High payload long-endurance jetpack

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Technological Assessment of Jetpack technology using modern recuperated ceramic gas turbines and high efficiency ducted fans

Christophe Pochari, Pochari Technologies

The helicopter has been the benchmark of vertical flight for nearly a century.

While the helicopter’s hover efficiency and stability are impressive, it feature many characteristics that make in inapposite for a new generation of applications.

One of the single biggest limitations of the conventional helicopter is its large footprint, with wide rotor blades, these machines are not suitable to operate in urban and confined areas. Even small helicopters have rotor diameters over 20 feet, making it highly unsafe to land around residential homes.

The Pochari Technologies PortaPack air vehicle has a footprint no bigger than four feet by five feet, it can practically land on a balcony on a high rise apartment, hover right up against a high rise building rescuing occupants, or loiter up rubbing against a cliff side to rescue stranded hikers.

Ducted fan technology has been around since the 1950s, the principal of generating lift using condensed propellers within a duct which acts to concentrate the airflow is a proven concept.

What strongly favors high disk loading VTOL technology is the non-linear relationship between disk loading and power loading.

For example, a 250mm EDF can generate 123 lbs of thrust using 60 hp, translating to 236 lbs/ft2 of disk loading a 2.05 lbf/hp.

In contrast, a Robinson R44uses 225 horsepower to generate 2400 of lift with 856 square feet of rotor area, yielding only 2.8 lbf/ft2 and 10.66 lb/hp.

This means 85 times more thrust can be generate in a given area with only a 5.2 times increase in power, strongly favoring higher disk loading.

A 15.5 times increase in the power loading increases disc loading by a 137.5 times.

Clearly, if one desires compactness, one should be able

With this in mind, it makes little sense to design ultra low disk loading helicopters since we have to accommodate such a large area.

The F-35 liftfan generates 20,000 lbf of thrust in a 50 inch diameter fan, generating 0.689 lbf/hp and 1466 lbf/ft2.

Sabrewing Aircraft Company claims to able to achieve a disk loading of 122.4 lbf/ft2 at a power loading of only 2.81 lb/hp with a 5.2 ft diameter ducted fan.

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.

Recuperated gas turbine technology is critical if not enabling to for jetpack technology.

The Russians achieved an SFC as low as 0.348 in a 500 hp turboshaft engine using an 80 lb recuperator. Saturn/Lyulka AL-34 produced 550 hp in cruise and a 1000 horsepower in takeoff. Up to a 28% reduction in SFC can be enabled with advanced recuperation. The lowest SFC gas turbine in the sub 1000 hp class is the HTS900 with an SFC of 0.52 lbs/hp-hr.

Heat transfer rates in a microchannel SiC heat exchanger could approach 58 kW per m2 with a tube section of 0.6mm.

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.

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 since a considerable portion of compressor mass flow is unavailable for combustion. 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 all of the centrifugal forces acting upon the blade. Conventional engineering ceramics have morphologically heterogenous 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 240,000 psi, more than sufficient to tolerate the tensile loads found in a turbine blade root connection. The tensile strength of Cannon Muskegon X-10 superalloy at 1000 C is around 108,000 psi. The primary issue is the lack of uniformity, engineering ceramics are immanently omnifarous, in addition, low flexural strength, and high brittleness further impede their 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 entirely 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 2 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

Helicopters are noisy, crash prone, and excessively bulky.

Jetpacks offer a miniaturized flight vehicle that can be used in urban, hazardous and confined areas.

The key technology needed to faciliate jetpack viability is not propulsion, it is in the domain of the powerplant.

Exonetik Inc, a Canadian company, claims they are designing a 50% efficient recuperated turboshaft engine with a power density of 5 kW/kg.

System weights:

The primary mass component of the jetpack system is the fuel load, comprising 35% at a three hour range, and 52% at the four hour range mode. with the second largest mass item being the powerplant. After the powerplant, the structure dominates.

The Pochari JetPack design makes use of an innovative integral fuel tank/fuselage design where the fuel tank structure serves as the structural member for holding the operator and thrusters into an integrated fuselage package.

Fuel tank/structure: 29.98 lbs

Side rail: 14.3 lbs

EDF weight: 3.4 kg

66 lbs

Vasyfan in Italy was able to generate 79.5 kg of thrust from a 395mm diameter ducted fan using 45 kilowatts of power.

