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.
Christophe Pochari, Pochari Technologies, Bodega Bay, CA
In April 2019, Pochari Technologies initiated research into the possibility of extending the endurance of medium-sized rotorcraft by incorporating a liquid-hydrogen propulsion system in combination with existing gas turbine propulsion architectures rather than fuel cell electro-propulsion.
The orthodox viewpoint in the nascent hydrogen aerospace community is that fuel cells enabling electro-propulsion is one of the key advantages offered by hydrogen, therefore it should be exploited. It is argued that electric propulsion is somehow superior, and that combustion is obsolete. This viewpoint is backed arguments of the superior efficiency of electric motors and fuel cells over “Carnot limited” thermal engines.
On the surface the claim is correct, but upon deeper analysis, it fails to remain valid.
This view is easily challenged by carefully analyzing the current technology landscape. The proponents of electropropulsion often hype the potential advances in electrotechnology while often overlooking if not ignoring potential advancements in thermal propulsion technology. In the case of rotorcraft, due to historical industry conservatism lack of material option options recuperation technology has lagged behind and failed to be implemented in aero turbine engines.
According to Mcdonald 2008, turboshafts in the 1000 kw range with conventional metallic recuperators can achieve an SFC figure of 0.35, applying silicon carbide heat exchanger technology enables minimal mass penalty imposed by the recuperator that was found with conventional Inconel heat exchangers. Silicon carbide possesses a thermal conductivity of 10x that of stainless steel with one third the mass.
In contrast to thermal powerplant, PEM fuel cells currently have an attractive power density of up to 5 kw/kg, currently achieved by Horizon fuel cells. It’s been speculated that replacing the graphite bipolar plates, which account for up to 80% of the mass of a PEM fuel cell, could potentially raise that number to 8 kw/kg (Kadyak 2018).
These numbers appear to enable fuel cells to supersede turbomachinery for aircraft propulsion. Unfortunately, there exists a physical phenomenon inherent to hydrogen fuel cells which erode the main advantage of the fuel cell.
“The maximum efficiency of a PEM fuel cell is achieved at very low power levels, and efficiency decreases almost linearly as power increases” (Veziroğlu 1993). This statement obviously poses problems for aero propulsion!
Let’s estimate the weight of the hybrid drivetrain, assuming the numbers provided by Kadyk 2018 are correct, that is fuel cell power density of 8 kw/kg are attainable with thin metallic bipolar plates, as he does not provide an efficiency estimate, if we take 2.2 w/cm2 at an attainable conversion efficiency of 50% as a baseline (Watkins 1993), since 4 w/cm2 is needed to achieve 8 kw/kg, so we can reduce to 40% since power density has been doubled.
With the added weight electromachines, the net powerplant weight is well above a turbomachinery setup, with net efficiency being lower when converter and motor losses are taken into account, is below an optimized turboshaft, netting 36%.
With current non-superconducting DC motors, power densities of 5 kw/kg have been achieved by Siemens. The DC motor is very advantageous, as current DC/AC inverter technology has low power density, 1 kw/kg (Brombach 2012).
Since fuel cells output low DC voltage, a step up and is needed, for a DC/DC step up converter, the weight is very minimal, 60 kw/kg (Fraunhofer). The majority of the weight found in power electronics lies in the DC/AC converter.
Estimate of powerplant mass
Turboshaft (conservative with recuperation): 2.5 hp/lb
Fuel cell: 6.5 kw/kg, midpoint between future of 8 kw/kg and current of 5 kw/kg
Total powerplant including motors: 1.72 hp/lb
Electric powerplant net efficiency: 36%
As we’ve demonstrated, fuel cell powerplants are only attractive if efficiency is at least 50%, 40% at high power densities fails to offset additional powerplant mass.
With propulsion aside, the next exigency lies in optimizing the LH2 tankage system.
Advances in thermal insulation and carbon fiber materials enable lightweight tankage. Working in our favor, it has been demonstrated that woven carbon fiber composites show an increase in fracture toughness at cryogenics temperatures (Kalarikkal 2006). Carbon fiber shows a 1.5x increase in tensile strength at cryogenic temperatures (Okayasu 2019). Carbon composites provide ultra-lightweight tank mass, with cryogenic compatibility.
A thin metallic permeation barrier is used on the interior of the composite tank, as CFRP has high hydrogen permeability.
Lightweight vacuum load-bearing panel insulation with densities of 11 lb/ft3 with low thermal conductivities of 0.15 w/m-K at atmospheric pressure developed by Nanopore Inc enable conformal tankage, as vacuum shells for MLI require spherical configurations due to the high loads. With surface to volume ratios of approximately 1:1, boil-off rates of 5% per hour can be achieved, although the number may sound high, it is perfectly acceptable for medium-range aircraft as boil rates remain below 38% of cruise fuel burn.
