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The LAPCAT-MR2 hypersonic cruiser concept


Abstract and Figures

This paper describes the MR2, a Mach 8 cruise passenger vehicle, conceptually designed for antipodal flight from Brussels to Sydney in less than 4 hours. This is one of the different concepts studied within the LAPCAT II project. It is an evolution of a previous vehicle, the MR1 based upon a dorsal mounted engine, as a result of multiple optimization iterations leading to the MR2.4 concepts. The main driver was the optimal integration of a high performance propulsion unit within an aerodynamically efficient wave rider design, whilst guaranteeing sufficient volume for tankage, payload and other subsystems.
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This paper describes the MR2, a Mach 8
cruise passenger vehicle, conceptually designed for
antipodal flight from Brussels to Sydney in less than
4 hours. This is one of the different concepts studied
within the LAPCAT II project [1]. It is an evolution
of a previous vehicle, the MR1 based upon a dorsal
mounted engine, as a result of multiple optimization
iterations [2] leading to the MR2.4 concepts. The
main driver was the optimal integration of a high
performance propulsion unit within an
aerodynamically efficient wave rider design, whilst
guaranteeing sufficient volume for tankage, payload
and other subsystems.
The aerodynamics for the MR2 is a
waverider form based upon an adapted osculating
cone method enabling to construct the vehicle from
the leading edge while reducing integration
problems between the aerodynamics and the intake.
The intake was constructed using
streamtracing methods from an axisymmetric inward
turning compression surface and was integrated on
top of the waverider in a dorsal layout. The shape of
the streamtraced intake surface changed during
various evolutions of the vehicle but the final shape
was elliptical with a ratio of semi-major to –minor
axes of 3. This 3D shaped intake feeds a dual mode
ramjet/scramjet combustion chamber and is foreseen
to operate between M4.5 and M8. Below M4.5 an
accelerator engine is required. Behind retracting
door panels, a 2D-intake with moveable ramps is
installed based upon the XB-70 intake and was
shown numerically to provide the necessary mass
flow and pressure recovery for the ATR engine
downstream [3].
The nozzle was constructed in two sections;
the first isentropic 2D nozzle has an area ratio of 3,
thus bringing the elliptical combustor cross section
to a circular cross-section. During Ramjet-mode, this
nozzle was used as a combustor that thermally
choked, allowing for supersonic expansion in the
second nozzle. The second nozzle itself was
streamtraced from an axisymmetric isentropic
expansion and truncated to a suitable length. Both
nozzles were designed for cruise conditions.
The final vehicle is shown below in Figure 1
while specific details of the design are expanded
upon in the next section.
Figure 1 MR2 Vehicle
Discussion of sub-systems
The deployment of a high-speed cruiser only makes
sense for long haul flights with ranges up to
antipodal destinations. Simply based on the Breguet
range equation, at least a high L/D is necessary in
combination with a low specific fuel consumption.
Therefore, a waverider design was laid out to enable
a L/D > 6 for a cruise Mach number around 8. To
maximize the available planform for lift generation
and to optimize the internal volume, the engine was
positioned on top. As the top surface of the wings
J. Steelant and T. Langener*
* ESA-ESTEC, Keplerlaan 1, 2201 AZ Noordwijk, Netherlands
Vehicle Design, Hypersonic Flight, Combined Cycle Engine, Dual-Mode Ramjet
Steelant J. and Langener T.
and the fuselage are nearly aligned with the flight
vector, the lift is mainly generated by the windward
side of the vehicle. This layout allowed furthermore
to expand the jet to a large exit nozzle area (lower
specific fuel consumption) without the need to
perturb the external shape which would lead to extra
pressure drag.
Though a conical shape for an air intake has
intrinsically a minimal wetted area for a given
compression ratio, one has still the choice between
an inward versus outward turning intake. To
evaluate the impact on the overall performance of
the engine and its integration into the fuselage, a
general analysis was carried out to address the pros
and cons of both intake types. Values related to
outward turning air intakes will be denoted with the
subscript out whereas inward turning intakes by inw.
Fig. 1: Outward (top) and Inward (bottom) turning
conical air intakes.
