1

Figure 1- Kestrel BWB flight demonstrator

ISABE-2015-20164

Airframe Integration for LH2 fuelled Distributed Propulsion Systems

Howard Smith & Panagiotis Laskaridis

Centre for Aeronautics

School of Aerospace, Transport & Manufacturing

Cranfield University

Cranfield, MK43 0AL

United Kingdom

Abstract

Research into the BWB concept has previously been

carried out by Boeing, NASA and Cranfield, amongst

others, identifying a number of claimed advantages

for this aircraft configuration. However, associated

with these advantages are a number of significant

challenges that continue to require consideration.

Some of these issues will be touched upon in this

paper as they are of particular relevance.

One of the claimed advantages of the BWB concept is

that it offers the possibility of better integration of the

propulsion system with the airframe. This results in

fundamental differences between the packaging of the

BWB and the conventional airliner configuration.

This benefit is explored in this paper with the

objective of integrating an LH2 fuelled Hybrid-

Electric propulsion system. Distributed propulsion

systems have been considered utilising hybrid electric

technologies, however, this study focusses on the

integration at a preliminary design level. The LH2

concept, with its low density induced packaging

issues, is contrasted with a Kerosene fuelled system.

Nomenclature

BLI Boundary Layer Ingestion

BWB Blended Wing Body

CESTOL Cruise Efficient, Short Take-Off and

Landing

CS Certification Specifications

EASA European Aviation Safety Agency

HTS High Temperature Superconductor

LH2 Liquid Hydrogen

LCH4 Liquid Methane

M Mach number

MTOM Maximum Take Off Mass

NASA National Aeronautics and Space

Administration

RPM Revolutions Per Minute

sfc specific fuel consumption

UHCA Ultra High Temperature Capacity Airliner

1 Introduction

There has been a steady increase in air traffic over

the last twenty years. This trend is likely to continue at

an anticipated rate of some 4-5% p.a. This has led to

growing concerns over the potential environmental

impact of aviation. In 2009 the International Air

Transport Association concluded, to ensure continued

viability of the industry, CO2 emissions should peak by

2020 but then reduce to half of the 2005 levels by 2050.

For this to be achieved it is likely that radically new

technologies will need to be utilized.

Research into the BWB concept has previously

been carried out by Boeing, NASA and Cranfield,

amongst others, identifying a number of claimed

advantages (Liebeck, 2002; Smith, 2000; Fielding and

Smith 2002) for this aircraft configuration including

more flexibility in packaging and the possibility of a

properly integrated propulsion system. Development at

Cranfield includes the Kestrel Flying demonstrator,

figure 1, in association with BAE-SYSTEMS and the

X-48 airframes production for Boeing/NASA, figure 2.

2

Figure 2 – X-48B BWB flight demonstrator

However, associated with the advantages are a

number of significant challenges (Liebeck, 2003;

Smith,2000; Fielding and Smith, 2002) that have

required due consideration. Issues associated with the

structural solution to the non-circular pressure cabin,

aerodynamic stability, control and departure modes,

propulsion integration and passenger emergency are

currently being explored. Some of these issues will be

touched upon in this paper as they are of particular

relevance.

One of the claimed advantages of the BWB concept

is that it offers the possibility of better integration of the

propulsion system with the airframe. This results in

fundamental differences between the packaging of the

BWB and the conventional airliner configuration. This

benefit is explored in this paper with the objective of

integrating a Hybrid-Electric propulsion system.

Distributed propulsion systems have been

considered utilizing conventional turbofans (Ko et

al,2003) and hybrid electric systems (Felder,2009)

however, this study focusses on the integration at a

preliminary design level.

(Smith, 2013) describes a preliminary design study

that explores the possibility of integrating a hybrid-

electric propulsion system into a BWB airframe. This

paper develops the concept further to utilise LH2 as the

primary fuel. These concepts are outlined below:

1.1 Kerosene/LH2 Hybrid Concept In the first concept (Smith, 2011;Smith, 2013) the

overall power-plant is comprised of four main elements,

as illustrated in figure 3. The first element being two

turbo-shaft engines that have the primary role of

generating shaft power. These engines utilize standard

kerosene fuel. The second element is the electric

generators which generate electrical power, from the

shaft power output, for both propulsion and secondary

power. The third element is an electrical power

distribution system and finally a series of distributed

electric fans.

