Compression loaded ceramic turbine disc technology.

Pochari Technologies is reviving long-forgotten turbine engine technology. With advancements in single-crystal nickel alloys, turbine designers have been able to increase turbine inlet temperatures to record levels. The downside of this is the need to utilize large amounts of compressor bleed air, reducing engine efficiency. Conventional engineering ceramics including silicon nitride and carbide have been forgotten in part due to the focus on ceramic matrix composites. A relatively elegant and simple design is the use of a composite (carbon-fiber) hoop holding the centrifugal forces acting on the blade. Conventional engineering ceramics have uneven properties, rendering them unable to tolerate concentrated forces found in conventional mating conditions use in current turbine disc designs. Contrary to popular belief, engineering cermaics do indeed have sufficient tensile strength to be loaded in tensile, such as a beam for example. Silicon carbide has a tensile strength of up to 230,000 psi, more than sufficient to tolerate the tensile loads found in a turbine blade root connection. The primary issue is the lack of uniformity, the material properties are highly heterogeneous, in addition, low flexural strength, and high brittleness further impede its use. The simple (in hindsight obvious), solution is to load the blades purely in compression, (where ceramics shine), in fact, silicon carbide has one of the highest compressive strengths available, 1600 MPa. In order to take advantage of the excellent high-temperature capacity of these materials, a novel architecture is needed, dispensing the existing orthodoxy of turbine disc design. Using this design, turbine inlet temperatures approaching 3000 F are feasible without air cooling. The blades transfer compression loads to hoop loads, the carbon fiber rim performs the same function as a pressure vessel. The carbon fiber parameter containment hoop is not exposed to high gas temperatures. With this technology, it is possible to design small-scale sub-500 hp turboshafts with the efficiency of diesel engines (40%+), enabling jet-pack propulsion with 1 hour plus range.

“Ceramic materials offer a great potential for high-temperature application. This,
however, means it is necessary to live – even in future – with a brittle material with
small critical crack length and high crack growth velocity. Thus it will not be easy to
ensure reliability for highly loaded ceramic components, keeping in mind that for
reaction bonded ceramics the material inherent porosity is in the same order of
magnitude as the critical crack length. A solution to increase the reliability of ceramic
turbines may be a compression loaded rotor design with fiber reinforced hooping”

R. Kochendrfer 1980

“A vaned rotor of the type comprising a central metal hub or rotor body carrying a plurality of rotor blades made of a ceramic material, in which the blades are simply located on the rotor body and held in place by a coil of carbon fibres or ceramic fibres which surrounds the blades. To form a support surface for the coil each blade has a transverse part at the radially outer end thereof, which is partly cylindrical and which together with the transverse parts of the other blades, forms a substantially cylindrical support surface for the coil. Although ceramic materials used for such vanes (silicon nitride, silicon carbide, alumina, etc.) have much better physical properties at high temperatures (i.e., over l,lC than any metal alloy, especially if undergoing compression loads, they are nevertheless very difficult to couple to metal parts because of their relative fragility, lack of ductility, and their low coefficient of expansion.Because of the lack of ductility of ceramic materials, the driving forces exerted during operation of the rotor give rise to a concentration of the load in parts of the coupling areas between the ceramic vanes and the metal body of the rotor. This frequently causes breakages in these parts. The various systems presently in use for attaching a ceramic blade by the root to a metal rotor body for a gas turbine are generally inadequate because these systems, including dovetail fixings having both straight and curved sides, do not take sufficient account of the rigidity and relative fragility of the ceramic vanes.This problem is exacerbated by the fact that present manufacturing techniques for ceramic materials are still not able to provide a complete homogeneity of composition and structure of the material, so that adjacent areas of ceramic material can vary by up to 200% in tensile strength. For this reason the known types of coupling between a support disc forming a rotor body and rotor vanes of ceramic material, which rely on a wedging action, are not satisfactory”

R Cerrato  Fiat SpA, U.S patent 3857650A, 1973

“A Compression Structured Ceramic Turbine looks feasible. A new engine aerodynamic cycle with effective working fins to off set windage loss, a reduced tip speed to enhance aeromechanics and
the possible utilization of leakage gas to augment thrust should be considered. Also, the prospect for more efficient energy extraction offered by inverted taper in the span of the turbine blade should be of prime interest to turbine designers in any future engine utilizing a Compression Structured Ceramic Turbine. Material property data and design refinements based on this data will also have to be seriously considered”

