Moveable/adjustable center of gravity helicopter technology

Christophe Pochari, Pochari Technologies, Bodega Bay, CA

Conventional rotorcraft with enhanced center of gravity envelope technology: Current status of development and future potential for implementation on production aircraft:

The rotorcraft industry is highly consolidated, mature and competitive. The industry is characterized by a very conservative attitude towards design changes that deviate considerably from the established criteria. One may sympathize with this viewpoint, considering the high degree of regulatory control, high cost of certification, and risk of unanticipated design flaws leading to periodic mechanical failure on production aircraft. This can in turn permanently damage the reputation of a particular model or brand name and result in high costs related to insurance and liability.
Predicting and anticipating design flaws that may go unnoticed during the design phase remains challenging even in today’s design environment where simulation and computationally assisted design are extensively relied upon. With this in mind, it can seem to be an excessively high-risk endeavor to implement major design changes on mature and proven architectures regardless of their innovative potential. In order to justify the implementation of these design improvements, it’s critical that these design innovations enhance performance, increase versatility, improve safety, or most importantly, enhance the operating scope, enabling the aircraft to perform a wider range of missions profiles, or to fulfill the requirements of more demanding missions that would increase the value proposition of the aircraft.

With this in mind, this article aims to provide an overview of a highly novel and unconventional rotorcraft drive train configuration. This configuration is called adjustable center of gravity, abbreviated ACG.

Exceeding the allowable CG enveloped imposes serious flight control limitations referred to as cyclic over saturation. As the disc is tilted in a certain direction as a result of the mass exceeding the lift distribution equilibrium around the pivot point (main rotor shaft), a cyclic response is applied to rectify the imbalance, this reduces the amount of cyclic pitch remaining and available to perform unrelated flight controls. Eventually, all the pitch movement of the blade is exhausted, rendering the aircraft uncontrollable. This potentially dangerous scenario is why precise load placement is so critical, and the reason conventional rotorcraft are designed with strict CG envelopes established which can not be exceeded without severely compromising the aircraft’s maneuverability and safety.
Shifts in center of gravity are inherently very sensitive in a conventional single main rotorcraft helicopter, for the simple reason that a helicopter is analogous to a load suspending from a crane, which acts like a pendulum.
If a load was suspending via a single cable attached upon the exact center of a rectangle load, a slight imbalance and the load would permanently droop to one side. A crane corrects this by using a triangular cable configuration, where the load concentration is distributed over 4 cables attached to each corner of the load, which then converge into a single cable at a higher point along the vertical cable.

For this reason in order to fully optimize the conventional rotorcraft, it is necessary to expand the allowable CG envelope.

A multitude of reasons exist to provide the impetus to implement the expanded CG envelope system.

