view of an aircraft using a tail rudder 8 and an azimuth control device 12 for azimuth control. Figure 29 is a representative side view of an aircraft using only a tail rudder 8 for azimuth control. Figure 30 is a representative side view of an aircraft utilizing only an azimuth control device 12 for azimuth control.

Comparison with Anti-Torque Mechanisms FIG. 27 is a side schematic view of a conventional helicopter showing how the engine 5 and other rotating components such as rotor shaft 1 and rotor blades 3 create a rotational torque that is applied to the fuselage 4 by a rigidly connected support mechanism 62. Because such torque applied to the fuselage 4 is not isolated from the fuselage 4, it will rotate the fuselage 4 around the axis of the rotor shaft 1 unless it is countered. Conventional helicopters use a counter-rotation (anti-torque) mechanism that creates a counter torque in the direction of the scene to keep the aircraft from spinning uncontrollably. One example of this concept is shown in FIG. 27, which shows an aircraft with a tail rotor 59 and tail boom as an anti-torque mechanism. In contrast, FIG. 26 represents an embodiment of the present disclosure in which the torque applied by the rotor shaft 1 and its appendages is isolated from the fuselage 4, so that the fuselage 4 is not forced to rotate around the rotor shaft 1, eliminating the need for a counter-torque device such as a tail rotor.

Vortex Ring Mitigation Conventional helicopters are subject to a dangerous phenomenon called vortex ring. This is an aerodynamic condition created when the tip vortex of the main rotor grows because the aircraft remains in its own downwash, which allows the tip vortex to build up and grow in size and strength. If the aircraft does not move out of its downwash and into newer, less disturbed air, the vortex ring can reduce the amount of lift generated by the main rotor due to positive air pressure waves forming above the rotor. This can result in a sudden and uncontrolled loss of altitude that can lead to a crash.

The techniques of the present disclosure can result in a strong countercurrent of air that can inhibit or inhibit the formation of vortices and air pressure above the rotor blades. In some embodiments, the countercurrent is generated by placing the thrust lines of the engines 5 directly in the path of the vortices. The thrust lines of the engines 5 create high velocity airflow that fans horizontally above/below the rotor blades 3 and generally in the same plane as the rotor blades 3. The high velocity airflow can also cause a low pressure area to form just behind the engines 5 and above the rotor blades 3, further improving lift and mitigating the effects of the vortex ring condition. The relationship of normal and disturbed airflow due to the vortex ring condition on a conventional aircraft and the aircraft of the present disclosure is illustrated in Figures 31, 32, 33, and 34.

Additional Embodiments Although the above description focuses on an embodiment in which the aircraft's propulsion system mechanical elements (e.g., rotor shaft 1, rotor blades, thrust support structure 7, etc.) remain in substantially the same orientation relative to the aircraft's fuselage 4, such a configuration is not required. Indeed, the present disclosure includes and contemplates an embodiment in which the orientation of various aircraft's propulsion system mechanical elements can change. Such an embodiment may be useful, for example, for tilt rotor, vertical take-off and landing (VTOL), short take-off and landing (STOL), and short take-off and vertical take-off and landing (STOVL) aircraft. In that regard, reference is made to FIG. 42, which illustrates one exemplary aircraft configuration in which the aircraft's propulsion system mechanical elements can transition between multiple orientations relative to the fuselage 4. More specifically, FIG. 42 illustrates one example of an aircraft including rotor shaft 1, rotor blades 3, fuselage 4, engine 5, clutch 6, thrust support structure 7, tail rudder 8, and rotor hub 11, and such mechanical elements can transition from a vertical orientation to a horizontal orientation as illustrated in FIG. 42. Movement of such components between vertical and horizontal orientations (and vice versa) is accomplished using pivot 15. In this embodiment, pivot 15 allows for a range of motion of approximately 90 degrees, however pivot 15 can be configured to provide any desired range of motion. Additionally, in embodiments, pivot 15 is configured to allow stable retention of the above-mentioned mechanical elements of the propulsion system at any position between the vertical and horizontal orientations shown in FIG. 42. Thus, for example, if pivot 15 allows for a range of motion of 90 degrees, it can be configured to stably hold the mechanical elements of the propulsion system at any angular orientation between approximately 0 degrees and approximately 90 degrees relative to fuselage 4.

The nature and function of the rotor shaft 1, rotor blades 3, engines 5, clutches 6, thrust support structures 7, and rotor hubs 11 in FIG. 42 are the same as those described above and will not be repeated for brevity. And, consistent with the above description, the arrangement of such mechanical elements is not limited to the arrangement shown in FIG. 42. For example, the thrust support structures 7 and engines 5 can be arranged above the rotor blades 3, below the rotor blades 3, flush with the rotor blades 3, or combinations thereof. Thrust from a single engine 5 or multiple engines 5 can be applied to the thrust support structures 7 in any suitable manner. For example, the engines 5 can be directly coupled to the thrust support structures 7 and/or thrust from the engines 5 can be applied to the outer ends of the thrust support structures 7 by ducted flow. Multiple relocatable propulsion units (each including a rotor shaft, rotor blades, engines, clutches, thrust support structures, rotor hubs, etc.) can also be used on an aircraft, as would be understood by one skilled in the art. Additionally, and as shown in FIG. 42, the wings 28 and one or more horizontal stabilizers 92 can be positioned in suitable locations on the aircraft to provide, for example, lift, control, and/or stabilization. In use, the wings 28 and/or horizontal stabilizers 92 can also be reoriented between multiple positions (e.g., horizontal and vertical orientations). To that end, one or more pivots or other control mechanisms for reorienting the wings 28 and/or horizontal stabilizers 92 can also be included on the aircraft. Finally, while FIG. 42 depicts an embodiment using a single set of rotor blades 3 and associated drive elements, multiple (e.g., two, three, four or more) sets of rotor blades 3 and associated drive elements can also be used.

Vibrations, harmonic motions, oscillations, instabilities, etc., of the fuselage 4 or other components may impart mechanical or other stresses to various elements of the aircraft described herein. For example, when conventional transmissions are used to interconnect the rotor shaft to the fuselage of a rotating aircraft, vibrations, harmonic motions, and/or stresses from the rotor blades may be transmitted through the transmissions to the fuselage. This may result in undesirable operating conditions for the aircraft and/or pilot. In particular, the size and location of conventional transmissions may impede or even prevent the use of certain options for mitigating the transmission of vibrations and/or harmonic motions from the rotor blades to the fuselage of the aircraft.

With that in mind, the propulsion systems described herein do not require the use of conventional transmissions to interconnect the rotor shaft with the fuselage, as discussed above. Rather, the propulsion systems described herein use a clutch 6, which can be located outside the fuselage 4. As a result, volume within the fuselage 4 that could be occupied by a conventional transmission can be made available for other uses of the aircraft of the present disclosure. For example, such volume can be used to house one or more damping elements, which are configured to mitigate or even prevent the transmission of vibrations and/or harmonic motions from the rotor blades 1 to the fuselage 4 (or mechanical elements therein).

To address such issues, one or more damping elements may be utilized to attenuate or mitigate the undesirable forces/stresses. In that regard, reference is made to FIG. 43, which is a schematic side view of one embodiment of an aircraft propulsion system including one or more active and/or passive damping elements consistent with the present disclosure.

In this embodiment, the fuselage 4 includes or defines an interior volume within which the fuselage structural member 56 is disposed. The rotor shaft 1 is coupled to the clutch 6 and the rotor blades 3 and extends into the fuselage 4. A number of damping elements are included within the fuselage 4 to mitigate, dampen, or even prevent the transmission of vibrations and/or undesirable (e.g., harmonic) motion from the rotor blades 3 to the fuselage 4.

More specifically, FIG. 43 depicts an example embodiment in which a snubber 82 is disposed between the support bearing 2 and the fuselage structural member 56. In general, the snubber 82 is configured to dampen or mitigate the transmission of mechanical and other forces (e.g., due to vibrations, harmonic motions, instabilities, etc.) from the rotor shaft 1 (and/or rotor blades 3) to the fuselage 4 and vice versa. In embodiments, the snubber 82 can be in the form of a passive damping element such as a gas, liquid, or mechanical snubber (e.g., gas/liquid struts, one or more damping springs, combinations thereof, etc.). For example, in some embodiments, the snubber 82 includes a housing (e.g., a metal body) and a resilient material (e.g., a polymer such as rubber) that can absorb or mitigate vibrations or other undesirable motions, thus limiting or preventing the transmission of such forces from the rotor shaft 1 to the fuselage structural member 56.

One or more active damping elements may also be used to mitigate or prevent the transmission of vibrations or other undesirable forces from the rotor shaft 3 to the fuselage 4. This concept is illustrated in FIG. 43, which depicts an example embodiment in which a vibration reduction actuator 106 is utilized. In general, the vibration reduction actuator 106 is an active damping element that acts to cancel or dampen vibrations by application of an appropriate force. For example, the vibration reduction actuator 106 may be configured to match and/or cancel harmonic frequencies or vibrations coming from the rotor blades 3 or the support structure 7, thereby mitigating or even eliminating their transmission to the fuselage 4. In an embodiment, the vibration reduction actuator 106 includes an outer (e.g., copper) coil, and an inner element (e.g., an iron rod or tube) is disposed within the open center of the coil. Appropriate application of electrical energy to the coil may result in the generation of an electromagnetic field that changes the position of the inner element. Movement of the inner element may be controlled by applying electrical energy to the coil in such a manner as to cancel or mitigate the incoming vibrations, harmonic frequencies, etc., as will be appreciated by those skilled in the art.

Although FIG. 43 depicts the use of a single damper 82 in combination with a single vibration reduction actuator 106, such a configuration is not required. It should be understood that any number of dampers 82 and vibration reduction actuators 106 may be used and such mechanical elements may be used independently of one another. Furthermore, the location and configuration of such mechanical elements are not limited to the locations shown in FIG. 43.