“The “Novel” feature of this ceramic turbine rotor design involves maintaining the ceramic
rotating components in astate of compression at all operating conditions. Many ceramic materials being considered for gas turbine components today display compressive strengths ranging from three to eight times their tensile strengths. Utilizing the high compressive strengths of ceramics in gas turbines for improving ceramic turbine structural integrity has interested engineers in recent years as evidenced by a number of patents and reports issued on Compression Structured Ceramic Turbines with one as early as 1968. Turbine blades designed to be in compression could greatly enhance the reliability of the ceramic hot section components. A design of this nature was accomplished in this contractual effort by using an air-cooled, high strength, lightweight rotating composite containment hoop at the outer diameter of the ceramic turbine tip cooling fins which in turn support the ceramic turbine blades in compression against the turbine wheel. A brief description of the detailed structural and thermal analysis and projected comparable performance between the Compression Structured Ceramic Turbine”.

P.J Coty, 1983

Ultra-low cost Alkaline electrolyzers using commercial-off the shelf (COTS) cost reduction methodology


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Alkaline electrolyzer technology is ripe for dramatic cost reduction. Current alkaline electrolyzer technology is excessively expensive beyond what material costs would predict, mainly due to very small production volumes, a noncompetitive market with a small number of big players, and relatively little use of COTS (commercial off the shelf) methodology of cost reduction.

Pochari Technologies research has thus applied this methodology to bring to market affordable hydrogen generators fabricated from readily available high-quality components, raw materials, and equipment procured on

An alkaline cell is a relatively simple system, consisting of three major components. The electrode (a woven wire mesh), a gasket (made of cheap synthetic rubbers, EPDM etc), and a material for fabricating the diaphragm membrane for separating and oxygen and hydrogen while permitting sufficient ionic conductivity (usually polypropylene sulphide or Polytetrafluoroethylene

(TEFLON). The fourth component are the “end plates” which consist of heavy-duty metallic or composite flat sheets which hold a series of rods tightly pressing the stacks to maintaining sufficient pressure within the stack sandwich.

Unlike batteries, noble mineral intensity is alkaline technology is relatively small, with nickel mesh loading of under 500 grams/m2 of active electrode surface area needed to achieve anode life of 5 or more years assuming a corrosion rate of below 0.25 MPY. With current densities of 500 miliamp/cm2 at 1.7-2 volts being achievable at 25-30% KOH concentration, power densities of nearly 10 kW/m2 is realizable. This means a one megawatt electrolyzer at an efficiency of 75% (45 kWh/kg-H2 LHV) would use 118 square meters of active electrode surface area. Assuming a surface/density ratio of a standard 80×80 mesh, 400 grams of nickel is used per square meter of total exposed area of the mesh wires. Thus, a total of 2.25 kg of nickel is needed to produce 1 kg of hydrogen per hour. For a 1 megawatt, the nickel would cost only $1000 assuming $20/kg.

For a lower corrosion rate of 1 um/yr, a total mass loss of 7% per year will occur with a surface/mass ratio of 140 grams/m2-exposed area, the nickel requirement is only $350 or 17.5 kg for one megawatt! Although this number is achievable, higher corrosion rates will likely be enounctered. To insure sufficient electrode reserve, a nickel loading of

Power density of 6.8 kW/m2 electrode area.

For the diaphragm separators using a half a milimeter thick sheet of Polytetrafluoroethylene, around 88 grams is used per kilowatt, at a typical cost of PTFE of $9/kg, around $0.79/kilowatt can be expected assuming an electrode power density of 6.8 kW/m2 (400 miliamps at 1.7 volts)

Gasket costs are virtually negligible, with only 4.8 grams of rubber needed per kilowatt.

For 30% NaOH at 117 C, a corrosion rate of 0.0063 millimeter per year (0.248 MPY) is observed for an optimal nickel concentration of 80%. This means 55 grams of Ni is lost for one square meter, if we choose 10% per year as an acceptable weight loss, we return to 550 grams per square meter as the most realistic target nickel loading, with much lower loading achievable with reduced corrosion rates. A lower concentration of KOH/NaOH and lower operating temperature can be utilized as a trade-off between corrosion and power density.

