Conference PaperPDF Available

Aerodynamic Database Development for a Future Reusable Launch Vehicle, the Orbital 500R

Authors:
  • UK Space Agency
  • Fluid Gravity Engineering Ltd

Abstract and Figures

The Orbital 500R is a commercial semi-reusable, two-stage launch system under development by Orbital Access. The focus of this paper is a numerical aerodynamic analysis of the reusable first stage spaceplane, capable of powered and glided flight. The vehicle is intended to release expendable upper stage(s) to inject 500-kilogram of payload into low Earth orbit. Ejection of the upper stage(s) is expected to be above 85-kilometres, after which the first stage spaceplane will perform re-entry and landing. Toolset validation and the establishment of best practice is an integral aspect of assessing and optimising system performance and controllability. Steps have been taken to progress this, related to the use of Fluid Gravity Engineer-ing's Navier-Stokes solver, ANITA. The validation activities were focused around the characterisation of the aerodynamics of two wing-body experimental models, tested across a range of angles of attack at Mach 4.0 and 8.2 respectively. The results have provided a sufficient level of confidence in the use of ANITA to assess the aerodynamic performance of spaceplane wing-body configurations for supersonic and hypersonic flow regimes. A computational assessment of the vehicle's aerodynamics has been performed by means of both engineering-based tools and Navier-Stokes CFD computations. Lessons learned from the validation activities fed into the use of ANITA to characterise the vehicle's aerodynamics from Mach 3.0 and above. The lift and drag characteristics of the vehicle were found to be very comparable to those documented for the X-34 reusable launch vehicle from Mach 2.0 up to Mach 6.0. Furthermore, strong code-to-code agreement was observed with the SU2 and ANSYS-Fluent Navier-Stokes solvers for aerodynamic characteristics at Mach 3.0.
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AERODYNAMIC DATABASE DEVELOPMENT FOR A FUTURE REUSABLE SPACE
LAUNCH VEHICLE, THE ORBITAL 500R
Tristan Stindta, Jim Merrifielda, Marco Fossatib, Lorenzo Ricciardib, Christie Alisa Maddockb,
Michael Westc, Konstantinos Kontisd,Bernard Farkine, Stuart McIntyree
aFluid Gravity Engineering, Emsworth, UK
bUniversity of Strathclyde, Glasgow, UK
cBAE Systems, Prestwick, UK
dUniversity of Glasgow, Glasgow, UK
eOrbital Access, Prestwick, UK
ABSTRACT
The Orbital 500R is a commercial semi-reusable, two-stage
launch system under development by Orbital Access. The fo-
cus of this paper is a numerical aerodynamic analysis of the
reusable first stage spaceplane, capable of powered and glided
flight. The vehicle is intended to release expendable upper
stage(s) to inject 500-kilogram of payload into low Earth or-
bit. Ejection of the upper stage(s) is expected to be above
85-kilometres, after which the first stage spaceplane will per-
form re-entry and landing.
Toolset validation and the establishment of best prac-
tice is an integral aspect of assessing and optimising system
performance and controllability. Steps have been taken to
progress this, related to the use of Fluid Gravity Engineer-
ing's Navier-Stokes solver, ANITA. The validation activities
were focused around the characterisation of the aerodynam-
ics of two wing-body experimental models, tested across a
range of angles of attack at Mach 4.0 and 8.2 respectively.
The results have provided a sufficient level of confidence in
the use of ANITA to assess the aerodynamic performance
of spaceplane wing-body configurations for supersonic and
hypersonic flow regimes.
A computational assessment of the vehicle’s aerodynam-
ics has been performed by means of both engineering-based
tools and Navier-Stokes CFD computations. Lessons learned
from the validation activities fed into the use of ANITA to
characterise the vehicle's aerodynamics from Mach 3.0 and
above. The lift and drag characteristics of the vehicle were
found to be very comparable to those documented for the X-
34 reusable launch vehicle from Mach 2.0 up to Mach 6.0.
Furthermore, strong code-to-code agreement was observed
with the SU2 and ANSYS-Fluent Navier-Stokes solvers for
aerodynamic characteristics at Mach 3.0.