Various types of engines may be used as the engines 5. Non-limiting examples of suitable engines that may be used as the engines include shaft engines such as reciprocating (piston) engines and turbine engines, reaction engines such as jet engines, pulse jet engines, turbofan engines and rocket engines, Wankel engines, diesel engines, electric engines, combinations thereof, and the like. In some embodiments, the engines 5 may be pulse jet engines. For example, in the embodiment of FIG. 44, the engines 5 are in the form of pulse jet engines. In this embodiment, each of the engines 5 is surrounded by a shell 84 and includes a vertical air intake 88. In a non-limiting embodiment, the shells 84 and the vertical air intakes 88 are configured to limit, attenuate, or even prevent sound from escaping from the engines to the surrounding environment, and in particular, toward the ground and/or fuselage 4. In this example, thrust from the engines 5 is directed through the air gap 83 before entering the input duct 10, which acts as a thrust augmenter that entrains additional air from the air gap 83 to increase engine power. The thrust is then discharged from the tip of the duct 10, rotating the thrust support structure. As will be apparent, the use of the shell 84 and duct 10 can limit or prevent sound waves from escaping into the environment, reducing the amount of noise generated during the operation of the aircraft.

As noted above, various mechanisms can be utilized to provide azimuth control to an aircraft consistent with the present disclosure. With that in mind, FIG. 45 depicts another example of an azimuth control device consistent with the present disclosure. Similar to azimuth control device 12 described above, FIG. 45 depicts an embodiment of an azimuth control device in the form of an electric reversible motor, the shaft of the motor being rotor shaft 1. However, in this case, the motor is a reversible motor including a rotating magnet 17 and a rotating magnet support 87, both of which are coupled to rotor shaft 1. Stator 18 is coupled to the aircraft fuselage 4 by fuselage structural member 56. By operating the motor in one direction or another, a reaction force can be transferred to stator 18, which (because it is attached to fuselage 4) applies a force to fuselage 4 for azimuth control.

The preceding description has focused on an embodiment in which the rotor blades 3 are of fixed length. However, such a configuration is not required, and any fixed or variable length rotor blades can be used as the rotor blades 3. In an embodiment, the rotor blades 3 are of variable length and can be configured to adjust and/or increase the thrust generated by the aircraft. For example, variable length rotor blades 3 can be used to improve the kinetic thrust of a rotor vehicle, where the length of the rotor blades 3 may be reduced when transitioning from vertical lift to horizontal thrust for faster forward flight. In that regard, reference is made to FIG. 46, which is a schematic diagram of one example of a system including variable rotor blades consistent with the present disclosure. As shown, the system includes a blade grip 121 connected to one (initial) end of each of the rotor blades 3 (in this case the immediately adjacent rotor shaft 1). Each blade grip 121 includes a motor 91 (e.g., a linear electric or hydraulic motor) coupled to a lead screw 119, which is in turn coupled (e.g., screwed) to each rotor blade 3. Operation of the motor 91 may rotate the lead screw, resulting in the extension or contraction of the rotor blades 3. In an embodiment, the motors 91 on the opposing blade grips 94 can be linked to a common shaft (e.g., a common lead screw) to provide linear positioning of the opposing rotor blades. Of course, any other suitable mechanism for extending and retracting the rotor blades 3 can also be used.

During operation, the engine 5 generates heat and one or more exhaust streams that can change the infrared signature of the aircraft. The operation of the engine 5 and the rotation of the thrust support structure 7 can increase the temperature of the thrust support structure 7. This can be undesirable in some applications, particularly military applications where the infrared signature of the aircraft is important. With that in mind, some aspects of the present disclosure relate to systems and methods for regulating the temperature of the thrust support structure consistent with the present disclosure. In that regard, reference is made to FIG. 47, which illustrates one example of a propulsion system consistent with the present disclosure, in which various components are cooled by airflow. In the illustrated embodiment, the engine 5 is in the form of a jet engine that directs thrust into ducting 10 within the thrust support structure 7. Compressor bleed air 85 (i.e., air taken from behind the compressor stage of the fuel combustion section of the jet engine) is directed toward or injected into the thrust periphery. Since the compressor bleed air 85 is cooler than the thrust, the mixing of the compressor bleed air 85 with the thrust cools the ducting 10 and therefore the thrust support structure 7. This technique can be applied in the same way in other types of engines 5. The above description has focused on embodiments in which the thrust support structure 7 is often exposed in whole or in part to the external environment. While such embodiments are useful, the aerodynamics of the thrust support structure 7 itself may not be ideal for some applications. For example, the shape of the thrust support structure 7 may affect the drag coefficient and efficiency of the aircraft in flight. For example, if the shape of the thrust support structure 7 is spherical, the drag coefficient is the same regardless of the aircraft attitude during hover or in forward flight. Although a spherically shaped thrust support structure may simplify the aircraft design, it may not provide the desired drag coefficient. For example, a smaller drag coefficient may be obtained using an airfoil shape as opposed to a sphere.

With that in mind, aspects of the present disclosure relate to a propulsion system that allows for aerodynamic tuning of the thrust support structure. In that regard, reference is made to FIG. 48, which depicts a side and longitudinal cross-sectional view of one embodiment of a thrust support structure 7 including a straightening vane consistent with the present disclosure. In the illustrated embodiment, the thrust support structure 7 is surrounded by an airfoil-shaped straightening vane 89. As understood in the prior art, the angle of the airfoil relative to the approaching air affects its drag profile. Thus, if the airfoil-shaped straightening vane 89 is always perpendicular to the rotor shaft 1 and the rotor shaft 1 is tilted forward in forward flight, the aircraft movement direction may no longer be parallel to the airfoil-shaped straightening vane 89, resulting in unnecessary drag. One option to address this issue is to use an airfoil-shaped shell that includes the thrust support structure, along with bearings that allow the airfoil-shaped straightening vane 89 to rotate along the thrust support structure to maintain a low drag coefficient regardless of the aircraft's attitude or motion. In this way, the airfoil-shaped fairing 89 allows the thrust support structure 7 to "weathervane" into the prevailing airflow it encounters. Motors can be used to automatically move the airfoil to optimize its angle based on possible sensor inputs, or the motors can be manually controlled by the pilot to improve the aircraft's overall lift or create an aerodynamic braking effect as needed.

Thus, in an embodiment, the fairing 89 is configured to allow the aerodynamics of the thrust support structure to be streamlined with respect to the prevailing airflow moving around the thrust support structure 7. To that end, fairing support bearings 90 are disposed at each end of the fairing 89 to allow the fairing 89 to move freely with respect to the thrust support structure 7 and the engine 5. One or more positioning motors 91 may be included in the thrust support structure 7 and disposed relatively close to the rotor shaft 1. Such motors 91 may be energized to actively change the orientation (and thus the aerodynamics) of the fairing 89 and the thrust support structure 7. By providing suitable control of the position of the fairing 89 (and thus the thrust support structure 7), the structure may be aerodynamically streamlined with respect to the airflow moving around the thrust support structure 7, regardless of changes in the orientation of the thrust support structure with respect to the airflow.

The above description has often focused on an embodiment in which the thrust support structure 7 and rotor blades 3 are connected to the rotor shaft 1 with a clutch 6. Although the clutch 6 can provide a number of advantages as discussed above, its use is not required. For example, FIG. 49 illustrates an embodiment of a propulsion system in which the thrust support structure 7 and rotor blades 3 are connected to the rotor shaft 1 without a clutch 6 (e.g., rigidly connected). While this illustrated embodiment does not provide the advantages of the other described embodiments, it may be of interest in cases where such advantages are not desired. For example, such a configuration may be used for special purpose aircraft, such as unmanned aerial vehicles.

As previously described with respect to FIG. 42, the orientation of a propulsion system consistent with the present disclosure can transition from a vertical orientation to a horizontal orientation and vice versa. In such an embodiment, any suitable mechanism for orienting the propulsion system can be used. In that regard, reference is now made to FIG. 50, which illustrates an embodiment of an aircraft propulsion system in which a sliding rail system facilitates reorientation of all or a portion of the propulsion system between a vertical orientation and a horizontal orientation. In this embodiment, a frame rail guide 93 is disposed along the outside of the aircraft fuselage 4. Although the frame rail guide can be disposed in any suitable location, in an embodiment, it is positioned along the center of the fuselage 4. In addition, a rotor system cart housing 95 is coupled to the base of the aircraft propulsion system such that the rotor system cart housing is between the fuselage and the propulsion system. The frame rail guide 93 is configured to guide rollers 94 (or other guide mechanical elements) coupled to the rotor system car housing 95 as the propulsion system transitions between the vertical and horizontal orientations, as shown. The rollers 94 (or other guide mechanical elements) can be moved directly or indirectly (e.g., with electric or hydraulic motors) to move the rotor system cart housing 95 to a vertical or horizontal orientation depending on the desired thrust line for vertical lift or horizontal thrust for forward flight. The wings 28 and horizontal stabilizers 92 can be oriented to remain relatively perpendicular to the rotor blades to provide lift and stability during forward flight.

Various embodiments of a propulsion system including an engine 5 directly integrated into a thrust support structure have been described above. For example, FIG. 47 depicts an embodiment in which one or more jet engines are integrated into a thrust support structure 7 to provide thrust through ducting 10 to rotate the thrust support structure. While the thrust generated by such engines may be fixedly oriented (e.g., through fixed nozzles), in some instances it may be desirable to control the direction of the thrust exiting the thrust support structure 7. In that regard, reference is made to FIG. 51, which is a side view of one example of a propulsion system including a thrust support structure including one or more articulating nozzles that allow for directional thrust control. As shown, the articulating nozzles 96 are coupled to corresponding extremities of the thrust support structure 7 and are configured to receive thrust from the ducting 10. The nozzles 96 each include an outlet (not separately labeled) through which thrust flows. The nozzles 96 can be articulated (e.g., by one or more drive motors) to move or realign their respective outlets, thus allowing for direct control of the thrust line exiting the thrust support structure 7. 36 and 37 show another embodiment in which the engine 5 is directly integrated into the thrust support structure 7 and located just outside the clutch 6. The engine 5 powers a drive shaft 27 that is connected to a gear box 26 and drives one or more ducted fans or propellers 24. In the embodiment of Figs. 38 and 39, the engine 5 is directly integrated into the thrust support structure 7 that is directly connected to the rotor shaft 1. The engine 5 powers a drive shaft 27 that is connected to a gear box 26 to drive the ducted fans or propellers 24. The clutch 6 is integrated into the rotor hub 11, allowing the rotor blades 3 to rotate independently of the drive system components when the clutch is disengaged.