Hydrous hydrazine for fuel cell transportation

Hydrous hydrazine as a superior alternative to high-pressure gaseous hydrogen storage fuel for current PEM fuel cell powered applications. 

Hydrazine is not frequently mentioned in low carbon energy circles. Despite this, hydrazine holds the title as being the only liquid hydrogen carrier that does not contain carbon atoms. 

Hydrazine, being a liquid at room temperature, allows for safe and easy transport, storage and dispensing. Hydrazine is always mixed with water, between 29%-64% hydrazine by weight in water, as anhydrous hydrazine is highly unstable. Ordinary plastic containers are used to transport hydrous hydrazine, conventional tanker trucks can be used and existing fuel station infrastructure can be utilized with modifications limited to the fueling nozzles.

Hydrazine can be produced from hydrogen peroxide using a new and improved process invented in the 1970s. The feedstock required is hydrogen peroxide and ammonia. 2.4 mols of ammonia plus 1.4 mol of hydrogen peroxide react with a hydrocarbon, methyl ether ketone, to eventually synthesize into one mol of hydrazine, generating 2 mols of water. The ketone is recycled and not consumed in the process. The cost of producing hydrogen peroxide from the anthraquinone process is $227/ton for a 100% solution for raw materials and energy input. 

The raw material costs for producing hydrazine hydrate excluding the original hydrogen feedstock for ammonia through the peroxide process is approximately

$6.66/kg of H2 based on current market prices assuming a price of $1.5/kg for hydrogen for the manufacturing of the hydrogen peroxide, this is roughly the cost of producing hydrogen from steam methane reforming which supplies most FCEVs today supplied with 350-700 bar technology. The cost of $6.66 is the cost above and beyond the hydrogen feedstock, as it is to be compared to the cost of transporting, compressing and dispensing hydrogen. The capital costs of the plant depend greatly on the production volumes, of which currently is very low due to the small demand for hydrazine. In ammonia synthesis, for example, 90% of the cost is the hydrogen feedstock, with the balance being the cost of the plant. Hydrazine production requires no compression and minimal temperature, conditions are very mild, 10 bar and 150 C.

Hydrazine is a very dense hydrogen carrier, containing 12.6% wt hydrogen. In a hydrous solution, it is up to 8%. Below 40% concentration in water, hydrazine vapors are not flammable. The volumetric density of hydrous hydrogen is up to 84 kg/m3, 4x higher than 700 bar storage. 

Hydrazine for fuel cells.

Current PEM fuel cell technology can be used with hydrazine supplied hydrogen. With the proper catalyst, preferably bimetallic rhodium and nickel at 4:1 ratios Ni/Rh, hydrous hydrazine readily decomposes into pure hydrogen and nitrogen at room temperature. The mixture can then be fed directly into the PEM fuel cell.

Hydrous hydrazine has a comparable risk profile to methanol, with flammability risk being much lower than its toxicity risk. Hydrazine has not been shown to be carcingeic. Below 40% solution in water, it has no flash point, meaning it can not be ignited, eliminating all safety concerns that hydrogen faces due to its tremendous flammability. 

In summary, hydrazine is safer than gaseous hydrogen owing to its low flammability as well as liquid nature, which minimizes the probability of direct contact. 

In summary, hydrazine offers the ability to develop a safe liquid “hydrogen economy” bypassing all the safety and infrastructure issues faced by current high-pressure gas storage technology.

“Don’t try to store hydrogen by compression or liquefaction, I’ve said it and will say it again, chemistry is your friend! A few chemical reactions and some rearranging of atoms can do wonders”

Small scale Birkeland Eyde reactors

Pochari Technologies is developing gliding arc plasma reactors based on the 120-year-old Birkeland Eyde process. Efficiencies are estimated to be 10 kwh/kg NOx based on non-thermal gliding ac plasma instead of 20 kw/kg NO2 using thermal plasma generated by the Birkeland Eyde electrodes.
With micro-nitric acid reactors, farmers can produce their own fertilizer without carbon emissions cost-effectively.