The tanks can be configured to be integral with the rotorcraft’s rear fuselage to minimize structural weight. A woven CFRP skin of 0.15″ with ribs placed at 16″ intervals with VIP panels placed on the exterior forms a highly rigid and volumetrically efficient fuselage integral tank.
Tank weights of 45% LH2 mass can be achieved with this configuration and materials.
At airspeeds of 140 kts at sea level, flat plate skin friction drag of 0.2 lbf/ft2 of wetted area is used as an estimate for the additional parasitic drag incurred from the LH2 tankage volume. The additional drag results in a minimal increase in fuel burn.
During hover, the additional tankage slightly increases download, the increase represents 3.5% of the disc area for a 250 ft3 tank.
Rotorcraft specifications and potential range increase.
Powerplant: 1000 kw recuperated turboshaft
0.35 sfc @ 2.5 hp/lb
Rotorcraft weight: 8000 lb GW
Power needed: 1090 cruise hp, 136.25 lbs LH2/hr (LH2 120 MJ, Jet A 42.8 MJ/kg)
55% ew/gw ratio
Tank system: 45% wt CFRP fuselage-integral
Tankage tank penalty: Dynamic pressure @0.2 lbs/ft2: 3.5% downwash penalty
Range: 1200 miles
LH2 capacity: 1000 lb tank, 230 ft3, 450 lbs.
Range increase 1.9x
kerosene payload: 750 lbs
Hydrogen payload: 2100
Figure above illustrates the rear fuselage conformal liquid hydrogen tank
Figure below illustrates the download penalty
Credit: Christophe Pochari
Credit: Christophe Pochari
All renderings performed in Autodesk Fusion 360.
The rotorcraft like any other technology, has a certain number of subsystems, which together form a broader system.
All technologies that embody a certain degree of intricacy and convolution will generate a large number of relevant subsystems, each with their own unique characteristics and functions.
Each of the subsystems in turn can be optimized, enhanced and modified to further augment the expansive system. The expansive system can be viewed as the culmination of a plethora of specific functions, working synergistically, in a particular order to form a network of unique processes.
Each originally may possess a domain-specific categorization, but undergo sudden or gradual transition to fulfill a broader range of functions. The degree of single-task functioning contrasted to multi-task functioning can be broadly categorization as the “versatility” exhibited by the particular system component.
The processes within the expansive complex in turn are inexorably linked in a series of hierarchical commands.
As a stratified system forming a minimum level of symbiosis and compatibility constitute the essence of the expansive mechanism’s working principle.
The degree of deviation allowable by the specific subsystems can be categorized as the overall “flexibility” tolerated. This can be referred to as “plasticity”, or the ability of the expansive system to accommodate shifts in the order, motion, and pace of the relevant integrants.
Any particular digression from the premeditated course of action can potentially result in ruinous implications for either the particular components, or even the broader expansive system altogether.
One may note the similarity of this description to the cascading principle, where a gradual or sudden deterioration in the functioning of a particular integrant can cause a series of subsequent events in unrelated components, potentially compromising the entire expansive system.
In order to evaluate and assess an expansive system’s sensitivity to this phenomenon, we must evaluate the degree to which the operation of the expansive system is dependent upon, directly or indirectly, upon a linkage, or chain of subsystems, and the degree to which these subsystems suffer congenital vulnerabilities.
In turn, the degree to which components can operate alone, unimpeded from downstream failures can heavily influence the outcome, as the expansive system may be able to tolerate a certain portion of its integrants to remain inoperable, yet retain its ability to operate.
A ranking of the capacity of an expansive system to endure and cope with unanticipated spasmodic fluctuations in the customary activity can be ascertained by evaluating the mentioned above factors:
#1 The degree of reliance upon a series of downstream operations and cycles to achieve the final output. The operation itself may be derived of cycles and process, but who later converges into single function.
#2 The degree of interconnectedness of the individual components. The extent to which an individual component can achieve a particular function without delegating particular tasks to secondary components. A system can form a homogenous structure, by performing all tasks in an all-encompassing manner.
A heterogeneous system is forced by the nature of the particular procedure to consign portions of the procedure to highly specialized components, which by themselves provide only a singular function.
The expansive system’s integrants can be regarded as possessing a varying degree of intra-system functionality in a specific direction or route.
Tasks are performed through defined courses and trajectories. The capacity to deviate from a premediated route and embark upon a surrogate course while retaining a preponderance if its intra-system functionality can be regarded as a decisive factor in evaluating the permissible latitude in the expansive apparatus’s cyclic functioning.
A degree of desultory activity may be engendered upon the utilization of a non-customary itinerary. A minimum degree of plasticity must be ensured and provided to enable a transmutation to take place in a frictionless manner to curtail the unforeseen disruptions caused by variability in the itinerary quality and extent of compatibility with the contents of the subsystem.