Based on the above figures, we start off with an
identical air capture A
, conical intake length L
combustor + cowl length L
and a compression ratio
CR resulting in a combustor cross section A
defined as:
The radius of the combustor for the inward turning
intake can be defined as function of A
and CR:
The corresponding wetted combustor + cowl area
is then given as:
For the outward turning intake, the height of the
combustor can be defined as follows:
The corresponding wetted combustor + cowl area is
then given as (the factor 2 applies for the double
amount of inner walls compared to an internally
turning intake):
The ratio of the wetted combustor areas can hence
be written as:
For a contraction ratio e.g. CR = 9, the wetted
combustor area is 6 times higher for an externally
versus an internally turning intake. However, for an
internal compression, there is also an extra wetted
area stemming from the casing which is not present
for an external compression. Assuming a worst case
scenario where we have a completely axi-symmetric
casing (from an operational point of view this is not
desirable due to not-startability addressed later on),
the wetted intake casing areas are respectively:
The wetted area on the conical intake, whether
internal or external, is given as:
For a specific flight dynamic pressure q
and a given
skin friction Cf, one can calculate the viscous casing
For the drag estimation on the intake cone, cowls
and inside the combustion chambers, one can apply
the general approach that the velocity is nearly
constant throughout the internal flowpath at these
high flight speeds. This means that the dynamic
pressure at the cowl and in the combustion chamber
can be linked to the flight dynamic pressure as:
Hence for the above given value of CR=9, the
dynamic pressure in the combustor is 9 times larger
than the flight dynamic pressure. The drag generated
within the ducts consisting of the cowls and
combustors is given respectively as:
Assuming an identical skin friction coefficient, the
ratio of both drag components is hence identical to
the ratio of wetted areas:
The drag on the cone surface itself is approximated
The total drag for the external compression intake
can then be expressed as the contribution form (1),
(2) and (3):
Hence the difference in drag between both intakes is
at most:
One has to consider that the casing for the internally
turning intake doesn’t necessarily needs to be taken
as a drag force on the intake. Depending on the
layout, if this casing can be integrated cleverly into
the fuselage it can actually serve as part of the lifting
geometry of the vehicle. In case of the LAPCAT
MR2, near half of this casing is located on the
windward side generating lift, which otherwise
needed to be provided by extra wing surface.
Moreover, in order to assure a startable intake, the
particular choice of the positioning of the elliptical
streamtracing shape slightly below the axis of the
conical template flowfield resulted in a nearly
triangular opening at the top. This opening doesn’t
contribute nor to the external drag or the intake
cone. Hence, the last term L
in the above equations
(1) and (2) can be dropped at best along with a
reduction of s representing the effective exposed
surface for a non-closed internal turning surface. In
any case, for the considered vehicles, L
is about 3
to 4 times larger than L
. With a CR = 9, the
multiplication factor in the above equation ranges
between 11 to 12. As a ratio we have:
After some simplifications (s L
), one can reduce
the ratio to:
For the same values for L
and CR as used
previously, the above ratio results into a factor
ranging from 2 up to 3 (the latter number due to a
reduced conical surface for a stream-traced intake).
The evolution of the drag ratio is shown in
Figure 2.
This total drag increase with 200% to 300% is
mainly linked to the 6 times larger drag within the
annular combustor compared to the circular duct.
Steelant J. and Langener T.
Figure 2: Drag ratio for outward vs inward turning
intakes; colours representing fully and not-
optimized integration into a hypersonic cruiser.
The front view of the MR2 in Figure 3
shows how the intake is integrated within the
waverider shaped fuselage. The elliptical capture
shape (12m x 4m) used for the streamtracing
procedure was projected to the most forward plane
which served as the leading edge of the waverider.
Figure 3: Front view of MR2 showing the
integrated intake and waverider leading edge.
Figure 4 shows a zoom of the integrated region. The
waveriders elliptical leading edge does not extend
from 0° up to -180° but instead goes from about -28°
to -152°. The portion missing between -152° and -
180° is circled in the figure.
Figure 4: Zoom of intake/waverider integrated
Although this introduced a small drag
surface downstream of the leading edge when the
oblique angle between the two was blended out, it
was necessary to limit the wingspan of the
waverider. Due to the boundary layer displacement
correction that has been applied to the intake
geometry, the cross-section of the combustor is not
exactly elliptical.