Photovoltaic technology has witnessed tremendous improvements in cost reduction in large part imputable to the dramatic decrease in the cost of polysilicon, from over $400/kg in 2008 to as little as $10kg in 2019. Modern photovoltaic systems wholesaling on Chinese marketplaces such as Alibaba sell for as little as $0.18/watt from RISEN ENERGY CO., LTD for monocrystalline architecture. Much as hydropower enabled the viability of the Birkeland-Eyde process in the early 20th century, modern-day photovoltaic technology with LCOE’s of under 2 cents/kWh opens up a world of opportunities for the electrification of ammonia production and other energy-intensive chemical resources.Degradation rates are typically around 15% for 20 years, or around 0.8% per anum. That means a 1 kW system will produce 84% of its original power output after two decades.Panel type: 275-280/330-335W Multi-Module Price per watt (USD): 0.28 High, 0.175 Low, 0.185 Average.The average price for 350-watt panels is 18.5 cents per watt.The second major cost input is the DC/AC converter. Using data from Alibaba, numerous products were sampled. The most cost-competitive DC/AC rectifiers were ones used for solar-powered well water pumps. 7.5 kW units were priced around $250, yielding a price per kilowatt of $34.The Levelized Cost of Energy (LCOE) is determined by the irradiance available much more so than it is by slight differences in the panel module costs. A solar array in Scotland (880 kWh/kWp/yr) won’t be nearly as cheap as one in Chile (2300 kWh/kWp/yr), or in Los Vegas (1900 kWh/kWp/yr). Using this data, we can estimate the cost of producing NH3 in a high irradiance region such as Los Vegas using a 300 kW unit producing 609,000 kWh per year. The size of the farm would include 15,000 square feet of panels. A 50 kW PV system at B-E efficiency levels translates into about 11 kW installed electrical capacity, although since PV output is cyclical, power is only produced from sun-rise to sun-set, so most of the power is concentrated during a short window. To get around this limitation, the power can be broken down into multiple smaller reactors that can each be switched on to adjust to the rising and falling power output.

Potential system cost. Bare Panel system: $9,250 Invertor: $1700 Wiring: $500 Metallic frame: Not included (insignificant) Land: Not included (assumed to be already owned)

Total: $11,500Price per kWh for one year: $0.12 Price per kWh for 20 years: $0.0071

Price per ton-HNO3 at original B-E efficiency: $108Price per ton-HNO3 at improved non-thermal plasma efficiency: $80Prevailing U.S market price of HNO3: $350-400.CAPEX cost of reactor:

The main CAPEX of the reactor system is the electrical transformer. I recently purchased two 3.5 kW transformer cores for $65/each, not including $300 in shipping. These cores weigh 16 kg each and are made of ferrosilicon alloy. The power density is 0.22 kw/kg. Adding the cost of enameled wire, I bought 10 meters of 1.2 mm transformer wire for $6 on Aliexpress. If the work is done yourself and the products purchased from Chinese marketplaces, what would otherwise be an expensive item if procured conventionally from a specialized supplier is in actuality a cheap item. Therefore, A 10 kW transformer is potentially a very cheap component. What about the electrodes? For the non-thermal design, the plasma reactor consists of two knife-shaped electrodes. To maximize conversion efficiency, the volume of air exposed to the plasma must be maximized, the reactors are thus more efficient at smaller scales. Rather than a cylindrical design where 4 or more electrodes are arranged in a circular pattern, two are placed inside a thin flat box with quartz covering the top of the electrode. The free air space is thus minimal and most of the air that passes through is exposed to the plasma flame. This completes the reactor, all that is left is the absorption system. Arguably, the absorption system is the most complicated and potentially costly. Plasma synthesis of NOx was demonstrated to be 3x more efficient in the kHz range than in the 50-60 Hz range. What this means is we can use the DC current from the solar panels directly to the rectifier to produce the necessary high-frequency AC power needed for the plasma module. The cost of a 1 kW rectifier PCB circuit module outputting 20 kHz is $30. The rectifier outputs 400 volts, so a high-frequency ferrite core can then be used to increase the voltage to around 10-15 kV, voltage above ten kV is unnecessary. The main disadvantage of the non-thermal plasma is the power density. While the original Birkeland Eyde furnaces used highly intense thermal plasma discs stabilized by electromagnets, non-thermal gliding arc discharges (GAD) suffer from lower power densities despite noticeably higher efficiencies. This is due to the difference in activation mechanisms. While the thermal process takes advantage of the non-linear thermal Zeldovich mechanism, whereby high-temperature breaks down the diatomic nitrogen and oxygen, create N-O radicals which immediately form, non-thermal plasma employs the electron excitation principle for activation.