Depending on the subsystem function, hermicity may be a requirement to enable the transfer of escapable susbtances, which may directly be performing kinetic functions or energetic transfers.
Partial or full hermeticity demands a temporary barricade to be erected to enable a blockage of substance transfer prior to the beginning of the itinerary transition to minimize of uncontrolled expulsion of reactive substances.
In the case of gravitationally resistive levitational systems, the criticality of ensuring continuous system activity without interruption imposes unique exigencies upon the design of the expansive, integrant, procedural and cyclic design parameters.
Furthermore, an expansive apparatus must generate and administer a segregated chain of commands specifically tailored to the unique exigencies generated by the particular system functions. This chain of command can be thought of as a highly elaborate organizational hierarchy.
This organization hierarchy can be structured in a fashion to ensure a constant flow of data and intelligence gathering by placing intermediate superintending devices within close proximity to operational processes which expend and generate commands which form sequential instructions needed to fulfill their operating nature.
The capacity to ensure uninterrupted communications between the different nodes within the apparatus is a momentous task when exogenous variables are taken into consideration.
Endogenous risk may be generated through internal processes which breakdown, generating hazardous byproducts which can potentially induce breakdown within the individual integrants.
It can be stated that the propagation of exogenous events form the basis for justifying of the strategic mitigation substructure which serves to militate against probabilistic exogenous variables that carry a high degree of inconclusiveness.
Inconclusiveness is the very nature of unpredictability, for this reason an exacting method to assiduously analyze incoming events is importune to form a comprehensive risk evaluation mechanism to perform selective elimination of incoming threatening actors.
It is probable upon further inquiry and analysis that the principle risk is manifested as primarily originating from external sources. These risks may be diverse in nature, possessing unique characteristics which may impose a selective danger to individual components. Certain components by the nature of their composition may be more immune to external degradation than others.
Components are evaluated and treated according to their congenital elementary compositional characteristics which be exploited to further buffer against exogenous risk.
The magnitude of the risk is evaluated based on the capacity to temporarily or permanently disrupt the operating procedures within the expansive apparatus.
It may be desirable to partially or fully encapsulate the apparatus within an isolating shield to achieve partial or full isolation from exogenous sources of compromising actors. A selective barrier can serve as a buffer between sensitive components and their respective source of endangerment.
A hierarchical distribution of derivate functions is analyzed based on an imput-ouput directional perspective.
Gravitational resistive leviation systems can be broadly categorized as being domain-specific or multi-faceted.
Each constituent within the expansive complex as demonstrated earlier performs a domain-specific function. It can be thought as performing within a narrow well-defined context, with clearly set parameters.
The scope of operation, degree of plasticity, was defined as the degree of versatility exhibited by the initial and subsequent functions.
Components serve to generate a directional series of increments, each forming an iteration within a broader package of processes. The direction can be single-pathed or multi-pathed. The number of itineraries admissible is determined by the variability of the carriers within the pathways.
Vertical lift componentry can be broadly classified as falling under the category of kinetic motion systems. Ancillary factors may necessitate additional forms of energetic transfer methods, specifically tailored to enable transmission of kinetic motion through a unique set of obstacles imposed by the inherent complexity found within the expansive complex.
Kinetic and inertial motion can be interpreted as being the derivative of a series of thermal inputs.
Thermal energy is transmutated into kinetic energy via a series of chemical imputs. The biproduct takes shape as a divergent form of energy from the initial constituent of thermal radiant energies.
The derivate is a form of kinetic motion, comprised only of molecular mass, devoid of mechanical or dynamic energy, possessing only embryonic energy.
The intermediate product of kinetic molecule compounds can be thought of as being antecedent to the mechanical kinetic motion derived in the final step.
It serves as only the precursor to a higher form of kinetic energy, yet to have entered the final and stage of stage of transformation.
The multi-step transformation of energies commences as a stable chemical compound and can be characterized as following a progressive trajectory.
The molecular compounds are chosen based on their propensity to engage in a series of atomic level excitation events induced by initial thermal precipitation. Precipitation is required to impose vibratory energy upon the molecular compounds to initiate the reaction event.
The formation of energies can only take place if the needed quantity of reactant
pairs can be sourced. As heavier than air flight vehicles are constraints by the gravimetric to propulsive ratio, the derived source of reactant partners are sourced from the copious supply of ambient compounds circulating throughout the atmosphere.
The reactions which take place are a series of events that can be classified as an oxidative process, consisting of a series of molecular interchanges, precisely described as series of atomic transmutation. The commencing mass content is preserved through the series of atomic interchanges.