The Dual Mode Ramjet unit was modelled
from 0D/1D engineering tools up to 3D CFD codes
with detailed combustion chemistry at different
levels of details [4] [5]. This allowed a general
layout of the combustor and a detailed injector strut
layout ensuring a good mixing and combustion
This detailed CFD analysis further allowed to assess
the spillage drag and their effect on the overall
installed thrust.
Figure 5: Layout of Injector strategy.
The Air-Turbo-Rocket (ATR), [6] [7], [8] inlet in
the MR2 design were integrated as shown in Figure
6. A detailed CFD study was conducted to optimize
the different ramp settings. A mass capture of over
50% could be achieved between Mach 1.2 and 4.5
matching the required mass capture for the ATR
engine based upon an expander cycle. Also the
needed pressure recovery could be achieved.
Figure 6: LAPCAT-MR2 ATR air intake
diversion door opening side view
The nozzle contour was designed using the
method of characteristics (MOC) and afterwards
streamtraced using a similar method as for the
intake. The complete nozzle consisted of an initial
2D isentropic expansion followed by a 3D isentropic
expansion. Whilst this is not the most efficient way
of expanding the flow, it provided a discontinuity in
the surface to fix the thermally induced normal
shock during Ram-mode.
The 2D expansion was designed for an area
ratio of three so that the combustor ellipse was
brought to a circle. This then minimized the length
of the following 3D nozzle. Despite this, the 3D
nozzle was 75m long and so was truncated to 43m to
fit with the vehicle length. This would seem to
produce quite a drop in thrust but in reality the final
30% of an isentropic expansion produces relatively
little thrust when the low pressures and skin friction
are considered.
Figure 7:
Zoom of MR2 nozzle. The blue region
denotes a thrust surface.
Figure 7
shows a zoom of the nozzle with
the truncation line. In order to properly integrate the
nozzle into the vehicle a rough taper was inserted
around the nozzle which in effect increased its area
ratio. CFD computations showed that the overall
thrust of the nozzle was approximately the same as
calculated for the un-truncated nozzle using the
isentropic expansion and the 0.85 nozzle efficiency.
The waverider shape was designed with
structural integrity in mind but a rigorous structural
mass analysis is needed to properly assess the
vehicle performance. The waverider shape does not
conform to typical vehicle topologies and so
methods such as WAATS method produced very
different results depending on the interpretation of
different parameters. A deviation of up to 40-50%
was found between different estimates of the
structural weight.
However, specific studies have focused on
the structural masses of waveriders indicating that
the lower estimate of structural mass found using
WAATS has some validity [9]. A lobed design
technique where the fuel tanks are incorporated into
the vehicle load structure (
Figure 8
) were proposed.
Figure 8:
Multi-lobe internal structure and
corresponding FEM model.
A finite element analysis was conducted of the
multi-lobed structure which confirmed its feasibility
with other studies in literature and the lower mass
estimates of the WAATS analysis. No uncertainty
factor for this rather new and innovative structural
design method is included whereas a 71% increase
Nozzle truncation line
Steelant J. and Langener T.
of the ideal body and wing weights to account for
cut-outs, gage penalties, fasteners, and machining
constraints is included.
In LAPCAT-I [10], it was clearly shown
that liquid hydrogen is the only fuel able to achieve
antipodal flight due to its high specific energy
content. Hydrocarbons lead too fast to a large
GTOW prohibiting a fuel-efficient acceleration to
cruise speed. Liquid hydrogen was selected as the
fuel for the vehicle due to its very high specific
energy content. Despite its low density and the
inherently low volume on board of a waverider
concept, the available internal tankage volume
allowed more than 180 tons of fuel mass taking into
account thermal insulation and ullage volumes. The
different tank compartments which were modelled in
a detailed CAD design made it possible to further
calculate the shift of the CoG of the fuel at different
fill levels.