Index TermsSpaceplane, Launch System, Orbital
500R, X-34, Aerodynamics, CFD, Descent Trajectory
1. INTRODUCTION
This paper will present a method for developing an aerody-
namic database suitable for early-phase design studies of con-
ventional spaceplane wing-body geometries. The problem
is approached by integrating different fidelity analysis meth-
ods to populate an aerodynamic database (AEDB) suitable
for three degree-of-freedom trajectory simulations. This ap-
proach is applied to the Orbital 500R spaceplane, which rep-
resents the first stage of a two-stage-to-orbit (TSTO) launch
system under development by Orbital Access.
During the early development stages of a launch system,
the assessment of the concept's aerodynamic performance
is an integral aspect of understanding the feasibility of the
system. Furthermore, the increased geometric complexity
of spaceplane configurations compared to other spacecraft,
such as entry capsules, makes them more susceptible to er-
rors in aerodynamic characterisation. The assessment of the
vehicle's aerodynamic behaviour will largely drive the form
of the descent trajectory. Lift, drag and pitching moments
are the primary aerodynamic parameters to inform on initial
estimates of range capability, controllability and thermal pro-
tection system (TPS) design requirements in terms of heat
fluxes/loads. At these preliminary design stages, the analysis
fidelity requirements also need to be balanced against compu-
tational power requirements to permit rapid design iteration.
As such, the use of validated engineering methods, suitable
across a broad flight envelope, is highly desirable.
2. THE ORBITAL 500R
The Orbital 500R is a multi-stage vehicle, using rocket
propulsion systems, that will be air-launched from a modified,
wide-body carrier aircraft. The main vehicle is a spaceplane
that will allow for glided return flight (see Fig. 1).
The second stage is stored within the main body of the
spaceplane. This allows for better control of the moments
Fig. 1. Orbital 500R spaceplane
induced by the movement of the centre of gravity, however, it
introduces complexity and release issues.
The primary mission of the Orbital 500R system is to
deliver payloads of 500-kilograms to a 600-kilometre Sun
Synchronous Orbit, whether polar or equatorial. The sec-
ondary extended mission is to deliver payloads to a maxi-
mum altitude of 1,200-kilometres, with payloads up to 150-
kilograms. These mission parameters derive from investi-
gations during the previous Future Small Payload Launcher
(FSPL) UK study [1], with the goal of establishing the Orbital
500R as a commercial logistics system for in-orbit delivery.
2.1. Vehicle configuration
The Orbital 500R spaceplane can be classified as a wing-
body. The main body has an overall length of 23.3-metre,
from the nose apex to the trailing edge of the body flap. The
front fuselage ends with a rounded nose constituted by a
quasi-conical shape closed by a 0.6-metre radius hemisphere.
The mid-fuselage is characterised by a quasi-constant, rounded-
rectangular section while the afterbody ends with a truncated
base. The configuration exhibits a double-delta wing con-
figuration with a main 45-degree sweep wing section and
an 80-degree sweep strake. To improve lateral stability, the
configuration has a dihedral of 3.5-degrees. The overall wing
span is 12.6-metres, while the strake root chord is 13.7-
metres and the tip chord 2.5-metres. Full span elevons are
incorporated at the wing trailing edge. Based on NASA Space
Shuttle Orbiter/NASA X-34 concept, a body flap is mounted
aft of the fixed fairing and engine housing for pitch control.
For directional stability and control, a V-tail solution has
been adopted: the two vertical tails have a dihedral angle of
75-degrees, a sweep of 60-degrees and a span of 1.8-metres.
Table 1. Spaceplane configuration comparison
Dimension Orbital 500R
spaceplane X-34
Length, m 23.3 16.5
Wing span, m 12.6 8.50
MAC, m 6.62 4.43
Wing area, m275.0 33.2
The Orbital 500R spaceplane is comparable in size and
planform to the NASA X-34 suborbital vehicle, refer to Ta-
ble 1. Specifically, it has a double-delta wing with the same
sweep angles as the spaceplane. Therefore, the aerodynamic
characterisation of the X-34 provides a useful performance
reference that will be revisited later in the paper.
3. ANITA VALIDATION
Toolset validation and the establishment of a best practice to
evaluate aerodynamic performance is an integral aspect of
assessing and optimising the launch system performance and
controllability. Steps have been taken to progress this with
FGE's Arbitrary Non-equilibrium Implicit Thermochemical
Algorithm (ANITA). This is intended to be the workhorse for
characterising aerodynamic and aerothermal environments
above Mach 3.0 as the Orbital 500R design matures.