One of the many incentives may be provided by the requirement to rapidly unload various kinds of munitions without requiring the munitions to be stored directly in the CG envelope. If munitions are stored on the aircraft, such as small arms fire or larger air-ground missiles, when these munitions are fired, the aircraft’s is quickly reduced, this requires the munitions to be stored in a narrow CG envelope. With ACG technology, the munitions can be stored outside the CG envelope without impairing stability.
With the ACG system there exists the possibility of placing the fuel tanks outside of the CG envelope where fuel is traditionally stored. This unorthodox fuel placement would permit a more flexible design by freeing up the designer to utilize more of the valuable space in the CG envelope, increasing the allowable volume by eliminating the traditional fuel storage by shifting the fuel mass to a region of the aircraft which has less storage value.
Fuel mass is considerable, one of the single largest loads carried by the aircraft. This necessitates placement within the CG envelope. fuel tanks could be placed in the rear of the aircraft rather than underneath the bottom of the fuselage as centering the fuel mass is no longer necessary.
Another benefit and arguable on of the most important is requirement to provide rapid and unforeseen unloading of personnel while simultaneously carrying other loads such as cargo or equipment.
A helicopter may be loaded with troops in the rear portion of the helicopter, with equipment placed in the front portion. This configuration may be calibrated to insure a even load distribution despite the greatly differing load densities and profiles. Since the load in the front portion is fixed and cannot be distributed and scattered along the fuselage, if suddenly the personnel were to exit the aircraft, the load becomes highly concentrated towards the front, requiring a sudden rearrangement of the equipment towards the rear portion to correct the sudden imbalance. The equipment could be moved, but this entails a considerable hassle and consumes time, loads may be strapped down onto the floor, and require special handling equipment that may not be available on the aircraft. Even if the load were moved, the personnel has to eventually reembark into the aircraft, requiring the load to be rearranged yet again.
This hypothetical scenario illustration makes vivid the need for CG flexibility. One may argue the above mission profile is not frequently performed, this may be the case, but only because current aircraft simply cannot perform it without considerable difficulty and inconvenciance.
To avoid this problem highlighted above, separate aircraft have to be provided to perform their respective missions and cannot readily deviated from the missions which they are tailored to perform. This reduces the flexibility and versatility of the aircraft.
Even if the imbalances are small, flight control response is still required.
Rather than requiring the pilot to constantly compensate by applying cyclic to correct the tendency of the aircraft to pitch forward or backward, the operator can carefully calibrate the rotor mast’s position relative to the fuselage, until the aircraft naturally finds it’s point of maximum stability during all flight regimes.
If a rotorcraft is traveling at higher speed, the rotor mast is tilted forward, along with the entire fuselage, leading to the center of gravity to move further forward, once the aircraft slows down, the rotor mast moves back to a level position, causing the load center to move further back. These load shifts and imbalances occur frequently during operation in different flight conditions. The horizontal stabilizer serves the function of leveling the aircraft during forward flight, the horizontal stabilizer thus also serves as a way to enhance the CG range, but only during forward flight. The limitation of the horizontal stabilizer is that it requires airspeed to provide lift, thus serves no function at zero airspeed during hovering. As a result, the horizontal stabilizer cannot provide any increase in CG envelope.

In order the solve this fundamental limitation faced by the single main rotor helicopter, three two options are allotted to the designer.

The most obvious option is to simply add an additional rotor disc. The development of the tandem rotor helicopter was partially developed as a need to enable more flexible center of gravity. A tandem rotor helicopter in spite of solving the GC problem, imposes other challenges.
A tandem rotor helicopter will have a significantly larger footprint, impairing confined landing capability.
A tandem rotor design will also increase the empty weight and add significant additional complexity by doubling the number of dynamic components, increasing the probability of mechanical failure. Although stability is no longer an issue with modern flight controls, a tandem rotor requires an interconnecting driveshaft, gearbox and redundant flight controls in the event of an engine failure as stability is completely lost in the event of a partial or complete lift in one set of rotor discs.

The ehanced CG evenlope capability of the tandem rotorcraft is highlighted below

“The ability to adjust lift in either rotor makes it less sensitive to changes in the center of gravity, important for the cargo lifting and dropping. While hovering over a specific location, a twin-rotor helicopter has increased stability over a single rotor when weight is added or removed, for example, when troops drop from or begin climbing up ropes to the aircraft, or when other cargo is dropped”

Nick Van Valkenburgh describes the need for enhancing the CG range of helicopters.

“single-rotor helicopters were successful in their limited military service in WWII, they were restricted in payload and had serious center-of-gravity limits.”

Valkenburgh describes one of the main advantages of the tandem rotor is the “ The ability to almost indiscriminately load personnel and cargo (extraordinary center of gravity range)”

Clearly this illustrates that center of gravity and load placement are on of the most salient issues facing helicopter operation and design. Thus there is great impetus for developing alternative designs which alleviate this limitation.
Can we combine the simplicity and legacy of the single main rotor helicopter with enhanced CG technology? This is the crux of this paper, as we believe the single main rotor helicopter is a candidate for this technology.

For single main rotor configurations, there exist two realistic options
The second option is to tilt the entire rotor disc, not through quasi tilting as provided by cyclic action, but rather the complete set of blades pivots in a back and forth analogous to a tiltrotor, albiet to a much more limit extent, as the fuselage and tailboom imposes limitation on the degree of tilting allowable.
This method was developed by Floyd Carlson at Bell Helicopter in 1975 and described in a patent.

Franz Weinhart in Germany conceived of an alternative design that appears to enable the helicopter to perform the basic flight controls by constantly adjusting its own center of gravity, thus it can be said that Franz Weinhart’s design seems to have been born out of a desire to develop an entirely new flight control method that eliminates the need for cyclic control rather than for expanding the center of gravity envelope, despite this, the design achieves both objectives. Franz Weinhart’s patent description is provided below.