The mass estimations for the other subsystems were
based upon existing correlations derived from
supersonic and hypersonic vehicles (based upon the
WAATS weight analysis tool) or directly obtained
by provided weights in the open literature (e.g.
landing gear). The total Gross Take-Off Weight
ended up to 400tons of which 60tons was allocated
to the payload and about180tons for the fuel. The
mass breakdown for the other subsystems are
detailed in
Figure 9
. The acronyms used in the
legend are representing the different systems on-
bard and are listed here below:
aerodynamic surfaces: wings and control
surfaces: WSURF
body structure: WBODY
Thermal Protection System (TPS): WTPS
propulsion: engines and tanks: WPROPU
take-off and landing gear: WGEAR
power supply: WPOWER
payload: WPAYLOAD (i.e. passengers, crew,
cabin, luggage,…)
fuel: WFUEL
margin: WMARGIN
The related weight values are embedded into the
Figure 9:
GTOW Mass Breakdown of LAPCAT
The trajectory calculations were performed
using ASTOS [11], a three/six degrees of freedom
trajectory code in a cartesian coordinate system with
its origin at the centre of a spherically rotating
The aerodynamic and propulsive forces are
based upon engineering methods or detailed nose-to-
tail computations. This entails also complete
thermodynamic cycles for the ATR and DMR.
The range trajectory included a 400 km
subsonic cruise to reach the ocean from take-off at
most European airports prior to acceleration to
supersonic. This acceleration phase of the trajectory
could take up to 45% of the fuel mass. It was shown
in [6] that the mission Brussels – Sydney as a
representation of an antipodal flight is in principally
feasible, given the available vehicle layout and the
databases for the engines and for the aerodynamic
performance. The flight time to Sydney would be
around 2h55m whereas all the available fuel on
board would be consumed. The missions to Tokyo
need 2h13m and the flight to Los Angeles 2h20m.
All three simulated routes lead over the North Pole
and cross the Bering Strait in order to avoid
supersonic cruise over inhabited land.
Fig. 2 View on complete trajectory with pole
crossing and Bering Strait passage.
Once the full trajectory was known along
with the vehicle weight change and the CoG-shift, a
trimming analysis could be carried out. However, a
particular difficulty is related to the influence of the
engine flowpath on the overall pitching moment.
Additionally to the external aerodynamics, one
needs to take into account the effect of intake
spillage along the full Mach range together with the
different moments induced both by engine on and
off operation. Thanks to a detailed Nose-to-Tail
computation, both for engine on and off conditions
at different Mach numbers allowed to evaluate the
trimmability of the vehicle. It indicated that both the
presence of canards and ailerons are needed.
In the presented paper the basic principle
and design choices of the LAPCAT MR2 hypersonic
cruiser concept have been described with particular
attention to the choice of the intake and its
integration into the aircraft layout as a dorsal
mounted engine. By this way a combined high
aerodynamic and propulsion efficiency could be
guaranteed which is a prerequisite to assure a long
range without excessive fuel burns and gross take-
off weights.
Also the other major subsystems have been
concisely addressed allowing to perform a first
feasibility study and a global performance. The
analysis indicated that a hypersonic cruiser at Mach
8 for antipodal flight is conceptually feasible
provided liquid hydrogen is used as a fuel. With a
GTOW of 400tons and a fuel burn of 180tons, the
antipodal range from Brussels to Sydney is
achievable within 3 hours.
Presently, further elaborations are ongoing
with respect to the optimization of an integrated
advanced thermal protection system including on-
board power generation while exploiting the large
heat capacity of the on-board cryogenic fuel.
This work was performed within the ‘Long-Term
Advanced Propulsion Concepts and Technologies II’
project investigating high-speed transport. LAPCAT
II, coordinated by ESA-ESTEC, is supported by the
EU within the 7th Framework Programme Theme7
Transport, Contract no.: ACP7-GA-2008-211485.
Further info on LAPCAT II can be found on
[1] J. Steelant, "Sustained Hypesonic Flight in
Europe: Technology Drivers for
International Space Planes and Hypersonic
System Technologies Conference, Bremen,
Germany, 2009.
[2] N. Murray, J. Steelant and A. Mack,
"Conceptual Design of a Mach 8
Hypersonic Cruiser with Dorsal Engine," in
Sixth European Symposium on
Aerothermodynamics for Space Vehicles,
Versailles, France, 2008.
[3] C. Meerts and J. Steelant, "Air Intake
Design for the Acceleration Propulsion
Unit of the LAPCAT-MR2 Hypersonic
Aircraft," in 5th European Conference for
eronautics and Space Sciences (EUCASS)
, Munich, 2013.