Much of this disclosure focuses on an embodiment in which a single set of rotor blades is utilized. Such a configuration is not required, and any suitable number of rotor blades may be used. For example, one, two, three, four or more sets of rotor blades may be utilized in the aircraft propulsion systems described herein. To illustrate the concept, refer to FIG. 52, which depicts one example of a propulsion system including multiple sets of rotor blades driven by a single thrust support structure consistent with this disclosure. In this embodiment, the thrust support structure 7 provides torque through the inner shaft 81. The torque is transferred by the inner shaft 81 through the clutch 6 to the outer shaft 80 to drive multiple sets of rotor blades 3, each connected to the outer shaft 80. The position of each set of rotor blades can be varied and set to achieve desired flight characteristics. In the embodiment of FIG. 52, for example, the two sets of rotor blades 3 can be mounted parallel to each other but offset from the same axis (e.g., about 1 degree to about 10 degrees), for example, to enhance aerodynamic and acoustic performance. Such an arrangement can also generate more lift by allowing the lower set of rotors 3 to catch the wake of the upper set of rotors 3. Additionally, both sets of rotors can be driven in the same direction to avoid wake interference problems created by counter-rotating blades. And the use of multiple sets of rotors 3 can allow for the generation of lift in an equivalent amount to a single set of rotor blades 3, but in a smaller footprint.

The techniques described herein may be implemented in a wide variety of aircraft designs and are not limited to use in relatively conventional helicopter designs such as those illustrated in many of the figures. Indeed, the techniques described herein have utility in any type of aircraft design that may benefit from their use. For example, the techniques of the present disclosure may be used in hybrid aircraft such as compound helicopters, compound gyroplanes, and the like, both of which may be referred to as heliplanes. In such aircraft lift, a helicopter-like rotor system may be used during takeoff and landing, with a secondary propulsion system (e.g., push or pull propellers, one or more jet engines, one or more turbofan engines, etc.) providing horizontal (forward and/or aft) thrust during flight. Lift may also be generated by wings during full-body/reverse flight, reducing the speed of helicopter-like rotor blades.

Figure 53 depicts one example of the use of the techniques of the present disclosure in a compound helicopter environment. In the illustrated embodiment, the aircraft is a compound helicopter including multiple sets of coaxial rotor blades 3 driven by a thrust support structure 7 (including engines 5 and ducting 10). The aircraft further includes thrusters 97 to provide forward thrust during flight. In this embodiment, thruster 97 is a single push propeller, although it should be understood that any suitable type and number of thrusters may be used. For example, thruster 97 may be one or more jet engines, turbofan engines, rocket engines, fans, propellers, combinations thereof, and the like.

In operation, the coaxial rotor blades 3 can be used for lift, and the thrusters 97 can be used to accelerate, decelerate, or hold the aircraft stationary depending on the desired flight conditions. Lift can also be provided by the wings 28 when the aircraft is in forward flight. In such an embodiment, the speed of the rotors 3 can be reduced, for example, by slowing down the engines 5 and/or temporarily shutting them down, since there is a reduced need for lift provided directly by the rotors 3 in forward flight. In such an example, the motors (not shown) used in hovering for azimuth control may be retasked to provide constant low power to the rotor shaft 1 (not labeled in FIG. 53) to maintain stability and control. Any yaw motion (if any) occurring in such a condition can be counteracted using stabilizers (e.g., vertical stabilizers, not shown) and reduced accordingly.

Figure 54 illustrates another example of the use of the techniques of the present disclosure in a composite helicopter environment. This example is substantially similar to the embodiment of Figure 53, except that a single set of rotor blades 3 is used instead of the multiple coaxial rotor blades used in the embodiment of Figure 53. As the nature and function of the machine elements of Figure 54 are the same as those shown and described above with respect to Figure 53, they will not be repeated here for the sake of brevity.

Many of the above-described embodiments utilize a propulsion system including a clutch 6, a thrust support structure 7, and one or more sets of rotor blades 3, where the thrust support structure 7 and rotor blades 3 are coupled to one another and vertically off-center from one another along the rotor shaft 1. Such a configuration is not required, and the present disclosure includes embodiments in which the clutch 6 is not used and/or the rotor blades 3 and thrust support structure are coupled to the rotor shaft 1 and positioned in the same plane. One example of this concept is shown in FIG. 55, which depicts a top and side view of one example of a propulsion system consistent with the present disclosure, where the propulsion system does not include a clutch 6, but includes thrust support structures 7 and rotor blades 3 that are in the same plane. More specifically, in this embodiment, the thrust support structure 7 and rotor blades 3 are integrated or coupled to a common structural hub 98, which is in turn integrated or coupled to the rotor shaft 1. The common structural hub 98 is integral with or fixedly attached to the rotor shaft 1 by any suitable means, such as through one or more fasteners, mechanical interference joints, welding, adhesives, combinations thereof, and the like. More specifically, the common structural hub is integral with or coupled to the rotor shaft 1 so that its position remains fixed relative to the position of the rotor shaft 1 (i.e., the common structural hub 98 does not rotate independently of the rotor shaft 1). Similarly, the thrust support structure 7 and the rotor blades 3 may be integral with or coupled to the common structure 98 by any suitable means, such as through one or more fasteners, mechanical interference joints, welding, adhesives, combinations thereof, and the like. During operation, the engine 5 rotates the thrust support structure 7. The resulting torque is transferred from the thrust support structure 7 to the common structural hub 98 and then to the rotor shaft 1 to rotate the rotor shaft 1. Because the common structural hub 98 is fixedly attached to the rotor shaft 1, the rotation of the rotor shaft 1 ultimately rotates the rotor blades.

Various embodiments have been described above in which jet or other engines are coupled to or integral with the thrust support structure and are operable to generate thrust in the form of airflow directed through ducting to rotate the thrust support structure. While such embodiments are useful, the use of ducted airflow through the thrust support structure is not required and other configurations may be used. For example, one or more engines coupled to or integral with the thrust support structure may be coupled to one or more drive shafts, which may in turn be coupled to one or more fans. In operation, the engines may drive the fans on the drive shafts to generate airflow that may be discharged in a fixed or variable direction to rotate the thrust support structure.

FIG. 56 depicts an embodiment of the present disclosure in which one or more fans are driven by a motor to generate an airflow that can be directed in a fixed or variable direction. More specifically, FIG. 56 depicts an embodiment in which multiple engines 5 are coupled to or integrated with the proximal end of the thrust support structure 7 and are located in close proximity to the rotor hub 1. A first (proximal) end of a drive shaft 27 is coupled to each engine 5. Each drive shaft 27 may be located outside the thrust support structure 7 or may pass through a channel in the thrust support structure 7. In either case, a second (distal) end of each drive shaft 27 is coupled to a fan unit 99 at the distal end of the thrust support structure 7. In operation, the engine 5 causes the drive shaft 27 to generate an airflow in the fan unit. Each fan unit 99 is oriented to create an airflow that is in the same or substantially the same plane as the thrust support structure 7. By doing so, it is possible to even facilitate the load on the fan blades and fan blade hubs (not shown) of each fan unit while the thrust support structure 7 is moving.

In this embodiment, an airflow guide vane nozzle 100 is coupled to each fan unit 99 and configured to direct the airflow generated by each fan unit 99. In the illustrated embodiment, the airflow guide vane nozzle 100 is a static nozzle configured to redirect the airflow from the fan unit 99 to a desired exit angle, in this case approximately 90 degrees relative to the plane of the thrust support structure 7. The illustrated configuration is by way of example only, and the guide vane nozzle 100 can be configured to redirect the airflow generated by the fan unit 99 to any desired exit angle. And, in an embodiment, the guide vane nozzle 100 and/or its exit can be articulated (e.g., by one or more motors, not shown), thereby allowing dynamic control of the airflow exit angle. Control of the exit angle can be performed to redirect the airflow to a desired angle to generate thrust for moving the thrust support structure 7 and the clutch 6. For example, the use of guide vane nozzles 100 may allow the airflow generated by the fan unit 99 to be redirected from 0 degrees to 180 degrees or more relative to the plane of the thrust support structure.

One advantage of positioning the engine 5 in close proximity to the rotor shaft 1 and fan unit 99 near the distal end of the thrust support structure 7 is that centripetal forces applied while the thrust support structure 7 is moving can be partially offset by blade loads created from the generation of thrust by the fan unit 99. Such blade loads can pull the fan blades of the fan unit 99 back toward the drive shaft 27 and the engine 5. Furthermore, because the drive shaft 27 can be much thinner than the thrust support structure 7, the configuration of FIG. 56 can enable a thinner intermediate portion of the thrust support structure 7 to be used relative to the portion of the thrust support structure 7 that includes the piping 10.

FIG. 57 illustrates another embodiment of the present disclosure utilizing one or more fan units located at or proximate to the distal end of the thrust support structure 7 to generate thrust and rotate the thrust support structure. In this embodiment, the fan units 99 are in the form of centrifugal blowers 101. Like the fan units 99, the centrifugal blowers 101 are driven by a drive shaft 27, which in turn is driven by one or more engines 5. In operation, each centrifugal blower 101 is driven by the drive shaft 27 to create a torque to drive the thrust support structure 7, which in turn creates an airflow that rotates the rotor blades 3. The orientation of the centrifugal blowers 101 can be set to achieve desired aerodynamic conditions. For example, and as shown in FIG. 57, the centrifugal blowers 101 may be relatively long along a first axis, but relatively thin along a second axis that is perpendicular to the first axis. In such an example, the centrifugal blowers 101 can be oriented such that their long dimension (along the first axis) is perpendicular (e.g., perpendicular or substantially perpendicular) to the axis extending along the thrust support structure 7. In contrast, the short dimension of the centrifugal blowers 101 can be oriented such that it is parallel or substantially parallel to the axis extending along the thrust support structure 7. In such a configuration, the relatively thin dimension of the centrifugal blowers 101 is presented in the direction of rotation of the thrust support structure 7, which can reduce or minimize the induced drag created by the centrifugal blowers 101 as the thrust support structure 7 rotates.