“In the case of the helicopter of the invention the rotor system is able to be moved in translation in the longitudinal direction of the helicopter fuselage together with the drive system and is able to be pivoted about a pivot axis running along the fuselage. Owing to the longitudinal displacement of the rotor and drive system in relation to the fuselage the center of gravity of the helicopter is so changed that the helicopter is inclined about its transverse axis forward and, respectively, backward so that forward flight may be accelerated and, respectively, retarded. The pivoting about the pivot axis running along the fuselage on the other hand causes an inclination of the helicopter to the left or to the right so that it is possible for corresponding curves to be flown.
The helicopter in accordance with the invention offers the advantage that cyclical blade control, that is to say the swash plate and its control elements, may be completely dispensed with so that the overall structure is substantially simplified. This means that there are lower costs of production, less wear, substantially longer intervals between servicing and therefore lower serving costs. Furthermore mechanical effort for cyclical blade angle addition is no longer necessary and accordingly the efficiency of the overall system is improved. A further advantage is that owing to the possibility of inclining the rotor system about the longitudinal axis of helicopter improved take-off and landings on hills become possible.
The control of the center of gravity of the helicopter of the invention further leads to a significant saving in weight and to operation of the rotor with less vibrations. Moreover following engine failure gliding with the rotor freewheeling (autorotation) is substantially simpler, the use of a suitable inclination of the rotor blades shortly before landing meaning that a relatively gentle touch-down is possible even without engine drive. The helicopter furthermore responds extremely rapidly to corresponding changes in the center of gravity so that extremely precise and simple control or steering of the helicopter is possible. Unlike known helicopters response to control commands is improved with an increase in load, that is to say a higher mass of the fuselage instead of being reduced.
In accordance with an advantageous embodiment of the invention a semi-cardanic suspension is provided for longitudinal transverse and pivoting of the rotor system, such suspension having at least one central support axle held on the fuselage. The central support axle in this case simultaneously defines both the longitudinal axis, along which the rotor system and possibly the drive system is able to be longitudinally moved, and also the pivot axis, about which the rotor and possibly the drive as system well may be laterally tipped.
In accordance with an advantageous embodiment of the invention the adjustment of the blade angle is performed by means of longitudinal displacement of a sliding sleeve, which is held on the rotor shaft in a manner allowing sliding but not rotary movement and is functionally connected with the rotor blades.
Such a sliding sleeve is preferably arranged to be longitudinally slid by means of a linkage, which is fixed on the sliding sleeve in the direction of sliding while being held in relation to same while allowing relative rotary movement and is able to be slid using a lever mechanism in the longitudinal direction of the rotor shaft”

Noboru Okada at Mishubishi Heavy Industries developed a smiliar concept involving a longitudinal sliding concept. The patent was unfortunately retracted, thus no images are available, the patent description despite being somewhat difficult to comprehend due to the translation, appears to describe a linear adjustable rotor mast relative to fuselage configuration similar to Franz Weinhart’s invention.

“The above-described conventional helicopter, there is a problem of the next to be solved. 
For the center of gravity movement allowable range is narrow, it can not operate as greatly barycentric position changes. Not always able to keep the aircraft horizontally during flight During forward flight, for center of gravity against the head down moment due to the fact that deviates from the main rotor shaft axis, it must cause fog up moment (downward lift) by the horizontal stabilizer.
The present invention aims to provide a center of gravity mobile helicopter capable of maintaining aircraft horizontally even weight shift easy and in flight in which the above-described problems.
The present invention is a solution to the above problem, each other and become more aircraft fuselage portion and a movable portion which is relatively movable split in the longitudinal direction, and a body portion and a movable portion relative a movable means for moving the, is intended to provide a center of gravity mobile helicopter, characterized by comprising; and a main rotor and Till rotor provided maintaining a predetermined distance to the movable portion.
Since the present invention is constructed as described above has the following effects.
Namely, the longitudinal direction with more becomes airframe and relatively movable divided body portion and a movable portion, and for a movable means for relatively moving the body portion and a movable portion, the body portion and a movable portion the relative movement by the movable unit, it is possible to move the center of gravity of the helicopter in an optimum position to operate.
This also barycentric position tolerance is greatly expanded.
Since the movement of the center of gravity is there is no need to perform by the pitch operation or the like of the main rotor as in the prior art, there is no need to tilt the aircraft. In addition, it is also possible to change the aircraft attitude by changing the position of the center of gravity in reverse.
According to the present invention has an effect such as the following because it is constructed as described above. It is possible to freely change the center of gravity position,
Conventionally been impossible can be also safe flying against substantial center-of-gravity position changes. Attitude of the aircraft in flight can be kept horizontally. By changing the position of the center of gravity, it is possible to change the aircraft attitude. It is possible to miniaturize the horizontal stabilizer.