“Nitric acid is one of the most important inorganic acids and it is used in the
production of fertilizers, dyestuffs, and resins. Further applications are stainless steel pickling and metal etching. About three-fourths of the nitric acid produced is used in the fertilizer industry, mainly for the production of ammonium nitrate, ammonium phosphates, and compound fertilizers. The nitric acid needed in the fertilizer industry is usually diluted nitric
acid with a concentration of 50-70%. For most other applications, such as
nitration reactions, 90-100% nitric acid is used”

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Personal jetpacks using ultra-high power density compression-ignition low SFC engine technology

Personal jetpacks using ultra-high power density compression-ignition low SFC engine technology

Christophe Pochari

Man has dreamed of micro-air vehicles that are highly compact, essentially fitting around a single person, a flying backpack, that enables him reach the most challenging and confined geographies that even helicopters cannot tackle. Conventional aircraft have large footprints and exposed rotors, even recent EVTOL designs all feature not only large physical footprints but exposed rotors which risk sustaining foreign object damage in highly confined environments. While the jetpack has tremendous application in a wide range of end uses, such as surveying, surveillance, law enforcement, defense, etc, the Jetpack is not simply a tactical or utilitarian technology, it is rather an attempt to attain complete freedom of mobility, to unleash man from his bondage to the earth.

Jetpack technology has been historically handicapped by poor propulsion efficiency. While Martin Jetpack had solved this by using ducted fans, the Martin Jetpack suffered from poor prime mover power density. While the propulsion efficiency was stellar, the engine used had high SFC and poor power-weight. Few people realize modern diesel engine technology can achieve enormous power density. Decades of NASA research on compound cycle diesel engines originally studied for rotorcraft propulsion achieved power densities of 1.66 lb/hp with SFCs of 0.32 using advanced turbo-compounding. The research was abandoned as increasing turbine inlet temperatures permitted by the advancements in single crystal alloy turbine blades overshadowed the best efforts of reciprocating propulsion. Fast-forward 40 years since the 1980s NASA/Garrett turbo-compounding program and we find ourselves needing high power density low-SFC propulsion more than ever for advanced micro-air vehicles to achieve desired endurance requirements. Turboshaft propulsion cannot be scaled down efficiently below 500 hp within an acceptable SFC range. More recently, SMA, a subsidiary of Safran, designed a 530 lb direct-injection compression ignition engine producing up to 800 hp. This does not include potential increases through turbo-compounding. NASA’s original estimate was 1.6 hp/lb.

The second major enabling technology are light-weight ballistic low-altitude parachute systems. This technology has come a long-way and is poised to make Jetpacks as safe as conventional air vehicles.

Micro tactical air vehicle with 2 hour endurance 750 lb gw

Propulsive system: High speed CFRP ducted fan 215 mm (20,000 rpm) 232 lbf/ft2 38.5 lbf/lb 2.5 lbf/hp 3x

Prime mover: 240 kW 2 stroke diesel engine, 0.35 lbs/hp-hr (2 hp/lb, currently achieved is 1.5 hp/lb by SMA)

Transmission system: Epicyclic titanium gears with CFRP casing

Weight breakdown:

Ducted fan shroud and blades: 21 lbs

Fuel tank: 15 lbs

Fuel: Exo-Tetrahydrodicyclopentadiene 0.976 g/cm3 42 MJ/kg

CFRP frame: 30 lbs

Seat 6 lbs

Belt and shafts: 10 lbs

Radiator: 7 lbs

Coolant: 4 lbs

Electronic flight controls: 5 lbs

Powerplant 340 hp: 170 lbs

Ballastic Parachute: 35 lbs

Fuel load: 178 lbs

Empty weight: 298 lbs

loaded weight: 476 lbs

Payload: 274 lbs

*Note: picture contains an turbine engine for reference, a compound cycle diesel engine has similar volumetric density.

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.

The potential of liquid hydrogen fuselage integral vacuum panel insulated carbon fiber cryotanks with recuperated turboshafts for ultra long range rotorcraft

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%

Turboshaft: 38%

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
Increase 2.8x.

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Figure above illustrates the rear fuselage conformal liquid hydrogen tank

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Figure below illustrates the download penalty