As with FIG. 56, the embodiment of FIG. 57 allows for the use of a drive shaft 27 and allows for the use of a lower profile thrust support structure 7 relative to the shape of the thrust support structure including the piping 10. And as with the guide vane nozzle 100, the orientation of the centrifugal blowers 101 can be fixed or articulated. In the latter example, the articulation of the centrifugal blowers 101 can allow for dynamic control of the exit angle of the airflow they generate.

FIG. 41 depicts another embodiment according to the present disclosure, in which concentric shafts are used to support and separate various components of an aircraft. As shown, the outer shaft 80 is connected directly to the rotor blades 3, supported by support bearings 2, and separated from the fuselage. The inner shaft 81 is disposed within the outer shaft 80 and connected directly to the thrust support structure 7, which in turn is connected to the engine 5. The inner shaft 81 is supported and separated from the outer shaft 80 by a concentric shaft bearing 79, which is disposed between the inner surface of the outer shaft 80 and the outer surface of the inner shaft 81. The clutch 6 is configured to engage or separate the inner shaft 81 and/or the outer shaft 80, and can be disposed in any suitable location. For example, and as shown in FIG. 41, the clutch 6 can be disposed within the outer shaft 80 (i.e., between the outer shaft 80 and the inner shaft 81) and configured to engage and separate the inner shaft 81. FIG. 41 illustrates an embodiment in which the thrust support structure 7 is connected to the inner shaft 81 and the rotor blades 3 are connected to the outer shaft 80, but it should be understood that such a configuration is not required and the thrust support structure and rotor blades 3 can be connected to the inner shaft 81 or the outer shaft 80, respectively.

FIG. 58 is a side view of another embodiment of the present disclosure in which separate shafts are utilized for different components of an aircraft propulsion system, such as the control system, rotor system, and drive system. The nature and function of many of the machine elements of FIG. 58 are identical to those described above in connection with other embodiments and will not be repeated for the sake of brevity. In this embodiment, the illustrated propulsion system includes an outer shaft 80, an inner shaft 81, and an intermediate shaft 102. The intermediate shaft 102 is disposed between the outer shaft 80 and the inner shaft 81. The outer shaft 80 is fixed (e.g., by a shaft mount 103) to the fuselage structural member 56 so that it does not rotate. The pitch control rod 41 connects to the swash plate 40, which is linked to the rotor blades 3 to provide rotor control. The intermediate shaft 102 is separated from the outer shaft 80 by a concentric bearing 79. The rotor hub 11 and the rotor blades 3 are directly connected to and supported by the intermediate shaft 102. The inner shaft 81 is separated from the intermediate shaft 102 by a concentric shaft bearing 79 and is only connected to the intermediate shaft 102 by a clutch 6 when torque from the thrust support structure 7 and the engine 5 driving the fan unit 99 generate thrust, causing the inner shaft 81 and the intermediate shaft 102 to rotate simultaneously. This arrangement provides a strong support system for the various components of the propulsion system, allowing direct replacement or service of such components. This also allows the components to be upgraded independently and modularly as aircraft upgrades or missions change. The fuel tanks 33 and the battery 74 are connected to the engine 5 through connections that run along the lower rotating union 16 and the interior of the inner shaft 81. Other components shown in FIG. 58 are described in other embodiments and are shown merely to illustrate a complete embodiment in which multiple shafts may be used. Of course, multiple shafts may be used in other embodiments as well. One advantage of using multiple shafts is that it allows the components of the propulsion system to be modular and upgradeable independently or as a unit.

In many of the above embodiments, the energy sources, such as fuel tanks and batteries, are located within the fuselage of the aircraft and are connected to components that are external to the fuselage, such as fuel lines, wires, etc. Such a configuration is not required, and the energy sources, such as batteries, fuel tanks, etc., may be located elsewhere. In that regard, reference is made to FIG. 59, which illustrates a side view of one embodiment of an aircraft propulsion system in which the energy sources, such as fuel tanks and batteries, are integrated into or connected to the thrust support structure. The rotor blades 3 and swash plate 40, which are connected to the outer shaft 80, are supported by a separate set of support bearings 2, which are connected to the fuselage structural member 56. The fuel tanks 33, batteries 74, and engines 5 are all mounted to the thrust support structure 7, eliminating the need to route power lines, fuel lines, etc. through the inner shaft 81. In addition, the wireless control/telemetry module 105 allows communication and data signaling to the engines 5. Thus, the inner shaft 81 can be serviced without the need to remove or move fuel lines, power cables, control lines, etc. that might otherwise be routed therethrough. For example, the inner shaft 81 can be removed from the fuselage 4 by isolating the connection at the support bearing 2, the concentric shaft bearing 79, and the clutch 6. Such a configuration can also eliminate the need for shaft seals and/or slip rings that can be used to route fuel lines, control lines, power lines, and the like between the shafts. FIG. 60 is a schematic side cross-sectional view of another embodiment of an aircraft including a propulsion system consistent with the present disclosure. In this embodiment, the rotor shaft 1 extends into the fuselage 4 and is supported by one or more support bearings 2 in the bearing support structure 43 and the thrust bearing assembly 57. In this embodiment, the support bearings 2 are disposed in the bearing housing 107 and include ball bearings between the inner and outer race cages. In either case, the support bearings 2 of the inner and outer races are configured to support the load of the rotor shaft 1.

The support bearing 2 is coupled to the bearing support structure 43 by any suitable means, for example, by a bearing housing 107 that is attached (e.g., by welding, mechanical fasteners, etc.) to the bearing support structure 43. In an embodiment, the bearing support structure 43 is in the form of a hollow tubular frame formed from any suitable material, such as a metal, alloy, composite, or other suitable material, having sufficient strength and properties to support the loads transferred from the rotor shaft 1 to the support bearing 2. The bearing support structure 43 is attached to the fuselage 4 in any suitable manner (e.g., by welding, mechanical fasteners, etc.).

In an embodiment, the outer race of the support bearing 2 is coupled to the bearing support structure 43 via the bearing housing 107, and the inner race of the support bearing 2 is coupled to the outer surface of the rotor shaft 1 (e.g., via mechanical fasteners, welding, or other suitable fastening means). In such an embodiment, the rotational torque caused by the rotation of the rotor shaft 1 does not exceed the ability of the support bearing 2 to isolate the fuselage 4 from such torque.

The embodiment of FIG. 60 further includes a rotating magnet 17 and a stator 18. The rotating magnet 17 is disposed on a disk-shaped structure that is connected to the rotor shaft 1 at its hub, so that the disk-shaped structure and the rotor shaft can rotate together simultaneously. The stator 18 in this embodiment is in the form of a pair of (e.g., metal) plates and metal (e.g., copper) coils that are disposed above and below the rotating magnet 17 and are connected to the fuselage 4 by exposing the support structure 43. Of course, the stator 18 is not limited to the illustrated configuration and can include one or more pairs of plates and coils depending on the amount of azimuth control required for the aircraft.

In this embodiment, the rotor blades 3 are coupled to a rotor hub 11, which is in turn coupled to the rotor shaft 1. Thus, rotation of the rotor shaft 1 rotates the rotor blades 3. The exemplary embodiment also includes a lower shaft flange collar 108 and an upper shaft flange collar 109, which are coupled to the rotor shaft 1 near its upper end. In this embodiment, the lower shaft flange collar 108 and the upper shaft flange collar 109 are above the rotor blades 3 and provide attachment points for the clutch 6 to the rotor shaft 1.

The clutch 6 is configured to engage/disengage the rotor shaft 1 from the thrust support structure 7. As noted above, the clutch 6 can be a one-way bearing or other similar device capable of engaging and disengaging two rotating structural members. The inner race of the clutch 6 is coupled, for example, via mechanical fasteners or through other means, between the lower shaft flange collar 108 and the upper shaft flange collar 109 so that it is attached to the upper frame rail flange 110 and the lower frame rail flange 116 and can be rotated by the rotor shaft 1.

60, the thrust support structure 7 includes an upper frame rail flange 100, a lower frame rail flange 116, a frame rail 112, a cross member 111, an engine mount 113, an engine 5, a piping flange 115, and a piping 10. The centerline of the thrust support structure 7 extends perpendicular to the rotor shaft 1 and is 180 degrees apart. The upper frame rail flange 110 and the lower frame rail flange 116 may be made of any suitable material, such as one or more metals, alloys, composites, etc. The upper frame rail flange 110 and the lower frame rail flange 116 are coupled to the outer race of the clutch 6 in any suitable manner, such as via welding, mechanical fasteners, etc.

In general, the frame rails 112 provide structural support and are connected to engine mounts 113, which are secured to the frame rails 113 in any suitable manner. In an embodiment, the frame rails 112 are formed from one or more structural rails including any suitable number of reinforcing cross members 111. The engine 5 is mounted longitudinally between the frame rails 112. For example, the engine 5 can include fastener mounting locations on both sides thereof (e.g., near their intake and exhaust) and can be mounted between the frame rails 112 via such fastener mounting locations. The frame rails 112 are also connected to piping flanges 115. The piping flanges 115 surround the exhaust passages of the engines 5 and allow their thrust to be transferred to the piping 10. The piping 10 can be a simple round, oval, or streamlined hollow tube made of metal or synthetic material or other material sufficient to draw air flow and maintain structural integrity against thermal and centripetal forces. In an embodiment, the piping 10 extends from the piping flange 115 near the exhaust passage of the engine 5 to an appropriate length to allow the high velocity airflow emanating from the engine 5 to provide thrust at the outer end of the piping 10, generating the rotational torque necessary to power the rotor blades 3 and provide sufficient lift for flight.

Rotation of a thrust support structure consistent with the present disclosure can provide lift within itself that can be utilized to facilitate flight of an aircraft and reduce or even eliminate the need for rotor blades separate from the thrust support structure. In other words, a thrust support structure consistent with the present disclosure can be configured to serve as a rotor blade for an aircraft. Such a configuration can provide various advantages, including, but not limited to, eliminating, reducing, or optimizing drag associated with structures above an aircraft fuselage. For example, when a thrust support structure serves as a rotor blade for an aircraft, the need for a separate rotor blade can be reduced or even eliminated, limiting or eliminating drag and other inefficiencies that result from the use of separate rotor blades. The elimination of separate rotor blades can also reduce or eliminate the need for a clutch to engage and separate the rotor blades from the rotor shaft, avoiding mechanical complexities and efficiency considerations associated with the use of a clutch. Thus, embodiments of the present disclosure relate to a propulsion system for an aircraft, in which a thrust support structure is configured to serve as a rotor blade.

