airframe integration for lh2 fuelled distributed
TRANSCRIPT
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
6
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
8
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.
17 References
Arnold, S. M., Bednarcyk, B. A., Collier, C. S. and Yarrington, P. W. (2007), "Spherical cryogenic hydrogen tank
preliminary design trade studies", Vol. 7, pp. 7125.
Felder J.L., Kim H. D., Brown G. V. (2009) “Turboelectric Distributed Propulsion Engine Cycle Analysis for
Hybrid-Wing-Body Aircraft”, 47th AIAA Aerospace Sciences Meeting Including The New Horizons Forum and
Aerospace Exposition, 5 - 8 January 2009, Orlando, Florida
Fielding, J.P., Smith, H. (2002), Development of Environmentally Friendly Technologies and Configurations for
Subsonic Jets, ICAS 2002. International Congress of the Aeronautical Sciences, 2002
Khandelwal, B., Karakut, A., Sekaran, P., Sethi, V., Singh, R., “Hydrogen Powered Aircraft: The future of Air
Transport” Progress in Aerospace Sciences 60 (2013) 45-49
Ko, A., Leifur , Leifsson T., Mason W.H., Schetz J.A. and Grossman B. (2003), “MDO of a Blended-Wing-Body
Transport Aircraft with Distributed Propulsion”, AIAA's 3rd Annual Aviation Technology, Integration, and
Operations (ATIO) Tech, 17 - 19 November 2003, Denver, Colorado
Liebeck R.H. (2002) “Design of the Blended-Wing-Body Subsonic Transport”, AIAA-2002-0002, 40th AIAA
Aerospace Sciences Meeting & Exhibit,14-17 January 2002 / Reno, NV
Liebeck R.H. (2003), “Blended Wing Body Design Challenges”, AIAA/ICAS International Air and Space
Symposium and Exposition: The Next 100 Years 14-17 July 2003, Dayton, Ohio
Mital, S. K. e. a. (2006), Review of current state of the art and key design issues with potential solutions for liquid
hydrogen cryogenic storage tank structures for aircraft applications, NASA/TM—2006-214346, Toledo, Ohio.
National Aeronautics and Space Administration (1996), Safety standard for hydrogen and hydrogen systems , NSS
1740.16, Office of Safety and Mission Assurance, Washington DC
Sass, J. P., Cyr, W. W. S., Barrett, T. M., Baumgartner, R. G., Lott, J. W. and Fesmire, J. E. (2010), "Glass bubbles
insulation for liquid hydrogen storage tanks", Vol. 1218, pp. 772.
Smith, H. (2000) College of Aeronautics, Blended Wing Body Development Programme. ICAS 2000.114.
International Congress of the Aeronautical Sciences. Harrogate, UK, 2000
Smith, H. (2011), “Advanced Blended Wing Body High Capacity Airliner BW-11 Project Specification”, DES1100,
Department of Aerospace Engineering, Aircraft Design Group, Cranfield University, U.K.
Smith, H (2013), “Airframe Intigration for Distributed Propulsion Systems”, ISABE-2013-1718, 21st ISABE
Conference, Sept 9-13, Busan, Korea
Smith, H. (2014), “Airframe integration for an LH2 hybrid-electric propulsion system”, Aircraft Engineering and
Aerospace Technology: An International Journal, Vol. 86 Iss: 6, pp.562-567