ANITA is an arbitrary cell, unstructured mesh, Navier-
Stokes solver for gases in thermochemical non-equilibrium. It
is designed for the calculation of aerodynamics and aerother-
modynamics at hypersonic and supersonic speeds. An ex-
ample of a past use case of the code in hypersonic flows is
provided in Ref. [2].
A set of validation studies were conducted using experi-
mental results from wind tunnel campaigns performed on two
wing-body configurations. Fig. 2 provides an example of the
ANITA simulations performed on these wing-body models.
An overview of cases is given in Table 2.
Table 2. ANITA validation cases
Parameter Singh [3] Pike [4]
Mach 8.2 4.0
Reynolds no. x1061.65 2.99
Angle of attack, deg 0,3,5,7,10 0,5,10,15,20
The management of CFD uncertainties is one of the main
challenges in the design process. By establishing credible
CFD results on relevant test cases, the analysis in this sec-
tion can be used to inform on the expected level of accuracy
achievable from the use of ANITA in the aerodynamic char-
acterisation of the Orbital 500R spaceplane.
The logic behind the testcase selection was two-fold.
Firstly, it was to target experimental campaigns that inves-
tigated flow conditions and geometries relevant to the most
important aspects of the Orbital 500R spaceplane and its de-
scent trajectory and, therefore, which exhibited flow fields
governed by the same physical phenomena. Secondly, it was
to evaluate this experimental data, to identify the sources
for which test conditions and geometries were well defined
and all experimental measurement accuracies and any post-
processing were well described.
Concerning the first selection criterion, the validation ac-
tivities presented have targeted the high-supersonic/hypersonic
portion of the descent trajectory, owing to the dominance of
(a) Symmetry plane Mach number contours
of the Pike wing-body at α= 20-deg
(b) Mach number contours of the Singh wing-body
at Y= 0.025-metres at α= 10-deg and δflap = 25-deg
Fig. 2. Wing-body flow fields from ANITA simulations
this regime regarding the duration of flight at these conditions.
In addition, potentially large angles of attack may be required
under these conditions. The physics of interest relates to the
strong bow shocks generated off the nose and leading edges
of the lifting surfaces and the resulting shock/vortical inter-
actions, in particular at the wingbody junctions. Capturing
the angle at which shock detachment increases significantly
at the wing leading edge will also play a pivotal role in how
the flow field evolves around the wing.
The challenge of assessing subsonic and transonic aero-
dynamic performance represents a major milestone in the es-
tablishing the controllability of a spaceplane configuration.
Vortex-induced effects, leeside separation and the movement
of the centre-of-pressure and sonic lines all contribute towards
making the transonic regime extremely challenging to model.
However, at this stage in the Orbital 500R design and devel-
opment, neither angle of sideslip nor elevon, aileron or body
flap deflection angles have been considered. Therefore, the
validation activities were focused on higher speed regimes,
where control surface authority was not deemed as critical.
3.1. CFD setup
The following section provides a brief overview of the CFD
strategy used on the experimental testcases. A summary of
the physical and numerical models employed in ANITA is
provided in Table 3.
Pointwise™ was used for grid generation. This offers
structured, unstructured, hybrid and overset meshing tech-
niques and a highly automated T-Rex (anisotropic tetrahe-
dral extrusion) technique for boundary layer resolved hybrid
meshes. The generation of the volume grids was performed
using the orthogonal advancing front marching process to
grow tetrahedral cells on the boundary grid. A consistent ap-
proach was adopted for the grid development of each model.
Table 3. ANITA setup for validation activities
Physical/numerical model Description
Viscosity Sutherland law
Conductivity Constant Prandtl
Fluid Calorifically perfect
Turbulence kωSST
Flux scheme HLLC or AUSMDV
Flux reconstruction Second order MUSCL
A grid convergence study was performed on a series of three
increasingly refined grids: coarse, intermediate and refined.
The computational domains consisted of an adjustable near-
field block, encompassed by far-field block. An example of
this for the Pike wing-body is depicted in Fig. 3.
Fig. 3. Pike wing-body computational domain definition: ad-
justable vehicle and near-field angle
Fig. 4. Pike wing-body grid convergence pitching moment
coefficient results
The results of the grid convergence study on the pitching
moment for the Pike wing-body are shown in Fig. 4.
3.2. Description of validation testcases
Singh conducted an experimental investigation of the hy-
personic flow over a wing-body configuration, for a range
of trailing edge flap deflections from 0 to 25-degrees [3].