Floyd Carlson’s patent describes the impetus for developing the system.

“Helicopters are loaded, unloaded and reloaded with different cargoes. The center of gravity of a loaded fuselage changes in location from load to load depending upon the load position in the cabin. This often requires readjustment of load position or careful initial distribution of load components in order to end up with the position of the center of gravity within a limited field. When the center of gravity is thus positioned within the limited field, the aircraft may then be maneuvered within the limits dictated by its design with out difficulty. When a given ship is certified by the governmental authorities, the CG field, size and location are certified and specified. The ship, once certified, cannot legally be operated with loads such that the CG is outside the certified field”
“The present invention permits in an aircraft equipped with the controls and linkages such as embodied in the present invention to accommodate load CG’s to be positioned anywhere within a wide range. Utilizing the present invention, there is a greater range over which the center of gravity may be positioned without exceeding the limits of control stick movement for the prescribed maneuvers of the aircraft. This means that aircraft loading may not require the same precision or discipline with the present invention as in ships that do not embody the present invention. This permits basic design of the aircraft to be altered to take advantage of the increased range of CG location. Stated otherwise, less control is required for the same CG range in a ship embodying the present invention compared with one that does not embody the present invention”

Push-pull rods are currently installed on the vast majority of turbine helicopters, these flight control transmission mechanisms are not conducive to allowing movement that alters their alignment, thus, it may be said, that fly by wire flight controls may be more attractive to enable a dynamic rotor drive system.
In addition to issues with the flight control system, are potential issues related to the hydraulic system. On a turbine helicopter, a hydraulic pump provides pressurized hydraulic fluid to servos (linear actuators) installed right beneath the stationary swashplate, these servos are typically mounted on the main rotor gearbox. Flexible hydraulic lines can be used to accommodate the movement, flexible hydraulic lines have been successfully used on tilting nacelle tiltrotors and proven to be perfectly reliable. Despite this, it’s likely that requirement of having to provide for the movement of flight controls and hydraulic lines imposes a certain reliability penalty, which could potentially compromise redundancy and subsequent implications for safety, although
An additional downside of this method is that it requires a series of clutches and freewheel units. In order to permit a shaft to rotate along another rotating shaft beneath it, a device called a freewheel unit is used. The input driveshaft delivering power from the engine to the gearbox is rotating at a fixed speed, if the mast is required to pivot, an elaborate configuration consisting of a total
Figure 2 illustrates a pivoting gearbox mechanism

of four dynamic units enables the shaft to pivot along the axis of rotating of the driveshaft. A single freewheel unit enables the outer shaft to rotating past the underlying shaft in the same direction, this requires two modules to provide forward and rear movement. The mechanism would be similar to a proprotor gearbox tilting mechanism found on tiltrotors such as the V-280 Valor.

Of the two options available, the first method suffers a slightly higher weight penalty than the latter method, although both methods add a perfectly acceptable weight penalty considering the performance benefits derived.
For the multitude of reasons discussed, the Carlson adjustable center of gravity system, despite having had the potential revolutionize helicopter technology, was never successfully implemented or further researched.