An example of the concept is shown in Figures 61-64, which represent various embodiments of a propulsion system consistent with the present disclosure, including a thrust support structure 7 configured to function as a rotor blade. In such an embodiment, the propulsion system includes a rotor hub 11 or a common structural hub 98, as previously described. The engine 5 is coupled to the rotor hub 11 or the common structural hub 98, such that a proximal end is located near (i.e., in close proximity to) the rotor hub 11 or the common structural hub 98. It should therefore be understood that at least a portion of the engine 5 is fully or partially supported by the common structural hub 98/rotor hub 11. The engine 5 is also coupled to the piping 10. Alternatively, the engine 5 may be integrated with (e.g., housed within) the rotor hub 11 or the common structural hub 98.

The tubing 10 includes one or more channels 10' therein. As shown, the proximal end of the tubing 10 is coupled to the distal engine 5. For purposes of illustration, the embodiment of Figures 61-64 shows the engine 5 positioned flush with the tubing 10 and the common structural hub 98/rotor hub 11, such that the proximal end of the tubing 10 is directly coupled to the distal end of the engine 5. That configuration is not required, and the engine 5 and tubing 10 can be positioned in any suitable manner. For example, the engine 5 can be off-center above, below, or to the side of the rotor hub 11/common structural hub 98, and the tubing 10 can be coupled by a connecting member that includes one or more channels therein that can be fluidly coupled to the channel 10'.

More generally, the engine 5 can be coupled to the piping 10 in any suitable manner. For example, and as shown in FIGS. 61-64, the engine 5 can be coupled to the piping 10 via a coupler 120, as shown. In general, the coupler 120 functions to fluidly couple the piping 10 to the engine 5, for example, via a rotary joint, a gas seal, or a blade grip. In any case, the coupler 120 can include at least one connected channel (not shown) configured to fluidly couple with the engine 5 and the channel 10' of the piping 10. Thus, thrust generated by the engine 5 can flow through the connected channel of the coupler 120 into the channel 10'. In an embodiment, the coupler 120 allows thrust generated by the engine to be transferred to the channel 10' while still allowing the piping 10 to pivot about an axis, for example, in the same manner as an articulating rotor blade. In that regard, the coupler 120 can be in the form of a rotatable joint, a rotatable gas seal, or a rotatable blade grip. In such an example, the orientation of the piping 10 can be controlled via a control arm 41 or a rotating swash plate 40, as shown.

The tubing 10 can have any external shape capable of generating lift as the thrust support structure rotates. For simplicity and ease of explanation, the tubing 10 is illustrated in FIG. 61 in the form of a hollow ellipsoid, but any other suitable shape can be used, as will be understood by those skilled in the art. In an embodiment, the tubing 10 is in the form of an airfoil having one or more channels 10' therein, but any known or later developed rotor blade shape can be used. In an embodiment, the external shape of the tubing 10 is configured such that the portion generates the majority of the lift as the thrust support structure 7 rotates about the axis of the rotor hub 11/common structural hub 98. For example, the external shape of the tubing 10 can be configured such that the outer third of the tubing 10 provides the majority of the lift. In such an embodiment, the external shape of the tubing can be the same as or similar to the shape of the rotor blades used in the tip-jet-powered Sud-Ouest Djinn helicopter.

In either case, the ducting 10 can be fluidly coupled to the engine 5 such that thrust (e.g., in the form of exhaust, compressor bleed air, etc.) is directed through the channel 10' toward the distal end of the ducting 10. The thrust (airflow) can be directed out of the channel 10' in such a manner as to rotate the thrust support structure. For example, and as shown in FIG. 62 (and consistent with the description of FIG. 56 above), the propulsion system can be configured to generate an airflow that is coplanar or substantially coplanar with the thrust support structure 7. To that end, and as shown in FIG. 62, in an embodiment, the propulsion system includes an airflow guide vane nozzle 100 coupled to the distal end of the ducting 10. Consistent with the description of FIG. 56, the airflow guide vane nozzle 100 is configured to direct the airflow produced by the engine 5 and flowing through the channel 10'. In the embodiment of FIG. 62, the airflow guide vane nozzle 100 is a static nozzle configured to redirect the airflow from the duct 10' to a desired exit angle, e.g., about 90 degrees relative to the plane of the thrust support structure 7. The illustrated configuration is by way of example only, and the guide vane nozzle 100 can be configured to redirect the airflow generated by the engine 5 to any desired exit angle. In an embodiment, the guide vane nozzle 100 and/or its exit can be articulated (e.g., by one or more motors, not shown), thereby allowing dynamic control of the airflow exit angle. Control of the exit angle can be performed to redirect the airflow to a desired angle to generate thrust for moving the thrust support structure 7 and clutch 6 (when used). For example, the use of the guide vane nozzle 100 can allow the airflow generated by the engine 5 to be redirected from about 0 degrees to about 180 degrees or more relative to the plane of the thrust support structure 7.

63 is a side view of one example of a propulsion system including a thrust support structure including one or more articulating nozzles that function as rotor blades to provide directional thrust control. As shown, the articulating nozzles 96 are configured to couple to corresponding ends of the piping 10 and receive thrust from the channel 10'. The nozzles 96 each include an outlet (not separately labeled) through which thrust flows. The nozzles 96 can be articulated (e.g., by one or more drive motors) to move or realign their respective outlets, thus allowing direct control of the thrust line (thrust) of air exiting the thrust support structure 7. Other components shown in FIG. 63 have been described in other embodiments and will not be described again here for the sake of brevity.

64 depicts another embodiment of a propulsion system including a thrust support structure functioning as a rotor blade consistent with the present disclosure. The nature and function of many of the components of FIG. 64 are not depicted for the sake of brevity, as they have been described above in connection with FIG. 60. However, unlike the embodiment of FIG. 60, the embodiment of FIG. 64 does not have separate rotor blades 3, but instead utilizes tubing 10 (or more generally, thrust support structure 7) as rotor blades. The embodiments of FIGS. 63 and 64 also differ from the embodiments of FIGS. 61 and 62 in that they include a clutch 6. The description of the use of the clutch 6 is provided to illustrate its use in the context of a thrust support structure functioning as a rotor blade, but is not required to be as shown in FIGS. 61 and 62.

Another embodiment of a propulsion system consistent with the present disclosure includes one or more thrust support structures, each of which serves as a rotor blade, where airflow from one or more fuselage-mounted engines is redirected through the rotor shaft and through ducts inside one or more of the rotor blades. Such a system can be used with or without the thrust support structure including one or more engines connected to ducting, as shown in FIGS. 63 and 64 and described above. In either case, one or more thrust diverters can be located in the exhaust flow of the fuselage-mounted engines and configured to control the flow of the airflow generated thereby. More specifically, the thrust diverters are oriented (pointed) toward the rotor shaft and configured to control the amount of airflow generated by the fuselage-mounted engines that is distributed to the ducting inside the one or more rotor blades.

During vertical flight, all or a significant portion of the airflow generated by the fuselage-mounted engines may be directed by the thrust diverter to the rotor shaft and then directed (e.g., by a plenum) to one or more rotor blades. The airflow thus directed may exit the rotor blades through openings proximate their tips, accelerating or decelerating the rotor's rotating blades to generate a desired amount of lift. During forward flight, the thrust diverter may be activated to allow a substantial amount of the airflow to exit directly aft of the thrust diverter (i.e., undiverted to the rotor shaft), thus increasing the aircraft's speed. Notably, during forward flight, less rotor power is required for lift, and a majority of the engine thrust may be utilized for forward propulsion.

To illustrate this concept, refer to FIGS. 65A and 65B, which depict one example of an aircraft propulsion system consistent with the present disclosure, where all or a portion of the airflow from one or more fuselage-mounted engines is directed into a duct in a pair of counter-rotating rotor blades. Such a system may be referred to herein as a reaction drive system. In the embodiment of FIGS. 65A and 65B, the engines 5 may be embedded in the aircraft fuselage 4 or may be externally mounted. In either case, the thrust diverter 117 may be located downstream of the engines 5 and in a diverter manifold 118. In general, the thrust diverter 117 is configured to redirect the airflow generated by the engines 5. The position of the thrust diverter 117 may affect the relative amount of airflow generated by the engines 5 that is directed into the rotor blades 3 to create lift and/or past the diverter 117 to provide thrust. The relative amount of airflow directed at or past the rotor blades 3 can directly affect the flight characteristics of the aircraft, including vertical lift production, forward speed, or a combination thereof.

In such an embodiment, the airflow ducts 29 in the fuselage are configured to allow air to be diverted to the rotor shaft 1. Airflow entering the rotor shaft 1 is routed to the plenum 14 and then distributed by the plenum 14 to the ducts in the rotor blades 3. In an embodiment, the plenum 14 is configured to distribute such airflow evenly. However, such a configuration is not required and the plenum 14 may be individually controllable to allow finer control of the distribution of airflow to the individual rotor blades. The airflow entering the rotor blades 3 may exit the rotor blades 3 near their tips, generating thrust to rotate the rotor blades 3 as described above. When the aircraft is engaged in vertical flight, the system may be configured as shown in FIG. 65(A) where the bulk airflow from the engines 5 is directed to the rotor blades 3 to generate vertical lift. When the aircraft transitions to forward flight, the thrust diverter 117 may be actuated to reduce the amount of airflow directed toward the rotor blades 3 and increase the amount of airflow directed toward the rear of the aircraft, as shown in the embodiment of FIG. 65(B). In other words, as the aircraft transitions to forward flight, the amount of airflow directed toward the rotor blades 3 may decrease, while the amount of airflow directed toward the rear may increase. In such a case, the amount of airflow directed toward the rotor blades 3 may be adjusted (by actuation of the thrust diverter 117) to maintain the lift, stability, and controllability of the aircraft.