Unfortunately, the utilization of conventional

technology in designing this propulsion system leads to

a build-up of poor efficiency factors that result in an sfc

that is higher than that associated with a conventional

system. However, this can be ameliorated through the

application of High Temperature Superconducting

technologies. Even within a 2030 timeframe it is

unlikely that cryogenic refrigeration systems will be

sufficiently light to permit their application here and so

liquid Hydrogen or Methane is used to cool these

components. After cooling the electrical systems the H2

is then burnt along with the jet fuel resulting in a

reduction in the kerosene fuel mass required.

Figure 3 – Distributed Propulsion System

Table 1, below, indicates the reduction in kerosene

required to achieve the design range due to the

additional LH2 fuel source. For comparison the

equivalent values are given for the utilisation of LCH4,

albeit at a higher temperature than the LH2.

Table 1

1.2 LH2 Concept The second concept proposes that the utilisation of

LH2 as a fuel implies a number of challenges in terms

of both the airborne system and the ground

infrastructure. Assuming that these can be resolved for

the Kerosene/LH2 hybrid concept then perhaps one

could dispense with the kerosene fuel and associated air

and ground systems entirely. Thus, this concept, briefly

described in (Smith, 2014), explores the possibility of

utilising solely LH2 as the propulsion system fuel.

Significant challenges result from the integration of

these systems and are discussed in the paper. Major

issues explored include packaging, certification and

safety of the distributed propulsion system. Many issues

relating to the BWB configuration are also discussed

particularly where they are compounded by the

propulsion issues.

2 Airframe Characteristics The case study aircraft (Smith, 2011) is an A-380

class ultra-high capacity airliner. Fin-tip to fin-tip wing-

Volume m3Mass kg Volume m3

Mass kg Volume m3Mass kg

Kerosene 260.8 204,750

Kerosene + LH2 237.4 186,320 94.9 6,710

Kerosene + LCH4 216.5 169,950 73.6 35,500

Kerosene LH2 LCH4

3

span is 80m and the body length is 48m.

Figure 4 – BWB Propulsion concepts

The aircraft is the latest development in a line of

designs, figure 4, dating back to 1998 with the BW-98,

a basic BWB configuration with externally mounted

engine nacelles. The BW-01 incorporated a properly

integrated propulsion system comprising of two engine

cores driving four boundary layer ingesting fans. Both

shaft and gas drives where investigated for power

transmission. The current incarnation is designated

BW-11 which is the subject of the present study.

The design mission is 555 passengers in a mixed

class arrangement with a range of 7650 nm at a cruise

Mach of 0.85. The MTOM is 468,000 kg. Figure 5

presents the general arrangement of the aircraft.

Figure 5 - -General Arrangement

3 Propulsion System The overall architecture of the Kerosene/H2

concept is similar to that proposed by NASA’s

CESTOL (Felder et al, 2009) concept. Primary power is

provided by 2 kerosene fuelled turbo-shaft engines

delivering shaft power to 2 electric generators.

Electrical power is then distributed, via power

converting circuitry, to electrical motors that deliver

shaft power to a number of thrust generating fans.

The system is comparatively complex and will,

necessarily, result in a chain of compounding efficiency

factors. To mitigate the reduction in efficiency many of

the system components exploit High Temperature

Superconducting technology. Cryogenic refrigeration of

the critical components results in significant efficiency

improvements.

The benefits of such an electrical propulsion are

many and varied. It offers the possibility of distributing

a number of small propulsive fan units. The electric

generators can perform the function of a gear-box very

efficiently allowing the decoupling of RPM from the

torque. This enables the turbo-shaft to be run at its

optimum speed whilst the propulsion fans can be

operated at their best performance speed. Failure of a

turbo-shaft engine will not result in a thrust asymmetry.

Failure of a single fan unit should result in a minimum

of system degradation. The distributed propulsion lends

itself to boundary layer ingestion thus increasing

aircraft airframe performance.

The pure H2 concept omits the kerosene tanks and

associated systems and replaces its chemical energy

with a greater quantity of LH2. The primary challenge

associated with this design is that the lower density of

the fuel implies that a significantly greater fuel volume

is required. This, of course, leads to packaging

constraints.