[4] T. Langener, J. Steelant, P. Roncioni, P.
Natale and M. Marini, "Preliminary
Performance Analysis of the LAPCAT-
MR2 by means of Nose-to-Tail
Computations," in 18th AIAA/3AF
International Space Planes and Hypersonic
Systems and Technologies Conference,
Tours, France, 2012.
Steelant J. and Langener T.
[5] P. Roncioni, P. Natale, M. Marini, T.
Langener and J. Steelant, "Numerical
Simulatoins of the LAPCAT MR-2 Vehicle
Scramjet Engine," in ISABE-2013-1656,
Busan, S. Corea, 2013.
[6] T. Langener, S. Erb and J. Steelant,
"Trajectory Simulation and Optimization of
the LAPCAT MR2 Hypersonic Cruiser
Concept," in ICAS 2014, St. Petersburg,
Russia, 2014.
[7] V. Fernández-Villace, G. Paniagua and J.
Steelant, "Installed performance evaluation
of an air turbo-rocket expander engine,"
erospace Scienceand Technology, vol. 35,
pp. pp. 63-79, 2014.
[8] I. Rodríguez Miranda, V. Fernández
Villacé and G. Paniagua, "Modeling,
Analysis, and Optimization of the,"
of Propulsion and Power, p. DOI:
10.2514/1.B34781, 2013.
[9] A. Bardenhagen, H. Kossira and W.
Heinze, "Weight Estimation of Hypersonic
Waveriders within the Integrated Design
Program Prado-Hy," in AIAA 96-4546,,
[10] J. Steelant, "Achievements obtained for
sustained hypersonic flight within the
LAPCAT project," in 15th AIAA
International Space Planes and Hypersonic
Systems and Technologies Conference,
Dayton, OH, 2008.
[11] ASTOS Solutions GmbH, ASTOS 7 -
ASTOS Model Library, Unterkirnach,
Germany, 2011.
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2014 proceedings or as individual off-prints from the
... In the EU FP7 project LAPCAT [80] and the EU H2020 project STRATOFLY [81], numerical and experimental studies of a conceptual design for antipodal hypersonic passenger travel have been performed. In LAPCAT, a flight vehicle model was developed based on a waverider design, LAPCAT-II MR2, having a dorsal-mounted, dual-mode ramjet engine [82]. In STRATOFLY this model was further refined, resulting in the STRATOFLY MR3 model. ...
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In this study the aerodynamic shape optimization of subsonic submerged intake is performed. The total pressure at the aerodynamic interface plane and lift are employed as objective functions. Drag is constrained and employed as penalty objective. Open-ware platforms are used including SALOME for solid modeling, GMSH for the hybrid mesh generation and SU2 for flow solutions and adjoint based shape optimization. Free form deformation box is employed for shape parameterization and surface deformation. Initially NACA intake placed on a flat plate is solved and result is compared with reference study at subsonic flow conditions. Thereafter single objective shape optimization is employed which enhances the CL/CD from 0.08 to 0.53 at the end of design cycles. The second case is of trapezoidal-entrance intake placed in slender aerodynamic body. The results are compared with reference study and baseline case for optimization is established. The shape is optimized using multi objective function and for multiple free form deformation boxes. The intake remains flush with the body during the shape deformation process. The intake internal surface is then allowed to deform along the optimization steps. The pressure recovery increases by 3%. The third case is similar but with circular entrance in slender body. A single free form deformation box encloses the intake region and the deformation is allowed such that optimization creates a semi-submerged shape. The total pressure recovery increases by 2% and CL/CD also increases by 10%. In the last case again the trapezoidal entrance intake is employed. The optimization process is similar to the third case. The optimum shape produces a substantial deformation at the intake surface and a semi-submerged intake is finally formed. The total pressure recovery improves by 6% and an additional CL/CD ratio of 0.04 is achieved.