66 illustrates another embodiment of an aircraft consistent with the present disclosure. Similar to the aircraft of FIGS. 65A and 65B, the aircraft of FIG. 66 utilizes airflow from an engine to rotate a thrust support structure (and thus one or more rotor blades). For example, in the embodiment of FIG. 66, a stationary (i.e., non-rotating) rotor shaft 1 is coupled at a proximal end to one or more engines 5 and includes one or more channels (shaft ducts) for receiving the airflow. A clutch 6 (e.g., a sealing clutch) couples the rotor shaft 1 near its distal end. An outer shaft 80 is disposed about the rotor shaft 1 and is located between the clutch 6 and the proximal end of the rotor shaft 1.

In the engaged state, the clutch 6 is coupled to the outer shaft 80, while in the disengaged state, the clutch 6 is disengaged from the outer shaft 80. As further shown, the thrust support structure 7 is coupled to the rotor shaft 1 via the clutch 6. Like the rotor shaft 1, the thrust support structure 7 in this embodiment includes at least one channel (thrust duct) for receiving airflow. In addition, the thrust support structure includes at least one outlet (thrust outlet, not labeled) for directing airflow at an angle (e.g., perpendicular) to the clutch 6 and/or the upward extension of the rotor shaft 1. The clutch 6 may be any suitable clutch capable of releasably engaging (and disengaging) the outer shaft 80, such as, but not limited to, the clutch illustrated and described above in connection with FIGS. 19 and 20. Of course, other clutch types and configurations may be used.

For ease of understanding, the operation of the aircraft of FIG. 66 will now be described, with the clutch 6 having the same or similar structure as the clutch described above in conjunction with FIGS. 19 and 20. In such an example, the operation of the engine 5 generates an airflow (indicated by the arrows) directed to an inlet opening in the rotor shaft 1. The airflow passing through the inlet opening flows into one or more rotor ducts of the rotor shaft 1 toward its distal end. Proximate the distal end of the rotor shaft 1, at least a portion of the airflow flows from the rotor ducts to one or more inlet openings in the thrust support structure 7. In the illustrated embodiment, the thrust support structure 7 includes at least two inlet openings, each leading to a corresponding thrust channel in its opposite side.

Airflow passing through the inlet openings flows into corresponding thrust channels toward opposite distal ends of the thrust support structure 7. The airflow in the thrust channels ultimately exits the thrust support structure 7 through one or more outlet openings, as shown. In the illustrated embodiment, the thrust support structure 7 includes a single outlet opening at each of its two distal ends, but it should be understood that any suitable number of outlet openings (e.g., 2, 3, 4, 5, 10, etc.) can be used and such openings can be located at any suitable location or combination of locations along the thrust support structure 7. Furthermore, the thrust support structure 7 is not limited to two distal ends, but can include any suitable number of distal ends (e.g., 3, 4, 5, 10, etc.).

In either case, the outlet opening can be configured to direct the airflow so that it exits the thrust support structure 7 at a desired exit angle. In an embodiment, the outlet opening is configured to direct the airflow so that the exit angle is perpendicular or substantially perpendicular to the upwardly extended portion of the clutch 6 and/or rotor shaft 1. Alternatively, or in addition, in an embodiment, the outlet opening is configured so that the airflow exits the thrust support structure 7 at an angle ranging from 0 degrees to 180 degrees, such as from 0 degrees to 120 degrees, or even from 0 degrees to 90 degrees, relative to the upwardly extended position of the clutch 6 and/or rotor shaft 1. In an embodiment, the exit angle can be fixed, while in other embodiments the exit angle can be variable. In the latter example, the outlet opening can be in the form of or include an articulating nozzle that can be moved so that the airflow exits the thrust support structure 7 at a desired angle, as previously described in connection with Figures 51, 56, 62, and 63.

In either case, the airflow through at least one outlet opening rotates the thrust support structure 7, as shown by the curved arrow in FIG. 66. The torque generated by the rotation of the thrust support structure 7 is transmitted to the clutch 6, engaging the clutch 6 with the outer shaft 80. For example, the torque applied to the clutch 6 can cause the clutch mobile bearing 52 to climb along the incline of the clutch outer case until the clutch mobile bearing 52 engages with the surface of the incline and with the surface of the clutch inner race 114 (which is connected to the outer shaft 80 in a manner similar to the engagement of the inner race 114 with the rotor shaft 1 in the context of the previous embodiment - see FIGS. 19 and 20). When the clutch 6 is so engaged, the torque applied by the rotation of the thrust support structure 7 is transmitted to the outer shaft 80 through the clutch 6 (or, more specifically, through the inner race 114), rotating the outer shaft 80 as shown in FIG. 66. The rotation of the outer shaft 80 in turn causes the rotor blades 3 to rotate. As described above in connection with other embodiments, the support bearings 2 function to decouple the torque generated by the rotation of the rotor blades 3 and outer shaft 80 from the fuselage 4.

When the rotational momentum of the clutch 6 (or, more specifically, the outer case 58) is absent or reduced, the clutch outer case 58 begins to decelerate in its relationship to the outer shaft 80. The clutch moving bearings 52, due to their momentum and centrifugal force, move away from and loosen their grip on the clutch inner race 114 (which, again, is attached to the outer shaft 80), thus disengaging the outer shaft 80 from the clutch outer case 58 and allowing the outer shaft 80 to rotate freely with respect to the latch 6. Also, the clutch 6 is disengaged from the outer shaft 80 when the speed of the outer shaft 80 is greater than the rotational speed of the clutch 6. The disengaged state of the clutch 6 allows the outer shaft 80 to rotate freely and to self-rotate without any induced drag from the thrust support structure 7 and the power system. Also, during self-rotation, the preserved inertia of the power system is transferred to the outer shaft 80 at any time when the speed of the outer shaft 80 is equal to or less than the rotational speed of the clutch 6. In that condition, the transfer of power system inertia can provide more time to reduce the overall pitch to the safe angle required for the aircraft to autorotate.

The following examples are additional non-limiting embodiments of the present disclosure.

Example 1: According to this example, an aircraft is provided that includes a fuselage, a support bearing connected to the fuselage, a rotor shaft attached to the fuselage by the support bearing, the rotor shaft being capable of rotating about a first axis extending through the rotor shaft, the support bearings allowing the rotor shaft to rotate relative to the fuselage about the first axis at all times, a rotor shaft, rotor blades connected to the rotor shaft, the rotation of the rotor shaft causing the rotor blades to rotate about the first axis, a thrust support structure connected to the rotor shaft, and an engine connected to the fuselage or the thrust support structure, in operation, the engine causes the thrust support structure to rotate about the first axis, the rotation of the thrust support structure causes the rotor shaft to rotate about the first axis, which in turn causes the rotor blades to rotate about the first axis, and the support bearings isolate the fuselage from at least the torque generated by the rotation of the rotor blades.

Example 2: This example includes any or all of the features of Example 1, and further includes a clutch coupled to the rotor shaft or a rotor hub coupled to the rotor shaft, the clutch having an engaged state and a disengaged state.

Example 3: This example includes any or all of the features of Example 2, and during operation, the engine causes the thrust support structure to rotate about the first axis at a first rotational speed that is greater than or equal to the second rotational speed of the rotor shaft, causing the clutch to enter the engaged state and grip the rotor shaft.

Example 4: This example includes any or all of the features of Example 3, and the clutch enters or is in the disengaged state when the first rotational speed is less than the second rotational speed.

Example 5: This example includes any or all of the features of Example 2, and further includes the clutch including a clutch outer case, a clutch inner race connected to the rotor shaft, and a clutch movable bearing, the clutch outer case including a ramp and a cavity, and in the engaged state, the clutch movable bearing is disposed between and in contact with both the ramp and the clutch inner race, thereby engaging the clutch outer case and the clutch inner race, and in the disengaged state, the clutch movable bearing is disposed within the cavity and only in contact with the clutch outer case.

Example 6: This example includes any or all of the features of Example 2, and the clutch is coupled to the rotor shaft.

Example 7: This example includes any or all of the features of Example 2, and the clutch is coupled to the rotor hub.

Example 8: This example includes any or all of the features of any of examples 1-7, wherein the support bearing includes a bearing housing having inner and outer race cages and at least one bearing between the inner and outer race cages, and the rotor shaft extends into the fuselage and into the support bearing.

Example 9: This example includes any or all of the features of example 8, and further includes a bearing support structure, the support bearing is coupled to the bearing support structure, and the bearing support structure is coupled to the fuselage.

Example 10: This example includes any or all of the features of Example 9, and the rotor shaft is connected to the fuselage only by the support bearing and the bearing support structure.

Example 11: This example includes any or all of the features of any of Examples 1 to 10, and the engine is coupled to the thrust support structure.

Example 12: This example includes any or all of the features of example 11, and the thrust support structure includes a proximal portion coupled to the rotor shaft and a distal portion coupled to the engine.

Example 13: This example includes any or all of the features of any one of examples 1 to 12, wherein the thrust in the form of an airflow generated by the engine is directed generally along thrust lines that extend horizontally above, below, or above and below the rotor blades.

Example 14: This example includes any or all of the features of example 13, and the airflow inhibits the formation of vortices above the rotor blades.

Example 15: This example includes any or all of the features of any one of Examples 1 to 14, and further includes a heading control device.

Example 16: This example includes any or all of the features of example 15, and the orientation control device includes an electric reversible motor including a motor shaft, the motor shaft being the rotor shaft.

Example 17: This example includes any or all of the features of example 16, wherein the orientation control device includes a rotor and a stator, and when the stator is energized, a force is applied to the rotating magnet, which in turn provides a force to the fuselage that rotates the fuselage about the first axis.

Example 18: This example includes any or all of the features of example 17, wherein the orientation control device includes a spacer coupled to the rotor shaft, the rotating magnet is coupled to the spacer, and the stator is coupled to the fuselage.

Example 19: This example includes any or all of the features of example 15, and the heading control device includes a rudder.

Example 20: This example includes any or all of the features of example 16, and the aircraft does not have a tail.

Example 21: This example includes any or all of the features of any one of examples 1 to 20, wherein the rotor shaft is movable between a first position and a second position relative to the fuselage.