4 Fan Intake Integration The fan units, figure 6, are semi-submerged into the

rear of the upper surface of the aircraft to permit

boundary layer ingestion.

Figure 6 – Fan propulsor arrangement

The limited available depth across the outer wings

confines installation to the body region. Installation in

this location does, however, heavily constrain the

structural design in an area that is already fairly

complex.

5 Turbo-shaft Integration The installation of the engines, figure 7, was

constrained by a number of considerations. Prime

amongst these were the proximity to the generators to

minimize mechanical power transmission distance,

4

proximity to the cryogenic coolant to reduce H2

distribution issues and proximity to the electric fans to

minimize electrical power transmission distances.

Figure 7 – Turboshaft engine installation

Packaging related issues included matching the

physical size of the engines to the available internal

volume. Other considerations included the proximity to

an appropriate external surface to permit the intake and

exhaust integration, access for inspection, removal and

other maintenance requirements. Figure 8 depicts the

arrangement of the fuel tanks and engine layout. Span-

loading benefits were considered, as well as centre of

gravity constraints.

Figure 8 – Fuel Tank and Engine Layout

6 Generators and Transmission The generators utilize the shaft power output from

the turbo-shaft engines to generate electrical power. The

two electric generators, sized to produce 80MW each,

weigh 2325kg apiece.

The efficiencies of the first design iteration of the

system are Fan: 81%, Cables: 99.7, Inverter with

cooler: 98.8%, Generator with cooler 99.7%, Gear

box:89% giving a total of 70.5%. The second iteration

involed the adoption of a single stage fan, decreasing

the RPM of the generator and removing the gearbox.

The resulting system efficiencies are Fan: 90%, Cable:

99.7%, Inverter with cooler: 98.8%, Generator with

cooler 99.7%, Motor with cooler: 99.5% giving a total

of 87.9%.

7 Airframe Systems The packaging of the propulsion and fuel systems

is, of course, constrained by the packaging of the other

airframe systems in addition to the cabin and baggage

volumes. Figure 9 indicates the location of the

Environmental Control System in addition to the

passenger cabins.

Figure 9 – Airframe systems

8 Airframe Structures The airframe structure is designed to comply with

EASA CS-25. Within the scope of the study both

metallic and composite structures, figure 10, were

explored from mass, manufacture, maintenance and cost

perspectives.

Figure 10 – Structural concepts

9 Cryogenic Systems - Kerosene/LH2 Hybrid

Concept

The cryogenic systems are vital to the overall

efficiency of the propulsion system. They do, however,

present many interesting challenges. The storage

location of the LH2 is particularly constrained, figure

11. Whilst it would be preferable to clearly separate the

LH2 tanks from both the turbo-machinery and the

passenger cabin. Unfortunately, the pressurization of

the tanks means that they need to have a fairly high

diameter to be weight efficient (as compared to a larger

number of smaller diameter tanks). As a result they, too,

need to be located within deep areas of the body.

Figure 11 – LH2 Tank Arrangement

14 Electric Fan Engines

on the Top of the Fuselage

2 Turbo-Shaft

2 Generators

Wing Tanks

LH2 Tanks

Trim Tank

5

Whilst the locations chosen are close to the

passengers they are remote from the engines and fans.

The design of the fuel system considers fuelling,

venting, fuel transfer, system failures, maintenance and

safety – figures 12 depicts some of the components

The system comprises of the tanks (inner and outer

vessels and attachments), the refuel/defuel/jettison

system (gravity adapter, pressure adapter, shut off valve

and vacuum jacket pipe), the cooling feed system

(booster pump, non-return valve, shut off valve and

check valve), the transfer system (transfer valve and

shut off valve), the pressurisation and vent system

(pressure relief valve, vacuum jacketed pipe and

pressurisation unit), the vacuum insulation system

(vacuum pump, vacuum level sensor, multi-layer

insulation and vacuum jacketed pipe), the fuel

management (refuel panel, fuel management computer,

quantity probes, level sensors, fuel properties

measurement unit and pressure switch) and

miscellaneous attachments.

Figure 12 – Fuel tank concept

10 Cryogenic Systems - LH2 Concept The majority of the cryogenic systems of the LH2

concept are identical to those of the Kerosene/LH2

concept. The primary difference being that the volume

of LH2 is considerably greater. By comparison to the

pure kerosene concept, which carries 260m3 of fuel, the

pure LH2 concept would need to carry 1060m3 of fuel

to replace the chemical energy of the kerosene, table 2.

