*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 It is argued that electric propulsion is somehow superior, and that combustion is obsolete.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.

Figure above illustrates the rear fuselage conformal liquid hydrogen tank

Figure below illustrates the download penalty