Increasing complexity associated to advanced aircraft concepts for high-speed flight is pushing the limit of engineering practices concerning design phases of products lifecycle. Specific methodologies aimed at managing requirements over the entire vehicle architecture definition shall then be properly formalized and supported by means of standardized processes and languages in order to be effective. The approaches suggested by the Systems Engineering practices, especially when considering the model-based environments as main tools for systems design, can be exploited to face these challenges, if properly tailored to suit the specific engineering field considered. This paper aims at proposing a model-based process for the design of high-speed aircraft exploiting a formalized methodology and the typical tools of Systems Engineering. A conceptual design exercise, based on the STRATOFLY MR3 hypersonic cruiser case study, is performed from high-level stakeholder objectives to vehicle performance estimations, passing through functional and interface analyses. Additionally, some insights concerning on-board subsystems sizing activities are provided in terms of integration within the high-level design process. The whole work highlights the capability of performing a seamless requirements management and development within the design procedure, particularly focussing on the relationships among the different stages of the workflow.
The STRATOFLY MR3 vehicle is the main objective of the STRATOFLY project, which aims to develop a hypersonic air breathing concept capable of covering antipodal routes in less than three hours. The aircraft architecture features a waverider configuration, internally supported by multi-bubble integral cryogenic tanks hosting LH2 propellant, being one of the major challenges the integration of lightweight structures with the high-speed propulsion system. The objective of this research is to completely define an efficient structural scheme of the multi-bubble structures. To do so, a multidisciplinary analysis of the full-scale aircraft model is carried out to assess the viability of the vehicle prototype. Once the flaws of the initial structural layout are identified, a set of stiffener elements were developed to generate a scheme which can withstand the loads that hypersonic flight entails. In the multi-bubble structures, a topology optimization strategy was applied to obtain a set of tension rods connecting the top and bottom parts of the bubbles to support the pressure loads. The proposed configuration was sized and analyzed for multiple points of the aircraft mission, obtaining stress levels below the failure criteria adopted for each material. In addition, the results show low displacements that guarantee an adequate aerodynamic behavior and engine performance, while mantaining global natural frequencies in the range of commercial airplanes.
Experimental determination of NO and \(H_2O\) production during hydrogen combustion applying laser absorption spectroscopy will be performed in the High Enthalpy Shock Tunnel Göttingen of the German Aerospace Center. The combustion process forming these species within the propulsion unit of a small scale wind tunnel model is to be studied. A novel experimental approach to obtain absorption spectra of both species in the infrared region within \(\mu \)-second range is reported. This technique is based on the use of frequency combs generated with quantum cascade lasers. This method enables multi-species measurements of absorption spectra from which concentrations can be inferred. The generated frequency combs are within the \(1760\,\pm \,30\,\mathrm cm^{-1}\) wavenumber region which features strong NO and \(H_2O\) spectral lines allowing characterization of these species. The technique, the application in short duration ground testing and its calibration procedure will be reported.
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To show the feasibility of high-speed aircrafts, one needs to have reliable performance indicators and figures of merit. As high-speed vehicles only make sense on long-haul routes, various trajectories and related optimizations were evaluated, covering classical transatlantic routes up to antipodal flights from Europe to Australia. The latter is actually the reference mission for the EC-funded projects LAPCAT I and II, enabling the potential reduction of antipodal flight times to about 4 hours without stopover. As these trajectories are different to classical routes, due to the imposed constraints of over-land flights due to sonic boom, the necessary information on range extension is then used in the as well EC co-funded project HIKARI. This enabled a more reliable dataset for a market analysis carried out by AIRBUS but is not part of study presented in this paper. The trajectory simulations and optimizations presented here are performed for the Mach 8 LAPCAT-MR2 hypersonic cruiser concept, given the GTOW and the amount of fuel on-board. This evaluation is based upon detailed aerodynamic and propulsion databases which on their turn are composed from experiments and numerical simulations.
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As part of the development of the LAPCAT-MR2 hypersonic aircraft, a supercritical supersonic air induction system was designed to provide high total pressure recovery to the auxiliary propulsion system from take-off conditions up to Mach 4.0. The design strategy was progressively established through the reverse engineering of the XB-70 air induction system. Based on the observations carried out throughout this first phase, an adaptation of the LAPCAT-MR2 external layout was studied to ensure supercritical operability and proper mass flow capture of the additional air intake. Due to the very inhomogeneous flow conditions in the captured flow vein, a CFD-based iterative design process was found necessary to converge towards an operable air intake configuration at the design speed of Mach 3.0.