Example 22: This example includes any or all of the features of example 21, and further includes defining an angle between the rotor shaft in the first position and the rotor shaft in the second position, the angle being less than or equal to about 90 degrees.

Example 23: This example includes any or all of the features of example 21, and further includes a pivot coupled to the fuselage, the rotor shaft coupled to the pivot, the pivot rotatable about a second axis, and rotation of the pivot about the second axis causes the rotor shaft to move between the first position and the second position.

Example 24: This example includes any or all of the features of example 23, wherein an angle is subtended by the rotor shaft in the first position and the rotor shaft in the second position relative to the pivot, the angle being less than or equal to about 90 degrees.

Example 25: This example includes any or all of the features of example 21, and further includes a frame rail guide coupled to the exterior of the fuselage, and a rotor system cart housing coupled to the frame rail guide, the rotor shaft coupled to the rotor system cart housing, and the rotor system cart housing configured to move along the frame rail guide to cause the rotor shaft to move from the first position to the second position, and vice versa.

Example 26: This example includes any or all of the features of example 2, wherein the fuselage defines an interior volume and includes at least one fuselage structural member, and the aircraft further includes a damping element between the support bearing and the fuselage structural member.

Example 27: This example includes any or all of the features of Example 26, where the damping element is an active damping element or a passive damping element.

Example 28: This example includes any or all of the features of Example 27, and wherein the damping element is a passive damping element, and the passive damping element is a gas shock absorber, a liquid shock absorber, a mechanical shock absorber, or a combination thereof.

Example 29: This example includes any or all of the features of example 28, and the damping element includes a gas strut, a liquid strut, a damping spring, an elastic material, or a combination thereof.

Example 30: This example includes any or all of the features of example 27, and the damping element is an active damping element.

Example 31: This example includes any or all of the features of example 30, and wherein the active damping element is configured to match or cancel harmonic frequencies or vibrations caused by the rotor blades, the thrust support structure, or a combination thereof.

Example 32: This example includes any or all of the features of any one of examples 1 to 31, wherein the rotor blades have a fixed length or a variable length.

Example 33: This example includes any or all of the features of example 32, and the rotor blades have a variable length.

Example 34: This example includes any or all of the features of example 32, and further includes a blade grip including a motor, the blade grip coupled to the rotor blade, and operation of the motor changes the length of the rotor blade.

Example 35: This example includes any or all of the features of any one of examples 1 to 34, wherein the thrust support structure includes a distal end including at least one duct and an opening, the at least one duct adapted to receive a thrust airflow generated by the engine and transmit the airflow to the opening.

Example 36: This example includes any or all of the features of example 35, wherein the thrust airflow has a first temperature T1, the engine is further configured to produce a compressor bleed airflow (compressed bleed airflow) having a second temperature T2, T2 being less than T1, and the at least one duct is further configured to receive the compressor bleed airflow, such that during operation, the compressor bleed airflow is injected into the thrust airflow to produce a mixed airflow having a temperature T3, T2 being less than T1.

Example 37: This example includes any or all of the features of example 35, wherein the thrust airflow has a first temperature T1, the engine is further configured to produce a compressor bleed airflow having a second temperature T2, T2 being less than T1, and the at least one duct is further configured to receive the compressor bleed airflow, such that during operation, the compressor bleed airflow is directed about the thrust airflow.

Example 38: This example includes any or all of the features of any one of examples 1 to 37, and the thrust support structure has an airfoil shape.

Example 39: This example includes any or all of the features of any one of Examples 1 to 38, and further includes an airfoil-shaped fairing disposed about the thrust support structure.

Example 40: This example includes any or all of the features of example 39, and further includes a vane positioning motor coupled to the airfoil-shaped vane, the vane positioning motor operable to change the orientation of the airfoil-shaped vane.

Example 41: This example includes any or all of the features of example 1, wherein the thrust support structure includes a distal end including a first opening and at least one duct, the engine is coupled to or incorporated into the thrust support structure, and during operation, airflow generated by the engine is directed through the at least one duct and through the first opening of the thrust support structure.

Example 42: This example includes any or all of the features of example 41, and further includes a nozzle in fluid communication with the first opening, the nozzle configured to receive and redirect airflow passing through the first opening.

Example 43: This example includes any or all of the features of example 42, wherein the nozzle includes an outlet, air flow passing through the first opening flows through the nozzle and through the outlet, and the nozzle is configured to articulate between at least a first nozzle position and a second nozzle position, and a direction of the outlet at the first nozzle position is different from a direction of the outlet at the second nozzle position.

Example 44: This example includes any or all of the features of example 43, and further includes a drive motor coupled to the nozzle, the drive motor operable to articulate the nozzle between at least the first nozzle position and the second nozzle position.

Example 45: This example includes any or all of the features of any one of examples 1 to 44, wherein the rotor shaft includes a first shaft and a second shaft, the first shaft is disposed within a lumen of the second shaft, the thrust support structure is coupled to the first shaft, and the aircraft further includes a clutch having an engaged state in which the second shaft is coupled to the first shaft and a disengaged state in which the second shaft is disengaged from the first shaft, and during operation, rotation of the thrust support structure about the first axis causes the clutch to enter the engaged state, which in turn rotates the second shaft and the rotor blades about the first axis.

Example 46: This example includes any or all of the features of example 45, wherein the aircraft includes a first rotor blade coupled to the second shaft and a second rotor blade coupled to the second shaft above the first rotor blade, each of the first and second rotor blades configured to rotate about the first axis, and during operation, rotation of the thrust support structure about the first axis causes the clutch to enter the engaged state, which in turn rotates the second shaft and the first and second rotor blades about the first axis.

Example 47: This example includes any or all of the features of example 45, and further includes a third shaft, the second shaft is disposed within a lumen of the third shaft, the third shaft is coupled to the body and does not rotate, and at least one concentric bearing is disposed between the second shaft and the third shaft to separate the second shaft from the third shaft.

Example 48: This example includes any or all of the features of any one of examples 1 to 47, and further includes a thruster for providing horizontal thrust during flight of the aircraft.

Example 49: This example includes any or all of the features of example 48, and further includes a second rotor blade configured to rotate about a second axis, the second axis being perpendicular to the first axis, the thruster.

Example 50: This example includes any or all of the features of example 45, and further includes a thruster for providing horizontal thrust during flight of the aircraft.

Example 51: This example includes any or all of the features of example 50, and further includes a third rotor blade, the thruster configured to rotate about a second axis that is perpendicular to the first axis.

Example 52: This example includes any or all of the features of any one of examples 1 to 51, and the thrust support structure is coupled to the rotor shaft above or below the rotor blades.

Example 53: This example includes any or all of the features of any one of examples 1 to 52, wherein the thrust support structure and rotor blades are coupled to the rotor shaft and are disposed in substantially the same plane.

Example 54: This example includes any or all of the features of example 53, and further includes a common structural hub coupled to the rotor shaft, and the thrust support structure and the rotor blades are coupled to the rotor shaft and are disposed in substantially the same plane.

Example 55: This example includes any or all of the features of example 54, wherein the common structural hub is coupled to a distal end of the rotor shaft, and the proximal end of the rotor shaft is disposed within the fuselage.

Example 56: This example includes any or all of the features of any one of examples 1 to 55, and the engine includes a fan coupled to a distal end of the thrust support structure.

Example 57: According to this example, an aircraft is provided that includes a fuselage, a rotor shaft, a rotor hub coupled to the rotor shaft, an engine coupled to the rotor hub, and a thrust support structure coupled to a distal end of the engine, the thrust support structure configured to function as a rotor blade for the aircraft, the thrust support structure including piping having a proximal end, a distal end, and at least one channel, the at least one channel fluidly coupled to the engine such that, during operation, thrust generated by the engine is received and redirected by the channel to rotate the thrust support structure.

Example 58: This example includes any or all of the features of example 57, and the aircraft has no rotor blades other than the thrust support structure.

Example 59: This example includes any or all of the features of example 57, and further includes a support bearing coupled to the fuselage, the support bearing isolating the fuselage from torque generated by at least rotation of the thrust support structure.

Example 60: According to this example, an aircraft is provided that includes a fuselage, an engine mounted on the fuselage, a diverter manifold coupled to the engine, a thrust diverter disposed in the diverter manifold, and a rotor shaft coupled to the engine, the rotor shaft including a shaft duct in fluid communication with the diverter manifold, and in operation, the engine generates an airflow directed to the diverter manifold, and the thrust diverter is configured to control a relative amount of the airflow within the diverter manifold that is directed to the shaft duct.

Example 61: This example includes any or all of the features of example 60, and further includes a plenum coupled to the rotor shaft and in fluid communication with the shaft duct, and a rotor blade coupled to the plenum, the rotor blade including a blade duct in fluid communication with the plenum and further including an opening, the plenum configured to direct at least a portion of the airflow directed toward the rotor duct to the blade duct such that all or a portion of the airflow exits the opening.

Example 62: This example includes any or all of the features of example 60, and the engine is located on or within the fuselage.

Example 63: This example includes any or all of the features of example 61, and further includes a support bearing coupled to the fuselage, the support bearing isolating the fuselage from at least torque generated by rotation of the rotor blades.

Example 64: According to this example, a stationary rotor shaft is provided with a fuselage, an engine, a stationary rotor shaft coupled to the fuselage, the stationary rotor shaft including a proximal end, a distal end, and a shaft duct extending from the proximal end to the distal end, the shaft duct being in fluid communication with the engine, a thrust support structure coupled to the stationary rotor shaft, the thrust support structure including an inlet opening, an outlet opening, and a thrust duct in fluid communication with the rotor duct, the thrust support structure including an inlet opening, an outlet opening, and a thrust duct in fluid communication with the inlet opening and the outlet opening, a rotatable outer shaft disposed around at least a portion of the distal end of the rotor shaft, and a rotatable outer shaft disposed around at least a portion of the distal end of the stationary rotor shaft. An aircraft is provided that includes a clutch connected to a rotor shaft, the clutch having an engaged state and a disengaged state, and in operation, airflow from an engine 5 is directed through the rotor duct, through the inlet opening, through the thrust duct, and through the outlet opening, thereby rotating the thrust support structure, and the rotation of the thrust support structure causes the clutch to enter the engaged state in which it is connected to the outer shaft, so that the torque generated by the rotation of the thrust support structure is transmitted to the outer shaft, causing the outer shaft and the rotor blades to rotate.