An intrinsic and universal constant state of foundational energy can be thought of as being possessed by a series of electrical and magnetic attachments forming elemental compositions that, under exacting conditions, prompted to initiate chain reactions of constant or periodic discharges of potential energy.
The derivatives of the process are expelled from the enclosed reaction chamber enabling a mass-accumulation free process.
The geometric orientation of the complex exists as an aggregation or cluster of individual geometries forming a near uniformly distributed homogenous geometry whose shape is determined by its internal component configurations.
There exists a finite number of elementary or fundamental principles that govern the design and configuration of the expansive systems within the broader vertical lift complex.
These principles can be classified as being governed by fundamental physical laws imposed by natural forces. These stem from organic, thermal, kinetic and chemical properties each producing unique integration and compatibility exigencies forming the basis of design choices.
The second category can be broadly referred to as being manmade in nature. This category can include a wider scope of particular design configurations influenced and governed by the former principles.
The latter principles form a nearly infinite number of design possibilities and variables each influence in part by their relationship with their correlative members. As discussed previously, a systematic order of operation is formed largely by natural forces that impose a strict operational procedure for both the individual integrants and the broader expansive complex as a whole.
An immutable form of consanguinity exists between the constituents within the scope of procedural processes as a byproduct of the commencing requirements.
The order and sequence of operational procedures are contrived based upon strict adherence to the elementary principles which govern the admissible blueprint.
The degree of transgression and infringement permitted to take place from the guiding blueprint is conditional upon the congenital level of the rigidity imposed by ancillary factors derived from exogenous forces beyond the scope of alternation enabled by current techniques.
The origin and nature of exogenous forces produced by the intended operating environment and terrain is unpredictable in nature.
The intended operating scope of the vehicle has to be properly defined in order to ascertain the necessary information to perform domain-specific optimization of the respective componentry and to insure intra-component consanguinity.
The degree of deviation from a defined operating environment is influenced by the extent to which the expansive complex can acclimatize sudden variations imposed by natural fluctuations which have the propensity to take place within the unpredictable ambiance.
I’ve made some recent alterations to VED technology for helicopters. These changes include stretching the capsule from 36″ to 75″ to accommodate a stretcher in the longitudinal direction being offered as an option. In previous versions of VED for EMS operations in mind, the capsule was designed with a stretcher configuration perpendicular to the length of the helicopter. This meant the stretcher, typically being long enough to accommodate a tall person, had to protrude through the doors on each side of the capsule, provisions for this include specially designed doors with a “bubbles” on each side. Although the size of the capsule is nearly 70″ long, only minimal accommodation was needed. This design also posed some additional limitations. The previous design restricted the ability of the medics to freely move between the front and rear section of the fixed portion of the fuselage. This could potential hinder the medics from performing interventions and monitor the patient on both sides of the stretcher. As such, two versions of VED technology will be offered, with the difference being solely the length of the capsule. The image below depicts a VED fuselage configured with a 36″ wide capsule.
Silane is one of the few flammable compounds that readily combusts in a pure CO2 atmosphere producing a pure solid carbon byproduct.
Silane emerges as an attractive solution to dispensing with captured carbon dioxide. Currently, there exists no practical method to convert carbon dioxide to a more easily stored substance.
The principle of this cycle is combusting silane gas in a CO2 or CO2/argon atmosphere, using the CO2 as an oxidizer, producing energy via a supercritical CO2 turbine, emitting solid carbon and silicon dioxide, then separating the silicon dioxide from the solid carbon, to then reuse for other applications.
98% metallurgical grade silicon: $1500/ton, 0.85 per ton of silane
Potential silane cost including reactor capex, chlorine and hydrogen consumption: $3000-3200/ton
1 ton of silane = 2.75 tons/CO2 consumed
Biproducts to resale
1.87 tons silicon dioxide at $750/ton
0.75 tons carbon powder at $700/ton
Total downstream revenue potential: $2530
Potential price per ton of carbon dioxide reduced: $180-250 (no comparable method available to compare cost)
Trisilane is an ideal carbon free fuel as it’s a dense liquid at room temperature and pressure (740 kg/m3), burns efficiently and possesses high energy density (40+ MJ/kg). The only biproduct of combustion is silicon nitrade, a solid. Triilane is ideal for use in external combustion cycles such as S-CO2 Brayton cycles. Heavy duty propulsion applications such as marine and rail require an energy density liquid fuel, the issue is all current carbon free fuels suffer from storage and energy density constraints. Tisilane offers an ideal carbon free propulsion option for heavy duty propulsion. Trisilane is compatible with current liquid infrastructure and holds the title as being the only carbon free liquid fuel, excluding hydrazine, which is considered too toxic. All other carbon free fuels are gaseous or require carbon recycling. Potential power density of supercritical S-CO2 bottoming cycles are 1.4 MW/Ton.