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The propulsion plant of a prospective supersonic cruise aircraft consists of an air turbo–rocket expander and a dual–mode ramjet. A comprehensive numerical model was constructed to examine the performance of the air turbo–rocket during the supersonic acceleration of the vehicle. The numerical model comprised a one–dimensional representation of the fluid paths through the dual–mode ramjet, the air turbo–rocket combustor, the regenerator and the airframe–integrated nozzle, whereas the turbomachinery and the air turbo–rocket bypass were included as zero–dimensional models. The intake operation was based on the results of time–averaged Euler simulations. A preliminary engine analysis revealed that the installation effects restricted significantly the operational envelope, which was subsequently extended bypassing the air turbo–rocket. Hence the engine was throttled varying the mixture ratio and the fan compression ratio. Nevertheless, the performance was optimal when the demand from the air turbo–rocket matched the intake air flow capture. The heat recovery across the regenerator was found critical for the operation of the turbomachinery at low speed. The transition of the air turbo–rocket to ramjet operation was identified at Mach 4.5. During this regime, the propulsion plant was rather insensitive to the mixture ratio and was throttled with the air turbo–rocket throat area.
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Hypersonic airbreathing vehicle designs are challenging due to the complexity of airframe integration and engine cycle design able to operate over a wide Mach number range. Airframe integration needs to deal with the mutual effect of the aerodynamics on the engine performance. On the other hand, the engine performance affects the vehicle design and operation. This paper first presents the followed design methodology including a description of the different approaches used for the layout of a M=8 hypersonic cruiser vehicle. The simulation of combustion process with the Dual Mode Ramjet engine is still based upon on 1D model whereas a dedicated 3D combustor CFD-analysis assesses the reliability of the 1D combustion hypotheses. An assessment of the thrust minus drag balance is carried out on a conceptual vehicle design for different Mach numbers using three-dimensional nose-to-tail CFD simulations. Here, the loosely coupled one-dimensional code for the estimation of the Dual Mode Ramjet performance is applied as well. It could be shown that for the trajectory points investigated both lateral and axial acceleration is possible.
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This paper presents, in the framework of the European program LAPCAT-II, a numerical analysis of the hypersonic Mach 8 vehicle aiming at reproducing its cruise flight conditions. Calculations will focus on the scramjet engine combustor of the full scale vehicle, and will be performed with several CFD codes and analysis tools.
Conference Paper
The objective of the paper is the description and comparison of two different models for the structure weight estimation within the integrated design program PrADO-Hy (Preliminary Aircraft Design and Optimization program - Hypersonic). While the first method consists of one simplified hyersonic loadcases and a beam model to estimate the strucrural stresses, the second one includes subsonic and hypersonic loadings and uses a finite element method to calculate the stress. The fully-stressed design during the structure dimensioning takes into account strength and stability requirements. Both methods are used during the design of the lower airbreathing stage of a two-stage-to-orbit (TSTO) space transportation system. A known TSTO concept having a blended body configuration as its first stage is recomputed to adjust the new weight estimation methods an the programm system PrADO-Hy as a whole. A changeover to a waverider configuration indicates that it should be possible to design a TSTO concept with 11.7% lower maximum take-off weight (MTOW) consuming 35% less fuel compared to blended body reference. Even an error rate of 10% in the structure mass still leads to a convergent waverider design with a 2% higher MTOW. During these investigations both structure mass methods predict nearly the same weights and global sensitivities. But the first method is about five times faster compared to the higher sophisticated second method making the fast version the first choice for the preliminary design process. Remark: The main author of this paper is Dr.-Ing. Andreas Bardenhagen who received his doctor's degree to this theme "Fast weight estimation to hypersonic TSTO waverider concepts in an overall design approach"
Conference Paper
A Mach 8 cruise concept vehicle is defined within the LAPCAT I program to explore the feasibility of antipodal travel. Different engine configurations including Rocket Ejector, Air Turbo Rocket, Ramjet and Scramjet are integrated into a waverider type vehicle and flown as part of a numerical trajectory.
Modeling, Analysis, and Optimization of the
  • I Miranda
  • V Villacé
  • G Paniagua
I. Rodríguez Miranda, V. Fernández Villacé and G. Paniagua, "Modeling, Analysis, and Optimization of the," Journal of Propulsion and Power, p. DOI: 10.2514/1.B34781, 2013.