The second option developed was first invented by Franz Weinhart in 1998 which involves sliding the rotor shaft along the longitudinal direction of the fuselage. This is the option found to be most attractive but could be integrated with a pivoting system if one desired even additionally expanded CG range. The distance is determined by the desired increase in CG enveloped which may vary depending on the nature of the operation. The exact working mechanism is surprisingly simple.
A set of weight-bearing roller mechanisms slide on the bottom of the main structural element, comprised of dual frame/rail, with a hollow section in between. A set of rollers provide the load-bearing capacity and also serves to enable the linear forward and back movement. Lightweight composite hydraulic cylinders provide the actuation force to slide the gearbox mechanism along the driveshaft and rail, the driveshaft remains fixed along with the engines and main reduction gearbox, only the 90 gearbox and mast assembly are dynamic. The hydraulic cylinders also serve to provide additional longitudinal stability of the gearbox and mast. The hydraulic cylinders also lock and remain rigid to maintain the gearbox in a fixed position. Along the 40 inches of movement, the cylinder can stop extending at any time and hold the gearbox at a given point along the displacement line.
Lateral stability of the gearbox and mast is provided by the non-load bearing struts which also slide along the rail. The sliding driveshaft mechanism enables the 90-degree gearbox to slide as the shaft is spinning. The drive shaft contains a pattern of grooves, which permits the rolling of the ball bearings, which provide the torque transmission as well as permitting linear sliding. The driven unit, which turns the 90 degree gear, houses the bearing balls.

Two methods are considered to configure the gearbox sliding system.
In a conventional helicopter gearbox, a large portion of the reduction takes place in the gear assembly located beneath the rotor mast in the vertical direction. In light helicopters, the turboshaft typically has an output speed of around 6000 rpm, the power turbine is rotating at 30,000-50,000 rpm depending on the number of stages. Light rotorcraft turboshafts are often equipped with an integrated gearbox. A large portion, around 50% of the total speed reduction takes place in the engine integrated gearbox. On larger rotorcraft, the turboshafts output speed is equal to the power turbine speed, no reduction takes place at the engine, for example, the GE T700, T64 and T408 provide no speed reduction. On these large rotorcraft,
The main rotor gearbox provides 100% of the reduction, this requires a large amount of gearing placed in the vertical direction beneath the rotor mast, this means the main rotor gearbox assembly is large and elaborate, requiring more housing and occupying more space. Due to the large amount of space occupied by a reduction gearbox configuration, more room must be provided along the course of movement, in this case, 40 inches of displacement is provided. This enables the bevel gearbox and mast assembly to remain compact, occupying less volume during its course of movement. In order to design the bevel gearbox assembly to be as compact as possible, the gearbox consists only of a 90-degree bevel gear connecting the drive shaft to the rotor mast, all speed reduction takes place in an engine integrated gearbox located in the rear of the aircraft outside of the designated displacement boundary.
Directly beneath the 90-degree gearbox assembly is a grip, which connects to a composite strap, which then connects to the roller assembly. This unit bares the weight of the entire rotorcraft, as a result, the 90-degree bevel gearbox is a major structural element.
A large structural member shaped as a beam integral with the fuselage structure spans the distance of the displacement boundary directly in the center of the fuselage, right beneath this member a metallic liner isolated from the composite beam by a compressible rubber layer is placed, which forms the rolling surface. The compressible layer protects the carbon fiber beam from sudden compressive force during violent vertical acceleration. If rotor trust is quickly increased, the main gearbox moves upward, pulling with it the roller assembly, the compressible layer attenuates this tendency reducing load on the fuselage structure.
Stabilization of the gearbox is a critical design requirement to insure the phenomenon of “mast rocking” does not occur. Mast rocking is a serious phenomenon which can cause major damage to the airframe and gearbox.
It can be argued that the adjustable CG system elevates the risk of mast rocking, for this reason, several precautionary design features are implemented.
The main rotor gearbox (MRG) can be thought of as the “heart” of the helicopter. It bares the entire weight of the loaded aircraft in flight, it serves to connect the fuselage to the rotor assembly, arguable the most critical function of the helicopter.
In most rotorcraft main gearbox systems, the load is transferred to the fuselage directly beneath the gearbox, multiple struts comprised of a metal or composite tube are angled at a 45-50 degree angle connect from the fuselage deck to the top of the gearbox, just beneath where the mast protrudes.
This configuration forms a triangular highly rigid frame, forming a truss-like shape.
In some designs, rather than a direct load path, a “pylon” system is used that extends the load path a greater distance away from the center of the gearbox.
In these designs, the pylon also serves the function of the stabilization strut. Since this design is inherently less rigid, the extending pylon is usually placed at a higher point along the gearbox closer to the mast protrusion line. The mast protrusion line is the point where the mast extends out beyond the gearbox casing. Placing the pylon further up reduces the distance between the connection point to the fuselage and the rotor head, this distance determines the level of force imposed on the connection during strong banks. During sudden lateral movements, a tremendous amount of stress is placed on the lateral gearbox connection. For this reason, it’s important to minimize the distance between the lateral stabilization connection point and the rotor head. This obviously means a coaxial helicopter will require a stronger gearbox-fuselage connection.