Vasyvan also achieved 32.7 kg of thrust from a 174 mm ducted fan drawing 26.8 kW, translating to a net mechanical power loading of 2.18 hp/lb and a disc loading of 281.5 lbf/ft2.

35% Brake thermal efficiency at 2800 F Turbine inlet temperature 4.2 kW/kg, 2.55 hp/lb

117.5lb/ft 3.5 hp/lb

Lilium disc loading 1585lb GW 136.3 lbf/ft2 320 430 hp kw 2.91 lbf/hp (motor controller efficiency 90%, 3.14 lbf/hp)

390 mm ducted fan 175 lbf 1270 lbf 230 lbf/ft2 2.7 hp/lbf jetpack 2 hour range

511 hp

fuel burn: 201.8 lbs/hr

4 hour range:

634 lbs

Fuel: 807 lbs 13.5 ft3

Max fuel capacity: 831 lbs

Powerplant: 100 lbs

EDF: 28 lbs

Parachute 650 kg capacity 12.3 kg weight 27.10 lbs

System weight: 130 lbs

Drive package: 20 lbs

Loaded weight: 884 lbs

Gross Payload: 701 lbs

Payload with max range:

Payload minus occupant: 521 lbs

GALAXY GRS s.r.o from the Czech republic manufacturers light weight and high capacity ballistic parachutes that fit into a highly compact package.

Pochari 850 horsepower recuperated gas turbine

Insulation: Excelfrax 1800 Board 0.036 W-mK @800 C 230 kg/m3

Silicon carbon high surface area plate heat exchanger/recuperator: 30 lbs

Low pressure turbine wheel: 5.6 lbs

High pressure SiC compression loaded wheel: 5.79

Medium pressure SiC compression loaded wheel: 6.137 lbs

Housing Al-Li/Ti 6Al-4: 35.97 lbs

Shaft: 5.478

Low pressure 7075 Al centrifugal impeller: 3.45 lbs

High pressure Ti impeller: 4.87 lbs

Combustor Nimonic 90: 8.5 lbs

Rear Nimonic heat shield: 1.86 lbs

Bearings: 1.86 lbs

Titanium bolts: 1.2 lbs

Fuel pump: 5 lbs

Total: 110 lbs

Power density: 6.4 hp/lb

Redundant single power turbine blade

Fuel capacity: 959 lbs

Disk loading:

Max thrust: 1731 lbs

Power: 890 hp

Hourly fuel burn: 343 lbs

Hourly power reduction: 175 hp

Hourly reduction in fuel burn: 68 lbs

Max range: 2.9 hours

Hour 0.5 1700 lb 170 lb burned

Hour 0.5 152 lbs burned

hour 2 275 lbs 1380 lb GW

Hour 3 1080 lb GW 220 lb burned

Hour 4 171 lbs burned

Total 988 lbs

Engine mass: 130 lbs

Fuel tank: 21 lbs

EDF: 32 lbs

Parachute: 27 lbs

Drivetrain belt drive: 30 lbs

Loaded weights: 240 lbs

Small LiftPack

1267 lbf max lift

1180 lb gross weight

590 hp

hourly fuel burn: 1 230 lbs 2 185 3 149 lbs

Total: 564 lbs

Engine: 85 lbs

Parachute 30

EDF: 36 lbs

Fuel tank: 20 lbs

Gears: 10 lbs

Loaded weight: 730 lbs

Payload: 450 lbs

GW: 750 lbs

Fuel volume: 3.9 cubic feet: 215 lbs

Power loading: 5.5 lbs/hp

Disk loading: 68 lbs/hp

Power: 136 hp

engine weight: 90 lbs

Compounded cycle diesel engine 0.33 lbs/hp-hr bsfc 2 lbs/hp: 121 lbs

System weight: 110 lbs

Loaded weight: 415 lbs

Payload: 335 lbs

1st hour fuel burn: 54 lbs

2nd hour fuel burn: 44.88 lbs

3rd hour fuel burn: 42.3 lbs

4th hour fuel burn: 39.7 lbs

5th hour fuel burn: 34.67 lbs

Electric version

Generator: 32 lbs

Motors 35 lbs

3 minute battery reserve: 40 lbs

efficiency penalty: 40 lbs

Additional weight of electric powerplant: 147 lbs

Short range micro jetpack with movable thrusters

GW 430 lbs

430 LB GW: 94 hp

637 LB GW: 202 hp

135 lbs

62 lbs

Disk loading: 91 lbs/ft2

Power loading: 4.6 hp/lb

Generator: 10 kw/kg 18 lbs

30,000 rpm SiMo Motors: 10 kw/kg 19 lbs

Rectifier: 3.5 lbs

Electric motor RPM increasing gearboxes: 8.3 lbs

Bevel gears: 3.5 lbs

Shafts: 1 lbs

Ducted fan fuel tank integral structure: 20 lbs

Magnesium ducted fan blades: 10.3 lbs

Radiator: 3 lbs

Fuel load 1st hour: 31 lbs

Fuel load 2nd hour: 28.63

Fuel load 3rd hour: 26.6 lbs

Fuel load 4th hour: 24.7 lbs

Total: 111 lbs, 1.875 cubic feet

Loaded weight without fuel: 97.3 lbs

Loaded weight: 210 lbs

Payload 4 hours: 210 lbs

Payload 2 hours: high lift mode: 135

Fuel load 3 hours high lift: 180

Loaded weight high lift: 302 lbs

3 hour payload high lift: 282 lbs

2 hour payload high lift: 334 lbs

2 hour payload ultrahigh lift:

1315 lb gw

611 hp 120 358 1st 173 second lbs

Payload: 624

GW:499.4 lbs

Maximum installed power 110 hp:

Power: 68 hp

Engine weight: 66 lbs (Stuttgart recuperated turboshaft 130 hp, 1.97 hp/lb, 0.54 lbs/hp-hr BSFC)

Fuel 1 hour: 27

Fuel 2 hour: 25.5 lbs

Fuel 3 hour 24 lbs

Fuel 4 hour: 22.9 lbs

Total turbine: 253 lbs

Total diesel: 257

Duct and propeller: 20 lbs

Magnesium fuel tank: 3 lbs

Structural bar: 5 lbs

Hydrogen two hours: 63 lbs 53.45 lbs total 116 lbs

Loaded weight: 165 lbs

Horizon PEM: 3 hp/lb 22 lbs

Motor 10 kw/kg: 11 lbs

Loaded weight electric: 250 lbs

5 min battery reserve @400 wh/kg: 26 lbs

Payload electric: 249 lbs

Methane 55.5 MJ/kg

Horizon PEM 3 hp/lb 45% EF without graphite bipolar plates 14 lbs FC

High speed AC motor: 10 kw/kg

14 lbs

Liquid hydrogen propulsion

Nanopore insulation 6.8 lbs

15.3 kg LH2 tank: 11.3% per hour , 29.5 psi maximum pressure:

LH2 mass: 33.7 lbs

Tank mass: 7 lbs

Motor 14 lbs

Total LH2 weight: 68.7 lbs

Total diesel weight: 154 lbs

Hourly operating cost: $8

Payload: 381 lbs

Loaded weight turboshaft: 318

GW: 4140 lbs

Disk loading: 120 lbf/ft2

Power loading: 3.85 lbf/hp

Max power: 1300 hp

Number of engines: 2

Continuous power: 1075

BSFC: 0.348 lbs/hp-hr

374 lbs 1st hour 340 2nd hour 309 lbs 3rd hour 271 lbs 4th hour

Total: 1294

Engine: 200 lbs

Ducted fan, fuel tank, structure, parachute, gearbox:

185 lbs

Loaded weight: 1679

Useful load: 2461 lbs

Net payload: 2281 lbs

The net payload depends on the operator weight, male weight varies greatly across different geographic regions, with the global average being 152 lbs, but an average is of little use, we can assume most operation will take place in Europe and North America, so we use a pilot weight of 180 lbs.

Average weight globally ; Asia, 126.9 ; Europe, 155.8 ; Latin America and the Caribbean, 149.4 ; North America, 177.5.

4 hour range turbine powered Hyper-Craft

2840 lbs 1570 hp 478 lbs 395 lbs 340 lbs 274 lbs

Fuel load: 1487 lbs

Payload: 1043 lbs

System weight: 180 lbs

Engine: 224 lbs

Loaded weight: 1401

Payload: 732 lbs

Payload minus pilot: 542 lbs

Cost: engine: $98,000

System: $45,000

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