The present analysis focused on the 25-degree flap deflection
case. The tests were conducted in Cranfield University's gun-
tunnel. The configuration is characterised by a sphere-cone
forebody, followed by a cylindrical fuselage accompanied
with a 70-degree sweep delta wing with trailing edge flaps.
Pike conducted a series of tests in the 3x4-foot Tunnel
(H.S.S.T.) at R.A.E. Bedford to investigate the forces and
flow on a range of wing-body models at Mach 4.0 [4]. The
model selected for the validation activity is characterised
by a ogive-cylinder body and wings in the plane containing
the axis of symmetry of the body. The planform has a lat-
eral plane of symmetry with equal leading and trailing edge
sweeps of 26.5-degrees. Boundary layer transition was fixed
by a strip of 36 grade Carborundum to 10%chord.
3.3. Validation results
Fig. 5 shows the agreement found using ANITA to calculate
the pitching moment coefficients across the range of angles of
attack for both wing-body configurations.
The ANITA results show an agreement of within 5%
and 10%of Singh's and Pike's experimental pitching mo-
ment measurements respectively (excluding the comparisons
at 0-degrees). These findings should be interpreted with an
understanding of the main sources of uncertainty, believed to
arise from experimental error and difficulties with digitising
the source material. For Singh's experimental campaign, an
error analysis performed by Opatowski [5] showed that the
maximum possible error for the three balance channels was:
normal force ±3.5%, axial force ±6% and pitching moment
(a) Pike wing-body
(b) Singh wing-body
Fig. 5. Wing-body pitching moment characteristics
±3.5%. The measured forces obtained by Pike were cor-
rected for balance interactions, giving an estimated accuracy
of: lift coefficient ±0.003, drag coefficient ±0.005, pitching
moment coefficient ±0.005. For both campaigns, very good
repeatability of the results was reported.
Further uncertainties have inevitably been introduced into
the analysis during the digitisation of the experimental results.
This will be more significant for the smaller pitching moment
coefficients at the lower angles of attack. An uncertainty in
the region of several percent is deemed a reasonable approxi-
mation to this error.
All things considered, agreements of 5 to 10%provide
sufficient confidence in the use of ANITA for the Orbital
500R spaceplane, when implemented in the configuration
detailed above and using lessons learned from the grid devel-
opment strategies for both wing-body configurations.
4. AERODYNAMIC CHARACTERISATION OF THE
ORBITAL 500R SPACEPLANE
The aerodynamic characterisation of the Orbital 500R space-
plane used a design space approach. A multi-dimensional
space, defined by Mach number, angle of attack and altitude,
was constructed to accommodate all likely operational trajec-
tory dispersions. The results could then be incorporated in tra-
jectory optimisation routines for the spaceplane. The design
space was defined by a series of boxes, detailed in Table 4.
Table 4. Definition of the design space
Box Mach Angle of attack, deg Altitude, km
1 0 to 3 -10 to 25 0 to 50
2 3 to 4 -10 to 35 5 to 50
3 4 to 10 -5 to 45 20 to 60
4 6 to 10 -5 to 45 60 to 120
The intention of the chosen analysis approach was to gen-
erate a broad assessment of the aerodynamics to a fidelity
suitable for early-phase design studies. Engineering meth-
ods were used to a greater extent (when compared with CFD)
to populate the AEDB. The engineering methods were cali-
brated to the higher fidelity CFD results. The choice of the
CFD cases was made with the intention of targeting regions
where rapid changes are expected, to explore the linear or
nonlinear dependence of the aerodynamic environments on
Mach number, angle of attack and altitude. At this stage in the
spaceplane design and development, neither angle of sideslip
nor elevon, aileron or body flap deflection angles have been
considered.
The design space spans a range of different flow regimes;
ranging from free molecular at very high altitudes to contin-
uum deep in the atmosphere. The physical models are differ-
ent in these regimes. In particular, for slender vehicles at low
angles of attack, the free molecular drag will be significantly
larger than the continuum level. This is worthy of considera-
tion, owing to the shallow nature of the entry profile because
of the vehicle's high L
/D. Therefore, a considerable period is
spent in the higher altitude regime.
A broad overview of the lift and drag characteristics gen-
erated across the flight envelope is given in Fig. 6. This is
the result of a collaboration between FGE, BAe and the Uni-
versities of Glasgow and Strathclyde and each organisations
respective analysis toolset and approach. The focus of this
paper is to present the application of FGE's methodology and
computational tools to flow regimes from Mach 3.0 upwards.