Note that the resulting range would depend upon the

mass of the fuel (lighter) and the mass of the resulting

fuel system (heavier). The hybrid kerosene/LH2

concept carries 95m3 of LH2.

Table 2

If the additional fuel volume were to be located in

cylindrical tanks above the upper passenger deck and

throughout the lower deck (replacing the passengers), as

depicted in figure 13, the available volume would be

405m3.

Figure 13 – Cylindrical tank arrangement

Clearly, the tank design needs to be further

developed if a greater fuel volume is to be achieved. To

this end, two tank concepts have been explored; the

Obround tank and a complex conformal tank.

10.1 The Obround tank The Obround tank geometry can be described as

rectangular prismatic sections with semi-cylindrical end

surfaces as depicted in figure 14. This results in a tank

that is more volume efficient than cylindrical tanks but

still retains the pressure efficient end bulkheads. They

do, however, have large flat surfaces that require the

structure to carry the pressure loads in bending.

Figure 14 – Obround tank arrangement

10.2 The Complex Conformal tank To utilise the maximum available internal volume

the tanks would need to be more conformal in shape.

These tanks, referred to as Complex Conformal tanks,

are more irregular in shape, as shown in figure 15 and

carry more of the pressure loads in bending. This will

result in an increase in the tank structure mass.

Figure 15 – Complex conformal tank arrangement

In principle, one final concept will need to be

explored, though not considered in this paper, which is

the integral tank concept.

Fuel Total fuel volume m2

Kerosene 260.8

Kerosene + LH2 332.3

Kerosene + LCH4 290.1

LH2 1060

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10.3 Fuel Tank Packaging

In all concepts the fuel tanks need to be integrated

with the airframe systems, structure and passenger

cabins. Consideration needs to be given to mounting

structure and maintenance. Figure 16 depicts the

integration of the Complex Conformal tank concept.

Figure 16 – Airframe/tank integration

Whilst a true integral tank concept has not yet been

investigated, there are considerable advantages to be

gained, particularly for the Obround and Complex

Conformal tanks, by properly integrating the tank

structure with the airframe structure. This is the topic

of an ongoing study.

To increase the capacity of the LH2 tanks it is

necessary to displace passenger cabin volume. For a

given vehicle size, the greater the fuel volume the fewer

passengers can be accommodated, as shown in Figure

17. The difference in tank volume between the Obround

and Complex tanks can also be seen.

Figure 17 – Passenger / tank capacity

11 Performance There is a trade-off between the volumetric

efficiency of the fuel tanks (and hence range) and their

complexity (and hence mass, cost and maintenance

burden) as can be seen in figure 18

Figure 18 – Range achievable with tank concept

12 Airworthiness The BW-11 is intended to be EASA CS-25

compliant. Historically, the development of

airworthiness requirements has progressed step by step

with the development of the conventional airliner

concept. Consequentially, demonstrating that an

2,610 2,722

3,489

-

500

1,000

1,500

2,000

2,500

3,000

3,500

4,000

IntersectingCylinders

Obround tanks Complex tanks

Ran

ge (

nm

)

7

advanced aircraft concept meets existing airworthiness

requirements or can, at least, demonstrate comparable

levels of safety is challenging. The BWB configuration

has a number of characteristics that require special

consideration. The integration of the distributed

propulsion system further constrains the design.

Figure 19 – upper passenger deck

Emergency evacuation from a double deck BWB is

less straight forward than from a conventional

configuration due to the lack of proximity between the

passengers and external walls. The incorporation of the

broad fan intakes makes evacuation to the rear of the

aircraft difficult. Emergency exits are primarily located

to the sides and front of the cabins. The location of the

turbo-shaft engines further constrains viable lateral

evacuation routes. The proposed solution is illustrated

in figures 19, 20 & 21. The need for appropriate

evacuation routes further constrains the integration of

the LH2 tanks. Gaps in the tanks must be incorporated

to facilitate access to the emergency evacuation doors

as shown in figure 20.