In order for the ACG system to provide the required rigidity and stability it is necessary to minimize mast rocking. To achieve this, the rails that permit the struts from sliding along with the gearbox must be sufficiently rigid to prevent a small amount of displacement.
It’s also critical to minimize even a slight unintended creepage along the guide rail. The hydraulic cylinder that extends and retracts can be locked into position, but there remains a slight amount play simply due to the comprehensibility of hydraulic fluid. No fixed locking mechanism is in place in the longitudinal direction, this means if strong forward cyclic is placed, the disc will tilt forward, along with it the mast and gearbox, this force will cause a tendency for the entire rotor drive assembly to slide forward a small amount. In order to prevent this, when the operator decides on a CG position, both sides of the hydraulic cylinders are pressurized, forming a barrier of high pressure fluid on both sides of the piston, preventing the extending arm from being push or pulled in the event of strong forward motion.

Illustration of the CG dynamic linear movement

Sliding driveshaft mechanism, with spherical roller bearings permitting sliding along torque path.

CAD model depicting the 90-degree bevel gearbox along the sliding driveshaft.

CAD models depicting the load bearing sliding mechanism

Krishnamurthi and Gandhi 2015 investigated a swashplateless rotorcraft using CG adjustment for cyclic control.

“For the swashplateless configuration, a total forward cg travel of 2.48 ft was required to trim the aircraft with increasing speed up to 120 knots. The lateral cg travel required was only 0.32 ft. By changing the horizontal tail slew schedule so it provided larger nose-down moments on the aircraft at moderate- to high-speed, the longitudinal CG travel requirements could be reduced to 0.77 ft”
An altogether different approach to swashplateless primary control, eschewing the use of on-blade TEFs, was presented by Gandhi, Yoshizaki, and Sekula. In this study, the authors proposed using rotor RPM variation in lieu of collective pitch control and moving the aircraft center-of-gravity (CG) in lieu of cyclic pitch control.
The CG could be moved, for example, by placing a fuel tank, batteries or payload on tracks and using actuators to move them in the fore-aft and lateral directions. Results, based on a swashplateless variant of a Robinson R22 type aircraft, showed that trim could be achieved at high speeds, and forward CG movement requirements
could be reduced by introducing a forward tilt of the rotor shaft or setting the horizontal tail at a nose-up angle of attack relative to the aircraft waterline”

Although it was not the originally intention for developing ACG technology, Weinhart, Krishnamurthi and Gandhi realized the potentials of either fully swashplateless control architecutre or to simply provided enhancement maneuverability offered by adjustable center of gravity technology.
Pochari Technologies believes the main benefit derived will be enhanced load placement flexibility.

The image above illustrates a conventional gearbox-fuselage mounting system.

An alternative but heavier option is forgoing the sliding driveshaft in favor of a hybrid drivetrain. With advancements in high power density aerospace generators, such as the 2.5 MW Electrodynamics with a power density of 16 kw/kg paired with Siemens SP260D drive motors enables a completely detached primer mover drive unit architecture. A flexible electrical cord would permit longitudinal sliding. A 10% higher fuel burn will be incurred, assuming 95% efficiency for the motor and generator. A mass penalty of nearly 600 lbs would be incurred for the hybrid drivetrain, compared to less than 60 lbs for the driveshaft.
This renders a hybrid drivetrain option unattractive.

Leave a Reply

Fill in your details below or click an icon to log in: Logo

You are commenting using your account. Log Out /  Change )

Twitter picture

You are commenting using your Twitter account. Log Out /  Change )

Facebook photo

You are commenting using your Facebook account. Log Out /  Change )

Connecting to %s