4.1. Panel method analysis
Engineering-level aerodynamic analysis was performed using
FGE's ASPEN panel method which computes free molecu-
lar and continuum aerodynamics for arbitrary geometries, for
high-supersonic/hypersonic speeds. The simplified approach
implemented within ASPEN, necessitates an awareness its
limitations. These include a lack of capabilities to handle flow
separation, real gas effects and flow interaction effects.
The geometry of the vehicle is represented by a system
of triangular panels. The only parameter required to calculate
the pressure is the impact angle of the free stream flow with
the panel. The surface elements that see the oncoming flow
directly are said to be in the impact region, whilst the elements
shielded from the flow are in the shadow region. Depending
on the region the element falls in along with the nature of
its surrounding geometry, an appropriate method can be ap-
plied. Accordingly, the spaceplane was divided into separate
analysis regions. A description of the panel method imple-
mentation is provided in Table 5. All regions of high cur-
vature/bluntness were automatically assessed using the mod-
ified Newtonian approach.
Table 5. Local surface inclination methods used for the as-
sessment of the Orbital 500R spaceplane aerodynamics
Component Impact flow Shadow flow
Nosecone Modified Newtonian
and Prandtl-Meyer N/A
Fuselage Tangent wedge Prandtl-Meyer
Base N/A p/3
Inboard delta Delta wing empirical Prandtl-Meyer
Outboard delta Delta wing empirical Prandtl-Meyer
V-tail Delta wing empirical Prandtl-Meyer
Body flap Tangent wedge Prandtl-Meyer
Modified Newtonian, tangent wedge and Prandtl-Meyer
methods are well reported in the literature. The ‘Delta wing
empirical’ method, that has been applied to the lifting sur-
faces, was derived by Gentry et al. [6] who performed experi-
ments on blunt, highly swept delta wings in hypersonic flows.
The models showed behavioural characteristics in accordance
with tangent wedge theory for angles of attack between 5-
degrees and 15-degrees. For angles above 15-degrees, the
nature of the flow appears to change such that the tangent
cone approximation appears valid up to 40-degrees angle of
attack. In accordance with this, Gentry et al. [6] devised an
approximate method that was found to perform well on the
spaceplane configuration.
4.1.1. Viscous bridging
ASPEN has an in-built capability to compute the approximate
viscous contribution based on flat plate theory with compress-
ibility corrections. The local skin friction coefficient, cf, for
incompressible laminar and turbulent flows over a flat plate is
given by Eq. (1) and Eq. (2) respectively.
cf,lam =0.664
Rex
(1) cf,turb =0.0592
Re0.2
x
(2)
where Rexis the running-length Reynolds number based
on boundary layer edge properties. The compressibility cor-
rections have been applied to Eq. (1) and Eq. (2) using the
Fig. 6. Lift (left) and drag (centre) coefficients, values at selected altitudes. CL(right-top) and CD(right-bottom) variations
with altitude at 0-deg and 20-deg
Eckert reference temperature and Van Driest II methods re-
spectively. Summation of the panels' local shear stresses,
resolved in the vector defined by the appropriate stagnation
point to the panel centre, permits an estimation of the global
shear stress. This model was subsequently calibrated to vis-
cous Navier-Stokes solutions from ANITA at a series of tra-
jectory points.
4.1.2. Free-molecular bridging
For initial scoping studies, as applicable to the current stage
of development of the Orbital 500R spaceplane, ‘bridging
methods have been used. These methods implement em-
pirically derived expressions to cover the transitional flow
regime, linking continuum to free molecular flow regimes.
Consequently, both the continuum and free molecular aero-
dynamic databases are required to generate the intermediate
transitional data. The calculation of aerodynamics in free
molecular flow has been conducted using the pressure and
shear equations developed by Schaaf [7].
The main attraction of using bridging functions to char-
acterise the vehicle force, moment and heat transfer levels is
their relative simplicity and ease with which they can be ap-
plied to determine the vehicle aerodynamic loads. However,
they are known to be highly geometry dependant.