Figure 20 – Emergency exit / tank integration

Figure 21 – lower passenger deck

The integration of the cryogenic fuel tanks is an

example of a design feature that cannot be directly

mapped to the current airworthiness requirements. For

the purpose of this study (NASA, 1996; Sass et al,

2010; Arnold et al, 2007; Mital,2006) were used to

establish a position that would, in practice, need to be

negotiated with the airworthiness authorities.

13 Engine fragmentation zones A considerable volume fraction of the aircraft is

devoted to the accommodation of the fuel. This gives

rise to concerns about the proximity of the turboshaft

engines to the fuel tanks. Consideration has been given

to engine fragmentation zones and their relationship to

the fuel and passenger cabins. Figure 22 shows that the

15° fragmentation zone from the port engine only

intersects the outboard starboard tanks. This is due to

the careful placement of the tanks and engines.

Figure 22 – Fragmentation zones

Furthermore, figure 23 shows that a burst fragment

will only hit the tanks if it is projected within a 4°

segment of the entire disk i.e. a 356/360 chance of

missing the tank. Adequate venting of the void between

tank and external structure will permit any lost H2 to

dissipate. (Khandelwal, 2013) indicates that H2 is likely

to disperse more readily than a liquid fuel.

Figure 23 – Critical disk burst angle

14 Operational Aspects

Current operational issues that are being assessed

are turnaround time and maintenance. The inspection of

the tanks and surrounding structure is an important

issue. The complexity of this requirement was first

explored on a conventional aircraft configuration.

Figure 24 depicts an aircraft with four different tank

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concepts; box tanks (in forward fuselage), cylindrical

tanks (in aft fuselage), external tanks (stowed under the

wings) and integral tanks (within the wings). The box

tank may be packaged in a similar manner to standard

baggage containers. These can, in principle, be accessed

and removed via the baggage doors. This would also

give access to the surrounding fuselage structure. The

external tanks could, if required, be removed and

dismantled to enable thorough inspection access to be

achieved.

Figure 24 – Conventional aircraft with low density fuel tank

concepts

One could also conceive of a system whereby the

empty tanks could be unloaded and replaced with pre-

fuelled tanks.

The wing integral tanks are not a good solution on

this configuration due to the lack of depth in the wing.

After allowing for the thermal insulation the available

volume is limited. Furthermore, the insulation could

prevent easy inspection of the structure. A possible

solution would be a removable thermal lining-similar to

the bag tanks in C-130 military transport wings.

The cylindrical tank depicted in this concept

provides a structurally efficient way of containing a

pressurized fuel. It also makes good use of the available

volume - in this case in the aft fuselage, however, it can

be seen that access is now becoming challenging. To

enable the tanks and surrounding structure to be

inspected three solutions are presented. The first, figure

25, involves designing the tank to be a straight cylinder

of sufficiently small diameter that it may be extracted

through the tail cone of the aircraft. This does imply

that less volume can be utilised.

Figure 25 – Cylindrical tank removal concept

The second possibility, figure 26, utilises a more

conformal tank geometry that is removed through the

tail cone but requires dismantling within the fuselage.

Figure 26 – Tank dismantling concept

The third proposal is to segment the tanks into

sufficiently small units, figure 27, that each can be

removed individually through a large maintenance door.

This carries the penalty of complexity.

Figure 27 – Segmented tank concept

Whilst these designs appear complex they

demonstrate that possible solutions do exist.

Application to the BWB configuration is more

challenging due to the high volume fraction of the fuel

combined with the complexity of the geometry. Figure

28 depicts the geometry of an obround tank within the

upper fuselage zone of a BWB. Whilst any of the

previous solutions could be applied here, an alternative

solution might be to make use of more of the BWB

volume to allow direct access by leaving a greater

volume of clearance between tank and fuselage

structure.

Figure 28 – Obround tank access

15 Conclusions There are many challenges associated with the

BWB configuration as applied to the role of a civil

airliner. One advantage offered is the possibility of

integrating an advanced propulsion system such as

hybrid electric distributed propulsion. This preliminary

design of a case study aircraft highlights a number of

specific issues and demonstrates possible solutions. The

relationship between passenger capacity and fuel

capacity has been presented. Key challenges have been

identified.

9

16 Acknowledgements The author wishes to acknowledge the contribution

of his colleagues, graduate students and M L

Shamsuddin, A Kadimi & K B Kokane at Cranfield

University’s Centre for Aeronautics.

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