Example 65: This example includes any or all of the features of example 64, and further includes a support bearing coupled to the stationary rotor shaft and the outer shaft, the support bearing isolating at least the fuselage from torque generated by rotation of the outer shaft and the rotor blades. .

Example 66: This example includes any or all of the features of example 64, and in operation, the engine causes the thrust support structure to rotate about a first axis at a first rotational speed that is greater than or equal to the second rotational speed of the outer shaft, causing the clutch to enter the engaged state and grip the outer shaft.

Example 67: This example includes any or all of the features of example 66, and the clutch enters or is in the disengaged state when the first rotational speed is less than the second rotational speed.

Example 68: This example includes any or all of the features of example 64, wherein the clutch includes a clutch outer case, a clutch inner race connected to the outer shaft, and a clutch movable bearing, the clutch outer case includes a ramp and a cavity, and in the engaged state, the clutch movable bearing is disposed between and in contact with both the ramp and the clutch inner race, thereby engaging the clutch outer case and the clutch inner race, and in the disengaged state, the clutch movable bearing is disposed within the cavity and only in contact with the clutch outer case.

Example 69: This example includes any or all of the features of any one of examples 1 to 67, and the aircraft does not include a counter torque mechanism.

While the principles of the present disclosure have been described herein, it should be appreciated by those skilled in the art that this description is made by way of example only and is not intended to be limiting with respect to the scope of the invention as claimed. Features and aspects described with reference to specific embodiments disclosed herein are susceptible to combination and/or application with various other embodiments described herein. Such combinations and/or applications of such described features and aspects with such other embodiments are contemplated herein. Modifications and other embodiments are contemplated herein and are within the scope of the present disclosure.

Claims (23)

1. An aircraft,
The torso and
a support bearing coupled to the fuselage;
a rotor shaft mounted to a fuselage by support bearings, the rotor shaft rotatable about a first axis extending through the rotor shaft, the support bearings allowing the rotor shaft to rotate about the first axis relative to the fuselage at all times;
rotor blades coupled to the rotor shaft, such that rotation of the rotor shaft causes the rotor blades to rotate about a first axis;
a thrust support structure connected to the rotor shaft;
an engine coupled to the thrust support structure and configured to rotate at a radially offset distance from the first axis, the engine configured to receive fuel and combust the fuel to generate a thrust airflow;
a clutch disposed between the rotor blades and the rotor shaft;
the clutch is configured to selectively transfer torque to the rotor blades;
the thrust support structure further comprises a distal end including at least one pipe and a nozzle, the at least one pipe configured to receive the thrust airflow generated by the engine and transfer the thrust airflow to the nozzle, the nozzle configured to receive the thrust airflow through an outlet and redirect it against the atmosphere to cause the thrust support structure to rotate about the first axis;
During operation,
rotation of the thrust support structure causes the rotor shaft to rotate about the first axis, which in turn causes the rotor blades to rotate about the first axis;
the support bearings isolate the fuselage from at least the torque generated by the rotation of the rotor blades;
aircraft. 1. An aircraft,
The torso and
a support bearing coupled to the fuselage;
a rotor shaft mounted to a fuselage by support bearings, the rotor shaft rotatable about a first axis extending through the rotor shaft, the support bearings allowing the rotor shaft to rotate about the first axis relative to the fuselage at all times;
rotor blades coupled to the rotor shaft, such that rotation of the rotor shaft causes the rotor blades to rotate about a first axis;
a thrust support structure connected to the rotor shaft;
an engine coupled to the fuselage or the thrust supporting structure;
a clutch disposed between the thrust support structure and the rotor shaft and configured to selectively transfer torque to the rotor blades;
During operation, the engine causes the thrust support structure to rotate about the first axis at a first rotational speed that is greater than or equal to a second rotational speed of the rotor shaft, causing the clutch to enter an engaged state and grip the rotor shaft, and the clutch enters or is in a disengaged state when the first rotational speed is less than the second rotational speed;
rotation of the thrust support structure causes the rotor shaft to rotate about the first axis, which in turn causes the rotor blades to rotate about the first axis;
The support bearing isolates the fuselage from torque generated by rotation of at least one of the rotor blades. 1. An aircraft,
The torso and
a support bearing coupled to the fuselage;
a rotor shaft mounted to a fuselage by support bearings, the rotor shaft rotatable about a first axis extending through the rotor shaft, the support bearings allowing the rotor shaft to rotate about the first axis relative to the fuselage at all times;
rotor blades coupled to the rotor shaft, such that rotation of the rotor shaft causes the rotor blades to rotate about a first axis;
a thrust support structure connected to the rotor shaft;
an engine coupled to the fuselage or the thrust supporting structure;
a clutch disposed between the thrust support structure and the rotor shaft and configured to selectively transfer torque to the rotor blades;
the clutch includes a clutch outer case, a clutch inner race connected to the rotor shaft, and a clutch movable bearing;
the clutch outer case includes a ramp and a cavity;
In an engaged state , the clutch movable bearing is disposed between and in contact with both the ramp and the clutch inner race, thereby engaging the clutch outer case and the clutch inner race;
In a disengaged state , the clutch movable bearing is disposed within the cavity and only contacts the clutch outer case. the support bearing includes a bearing housing having inner and outer race cages and at least one bearing between the inner and outer race cages;
the rotor shaft extends into the fuselage and into the support bearings;
2. The aircraft of claim 1. further comprising a bearing support structure;
the support bearing is connected to the bearing support structure, and the bearing support structure is connected to the fuselage;
the rotor shaft is connected to the fuselage only by the support bearings and the bearing support structure;
5. An aircraft as claimed in claim 4 . The aircraft of claim 1, wherein the engine is coupled to the thrust support structure, the thrust support structure including a proximal portion coupled to the rotor shaft and a distal portion coupled to the engine. The aircraft of claim 1, wherein thrust in the form of airflow generated by the engine is directed generally along thrust lines extending horizontally above, below, or above the rotor blades. The aircraft of claim 1, further comprising a heading control device. the orientation control device includes an electric reversible motor including a motor shaft;
The motor shaft is the rotor shaft.
9. An aircraft as claimed in claim 8 . 10. The aircraft of claim 9, wherein the azimuth control device includes a rotating magnet and a stator, and when the stator is energized, a force is applied to the rotating magnet which in turn provides a force to the fuselage to rotate the fuselage about the first axis. the orientation control device includes a spacer coupled to the rotor shaft;
the rotating magnet is coupled to the spacer;
The stator is connected to the fuselage.
11. The aircraft of claim 10 . 10. The aircraft of claim 9 , wherein the aircraft is tailless. the rotor shaft is movable between a first position and a second position relative to the fuselage;
an angle is defined between the rotor shaft at the first position and the rotor shaft at the second position, the angle being less than or equal to about 90 degrees;
2. The aircraft of claim 1. a pivot coupled to the fuselage, the rotor shaft coupled to the pivot;
the pivot is rotatable about a second axis;
Rotation of the pivot about the second axis causes the rotor shaft to move between the first position and the second position.
14. An aircraft as claimed in claim 13 . A frame rail guide connected to the outer side of the fuselage;
a rotor system cart housing coupled to the frame rail guide;
the rotor shaft is coupled to the rotor system cart housing;
the rotor system cart housing is configured to move along the frame rail guides to cause the rotor shaft to move from the first position to the second position, and vice versa.
14. An aircraft as claimed in claim 13 . the fuselage defines an interior volume and includes at least one fuselage structural member;
the aircraft further includes a damping element between the support bearing and the fuselage structural member, the damping element being an active damping element or a passive damping element;
2. The aircraft of claim 1 . 17. The aircraft of claim 16 , wherein the damping element is a passive damping element, the passive damping element being a gas shock absorber, a liquid shock absorber, a mechanical shock absorber, or a combination thereof. The aircraft of claim 1, wherein the rotor blades have a variable length. the rotor blades having a variable length, the aircraft further comprising a blade grip including a motor;
the blade grip is coupled to the rotor blade;
Operation of the motor varies the length of the rotor blades.
20. The aircraft of claim 18 . the thrust airflow has a first temperature T1;
The engine is further configured to produce a compressor bleed airflow having a second temperature T2, T2 being less than T1.
2. The aircraft of claim 1. During operation,
the at least one conduit is further configured to receive the compressor bleed airflow, such that during operation, the compressor bleed airflow is injected into the thrust airflow to create a mixed airflow having a temperature T3; or
the at least one duct is further configured to receive the compressor bleed airflow, such that during operation, the compressor bleed airflow is directed about the thrust airflow.
21. The aircraft of claim 20 . 1. An aircraft,
The torso and
a support bearing coupled to the fuselage;
a rotor shaft mounted to a fuselage by support bearings, the rotor shaft rotatable about a first axis extending through the rotor shaft, the support bearings allowing the rotor shaft to rotate about the first axis relative to the fuselage at all times;
rotor blades coupled to the rotor shaft, such that rotation of the rotor shaft causes the rotor blades to rotate about a first axis;
a thrust support structure connected to the rotor shaft;
an engine coupled to the fuselage or the thrust supporting structure;
During operation, the engine causes the thrust support structure to rotate about the first axis;
rotation of the thrust support structure causes the rotor shaft to rotate about the first axis, which in turn causes the rotor blades to rotate about the first axis;
the support bearing isolates the fuselage from torque generated by rotation of at least one of the rotor blades;
the rotor shaft includes a first shaft and a second shaft, the first shaft being disposed within a lumen of the second shaft;
the thrust support structure is connected to the first shaft;
the aircraft further includes a clutch having an engaged state in which the second shaft is coupled to the first shaft and a disengaged state in which the second shaft is disengaged from the first shaft;
In operation, rotation of the thrust support structure about the first axis causes the clutch to enter the engaged state, which in turn causes the second shaft and the rotor blades to rotate about the first axis. Further comprising a third shaft;
the second shaft is disposed within the lumen of the third shaft;
the third shaft is connected to the body and does not rotate;
At least one concentric bearing is disposed between the second shaft and the third shaft and separates the second shaft from the third shaft.
23. The aircraft of claim 22 .

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