The bridging function was selected as appropriate to
recreate the data within the Knudsen range 0.003 to 10,
which encompasses all of the altitudes above and including
80-kilometres, up through to the highest altitude considered
of 120-kilometres. Fig. 7 shows the classical, expected ‘s-
Fig. 7. High altitude lift and drag characteristics of the Or-
bital 500R spaceplane generated using calibrated engineering
methods with a free molecular bridging function incorporated
shape’ profiles exhibited by the lift and drag coefficients with
Knudsen number, Kn. Note that the characteristic length is
taken as the base height of 1.8-metres. This will be sensitive
to body size and will be significantly lower in altitude for
smaller, more detailed geometric regions. The form of the
bridging relation used in this study, in the case of the lift
coefficient, is:
CL=CL,cont +φ(CL,fm CL,cont)(3)
Where cont denotes continuum coefficients, fm stands
for free molecular and φ=Knn
/Knn+k. Values of n= 1
and k= 0.1have been used. Exactly the same approach de-
scribed by Eq.(3) was used to determine the drag coefficient.
This is what is termed a global bridging approach (i.e. the
coefficients are bridged directly, not the local pressures/skin
frictions).
4.2. CFD analysis
The lessons learned during the validation activities, presented
in Sec. 3, were used to inform on the strategy chosen to per-
form the ANITA simulations on the Orbital 500R spaceplane.
Pointwise™ was used for grid generation, to take advan-
tage of its hybrid meshing technique, discussed in Sec. 3.1.
The wall cell spacings ensured a y+value of less than one. A
similar grid convergence procedure to that outlined in Sec. 3.1
was performed, with the use of inviscid simulations at the out-
set to determine the shock shape and guide the further devel-
opment of the viscous grids. The final computational grids are
composed of about 2.5 million cells for the Eulerian versions
and 4.2 million cells for the viscous ones. The computational
grid, in the vicinity of the vehicle, used for the symmetric,
half body computations is shown in Fig. 8.
Fig. 8. Orbital 500R spaceplane computational grid
At high-supersonic/hypersonic speeds, the flow field is
dominated by strong bow shocks generated off the nose
and leading edges of the lifting surfaces and the resulting
shock/vortical interactions, in particular at the wing- body
junctions. As such, aerodynamic efficiency is affected to
a much greater extent by surface conditions of the wind-
ward rather than leeward regions of components because of
the stark differences in pressures. Geometrical entities with
small curvature radii inherently result in steep pressure gradi-
ents. The aforementioned reasoning motivated the approach
towards refining the resolution in specific regions of the ve-
hicle to sufficiently capture the flow physics. The windward
surface resolution was particularly important to accurately
compute the pitching moments.
4.3. Code-to-code validation
A code-to-code validation was performed at Mach 3.0, 20-
kilometres altitude across a range of angles of attack. In
addition to ANITA, the Navier-Stokes simulations were per-
formed by BAe using the ANSYS-Fluent solver and by the
University of Strathclyde using Stanford University’s SU2
solver. Independently generated computational grids were
used. Fig. 11 shows examples of the flow fields from the SU2
simulations. A variation of less than 1% was found for the lift
and drag coefficients. Similarly, Fig. 9 highlights an excellent
alignment of the pitching moment coefficients (referenced
around an estimated CoG).
Fig. 9. Pitching moment coefficient code-to-code comparison
referenced around a realistic CoG location
4.4. NASA X-34 comparison
Similarities between the X-34 and the Orbital 500R space-
plane, as outlined in Sec. 2.1, mean that it offers a valu-
able validation case in the supersonic and hypersonic flow
regimes. A remarkably good alignment of lift and drag coef-
ficients was found from Mach 2.0 up to Mach 6.0, to Brauck-
mann’s published results [8]. The comparison at Mach 6 is
shown in Fig. 10.
Fig. 10. Comparison between the Orbital 500R spaceplane
and the X-34 lift and drag characteristics
Fig. 11. Mach 3.0 contours for α=10-deg (left), α= 7.5-deg (centre), α= 25-deg (right) for Y= 4.263-metres
5. CONCLUSION
This paper has presented the appraisal of the Orbital 500R
spaceplane's aerodynamics in a computationally inexpensive
manner. This has been achieved through the effective im-
plementation engineering methods calibrated with validated
Navier-Stokes computations for a calorifically perfect gas
in both laminar and turbulent conditions. Bridging relations
have enabled a characterisation of the aerodynamics from the
continuum up to the free molecular regime. The vehicle's
configuration exhibits notable similarities to NASA’s X-34
reusable launch vehicle, therefore, it provided a useful vali-
dation case. The lift and drag characteristics of the Orbital
500R spaceplane were found to be very comparable to those
published for the X-34 from Mach 2.0 up to Mach 6.0. Fur-
thermore, strong code-to-code agreement with ANITA was
observed with the ANSYS-Fluent and SU2 Navier-Stokes
solvers for aerodynamic characteristics at Mach 3.0.
Validation of FGE's Navier-Stokes solver ANITA, that
was used for the Orbital 500R AEDB CFD computations
for Mach 3.0 upwards, has been performed on two simple
wing-body configurations at Mach numbers of 8.2 and 4.0
respectively. Extremely good alignment of the aerodynamic
coefficients was observed throughout. In particular, pitching
moment coefficient agreements within 10% were observed
across the range of angles of attack considered, with the
primary uncertainty drivers being experimental error and in-
accuracies associated with the data digitisation. This has
provided a sufficient level of confidence in the use of ANITA
to assess the aerodynamic performance of spaceplane wing-
body configurations in supersonic and hypersonic flows.
6. ACKNOWLEDGMENTS
This work has been partially funded by the UK Space Agency
and European Space Agency (ESA) General Support Tech-
nology Programme (GSTP).
7. REFERENCES
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[3] A. Singh, “Experimental study of slender vehicles at hy-
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[4] J. Pike, “Wind tunnel tests on six wing-body models at
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[5] T. Opatowski, “Gun tunnel force measurements on some
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229–239, March 1999.
... C p distributions on a cuboid in a flow with M = 9 and Re = 656 × 10 3 as generated by the ATD3 working group and published in [157]. TAU [63], MISTRAL [151], SU2 [49], and ANITA [161] are all high fidelity CFD codes, while PAMPERO and SAM are satellite demise prediction tools which use engineering models to predict heat fluxes. Note the wedge-shapes regions of increased pressure along the streamwise edge downstream of the sharp corner on the stagnation surface. . ...
... [157]. TAU [63], MISTRAL [151], SU2 [49], and ANITA [161] are all high fidelity CFD codes, while PAMPERO and SAM are satellite demise prediction tools which use engineering models to predict heat fluxes. Note the wedge-shapes regions of increased pressure along the streamwise edge downstream of the sharp corner on the stagnation surface. ...
Thesis
Full-text available
The goal of the work presented in this thesis is to improve the aerothermodynamic heating models used in modern satellite re-entry analysis tools. The current generation of heating models are usually based on correlations developed for shapes such as flat plates and hemispheres. These overly simplified models can lead to inaccurate predictions of the ground casualty risk associated with a re-entry event. In order to derive new heating models for shapes more representative of satellite geometries, this work uses a combination of CFD and wind tunnel measurements to study the heat fluxes experienced by a cuboid at two different orientations in a Mach 5 flow and at a range of Reynolds number conditions. 2D and 3D CFD simulations showed that the hypersonic flow around a cuboid geometry has a strong dependence on Reynolds number, with a breakaway separation bubble forming from the windward expansion edge at very high Reynolds numbers. This separation bubble can significantly lower the heat fluxes underneath it. In addition to this separation bubble, the 3D CFD simulations showed that the sharp corners on a 3D geometry can generate streamwise vortical structures along the streamwise edges of a cuboid. The surface heat fluxes induced by these structures can be as high as the stagnation point heating value and could therefore play a significant role in satellite fragmentation during re-entry. Finally, the CFD simulations suggested that the highest heat flux values experienced by the cuboid occur at the sharp edges and corners of the geometry. However, the heat flux predicted by CFD in these regions is seemingly non-physically large. The CFD simulations were experimentally validated with wind tunnel measurements of the Stanton number distribution over a cube. The Stanton number measurements were obtained by recording the temperature history of the wind tunnel model using infrared thermography and then calculating the convective heat flux using a 3D inverse heat conduction solver. The experimental Stanton number values generally showed very good agreement with the CFD results, and comparisons of both the stagnation point Stanton number and the average heating experienced by the cuboid were favourable. However, in contrast to the CFD results, the experimental measurements did not show any region of significantly increased heat flux near the sharp edges of the cuboid. We propose that this is likely due to abreakdown of the continuum assumption in these regions, which cannot be captured with conventional CFD. The combined CFD and experimental results were then compared to two currentgeneration satellite re-entry prediction tools, DRAMA and SAM. These comparisons showed that DRAMA overpredicts the average heat flux to a cuboid by 65-75% depending the flow conditions and cube orientation. On the other hand, the average heat fluxes predicted by SAM agreed well with the experimental and CFD values. Despite SAM’s success at predicting average heat flux values, comparisons of Stanton number distributions over the surface of the geometry showed that the tool did not predict the regions of increased heat flux associated with the streamwise vortex structures generated by the 3D expansions around the cube corner, while at the same time over-predicting the heat flux to other regions of the cube geometry. Future generations of re-entry prediction tools will need to be able accurately predict the Stanton number distributions across entire satellite geometries. This is important because different fragmentation phenomena may occur depending on which satellite components fail first.
... Surrogate models were used for the lift c L and drag c D coefficients as a function of Mach number M, angle of attack α and altitude h. An artificial neural network was employed, using an aerodynamic database for the training data coming from a mix of panel methods and CFD simulations (see Stindt et al. [32] for details on the generation of the aerodynamic database). The lift and drag forces were computed assuming no wind, with the relative velocity v rel = v − ω E r. ...
Article
Full-text available
This paper presents a novel method for multi-objective optimisation under uncertainty developed to study a range of mission trade-offs, and the impact of uncertainties on the evaluation of launch system mission designs. A memetic multi-objective optimisation algorithm, named MODHOC, which combines the Direct Finite Elements in Time transcription method with Multi Agent Collaborative Search, is extended to account for model uncertainties. An Unscented Transformation is used to capture the first two statistical moments of the quantities of interest. A quantification model of the uncertainty was developed for the atmospheric model parameters. An optimisation under uncertainty was run for the design of descent trajectories for a spaceplane-based two-stage launch system.
Thesis
Full-text available
Thesis (Ph. D.)--Cranfield University, 1996.
Article
The X-34, being designed and built by the Orbital Sciences Corporation, is an unmanned suborbital vehicle designed to be used as a flying test bed to demonstrate key vehicle and operational technologies applicable to future reusable launch vehicles. The X-34 will be air-launched from an L-1011 carrier aircraft, where an onboard engine will then accelerate the vehicle to speeds above Mach 7 and altitudes to 250,000 ft. An unpowered entry will follow, including an autonomous landing. A series of wind-tunnel tests on scaled models was conducted in four facilities at the NASA Langley Research Center to determine the aerodynamic characteristics of the X-34. Analysis of these test results revealed that longitudinal trim could be achieved throughout the design trajectory. Longitudinal control effectiveness (elevon, body flap) decreased for Mach numbers greater than one and was nonlinear in the transonic regime. In addition, there was a loss in elevon effectiveness for the largest negative deflection. The rudder effectiveness (of the all-moving vertical tail) decreased for supersonic Mach numbers and with increasing angle of attack, so that at combined high supersonic Mach numbers and high angles of attack the use of reaction control jets may be required. Deployment of the landing gear reduced the elevon effectiveness by 5% and destabilized the vehicle directionally.
Article
The volume contains the source language listings of the Mark IV supersonic-Hypersonic Arbitrary-Body Program (Mod 0 Version). The program as shown in this listing will operate on CDC 6500, 6600, and CYBER 74 computers. With a small converter program, the Mark IV program can be converted for operation on IBM 360 and 370 types of computers. This converter program is included with the listings.
Vehicle and mission design of a future small payload launcher
  • C Maddock
  • L Ricciardi
  • A Mogavero
  • S Rengarajan
  • K Kontis
  • A Milne
  • J Merrifield
  • M West
  • S Mcintyre
C. Maddock, L. Ricciardi, A. Mogavero, S. Rengarajan, K. Kontis, A. Milne, J. Merrifield, M. West, and S. McIntyre, "Vehicle and mission design of a future small payload launcher," in 21st AIAA International Space Planes and Hypersonic Systems and Technologies Conference. AIAA, 2017.
Navier-stokes computations for hollow cylinders and cubes relevant for satellite debris
  • B Couchman
  • M Haynes
B. Couchman and M. Haynes, "Navier-stokes computations for hollow cylinders and cubes relevant for satellite debris," in 8th European Symposium on Aerothermodynamics for Space Vehicles, March 2015.
Wind tunnel tests on six wing-body models at M=4," in Aeronautical Research Council Current Papers, Ministry of Technology
  • J Pike
J. Pike, "Wind tunnel tests on six wing-body models at M=4," in Aeronautical Research Council Current Papers, Ministry of Technology. Aerodynamics Dept., R.A.E., Bedford, 1969.
Gun tunnel force measurements on some thin delta wings suitable for cruising flight
  • T Opatowski
T. Opatowski, "Gun tunnel force measurements on some thin delta wings suitable for cruising flight," Aeronautical Research Council, 1971.