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Experimental characterization of the hypersonic flow around a cuboid

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Understanding the hypersonic flow around faceted shapes is important in the context of the fragmentation and demise of satellites undergoing uncontrolled atmospheric entry. To better understand the physics of such flows, as well as the satellite demise process, we perform an experimental study of the Mach 5 flow around a cuboid geometry in the University of Manchester High SuperSonic Tunnel. Heat fluxes are measured using infrared thermography and a 3D inverse heat conduction solution, and flow features are imaged using schlieren photography. Measurements are taken at a range of Reynolds numbers from $${40.0 \times 10^3}$$ 40.0 × 10 3 to $${549 \times 10^3}$$ 549 × 10 3 . The schlieren results suggest the presence of a separation bubble at the windward edge of the cube at high Reynolds numbers. High heat fluxes are observed near corners and edges, which are caused by boundary-layer thinning. Additionally, on the side (off-stagnation) faces of the cube, we observe wedge-shaped regions of high heat flux emanating from the windward corners of the cube. We attribute these to vortical structures being generated by the strong expansion around the cube’s corners. We also observe that the stagnation point of the cube is off-centre of the windward face, which we propose is due to sting flex under aerodynamic loading. Finally, we propose a simple method of calculating the stagnation point heat flux to a cube, as well as relations which can be used to predict hypersonic heat fluxes to cuboid geometries such as satellites during atmospheric re-entry. Graphic abstract
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Experiments in Fluids (2020) 61:151
Experimental characterization ofthehypersonic ow aroundacuboid
ThomasW.Rees1 · TomB.Fisher2· PaulJ.K.Bruce1· JimA.Merrield3· MarkK.Quinn2
Received: 20 December 2019 / Revised: 9 April 2020 / Accepted: 9 May 2020 / Published online: 12 June 2020
© The Author(s) 2020
Understanding the hypersonic flow around faceted shapes is important in the context of the fragmentation and demise of
satellites undergoing uncontrolled atmospheric entry. To better understand the physics of such flows, as well as the satellite
demise process, we perform an experimental study of the Mach 5 flow around a cuboid geometry in the University of Man-
chester High SuperSonic Tunnel. Heat fluxes are measured using infrared thermography and a 3D inverse heat conduction
solution, and flow features are imaged using schlieren photography. Measurements are taken at a range of Reynolds numbers
40.0 ×103
549 ×103
. The schlieren results suggest the presence of a separation bubble at the windward edge of the
cube at high Reynolds numbers. High heat fluxes are observed near corners and edges, which are caused by boundary-layer
thinning. Additionally, on the side (off-stagnation) faces of the cube, we observe wedge-shaped regions of high heat flux
emanating from the windward corners of the cube. We attribute these to vortical structures being generated by the strong
expansion around the cube’s corners. We also observe that the stagnation point of the cube is off-centre of the windward
face, which we propose is due to sting flex under aerodynamic loading. Finally, we propose a simple method of calculating
the stagnation point heat flux to a cube, as well as relations which can be used to predict hypersonic heat fluxes to cuboid
geometries such as satellites during atmospheric re-entry.
Graphic abstract
Extended author information available on the last page of the article
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151 Page 2 of 22
1 Introduction
A half century of human space flight and exploitation has
resulted in approximately 6000 satellites being placed in
orbit around the Earth. Due to their limited operational
lifespan, fewer than one in six of those satellites are still
operational, leaving a large number of decommissioned
satellites in orbit around the Earth (ESA’s Annual Space
Environment 2019). Due to the limited availability of
Earth orbits, the presence of decommissioned satellites
in space increases the risk of in-orbit collisions and the
associated risks of space debris. To address these prob-
lems, satellites must be disposed of at their end-of-life.
The method of disposal generally depends on the orbit
the satellite is placed in. Due to their high altitudes, satel-
lites in Geostationary Earth Orbit (GEO) are often raised
to a graveyard orbit well away from the other common
orbits. On the other hand, decommissioned satellites in
Low Earth Orbit (LEO) are often left alone at the end
of their life, as their low orbits will gradually decay due
to the atmospheric drag experienced in LEO. Eventually
this results in an uncontrolled atmospheric re-entry. For
smaller satellites, a re-entry event will induce temperatures
and forces large enough to destroy the satellite, with with
only limited parts making groundfall. As the satellite size
increases however it becomes more and more likely that
significant satellite mass will hit the ground.
Satellite debris impacting the Earth carries with it a
casualty risk which, although small for any given re-entry
event, cannot be neglected due to the sheer number of sat-
ellites in orbit above the Earth. As a result, it is generally
accepted that space users have a duty to minimize the risks
associated with re-entry events (Merrifield etal. 2014). To
this end, the European Space Agency issued an instruction
in 2014 that the casualty risk for any re-entry event should
be no greater than 1 in
(Dordain 2014). A number
of other national space agencies, including NASA, also
adhere to this figure (IADC 2007). Estimates of the ground
casualty risk associated with re-entry are calculated using
dedicated tools (Koppenwallner etal. 2005; Martin etal.
2005) which must take into account the number of objects
involved, their fragmentation and demise mechanisms, the
effective cross-sections of the surviving components, their
most likely locations as they hit the ground, as well as an
accurate population density map of the Earth.
In particular, there remains considerable uncertainty
in predictions of the aerothermodynamic heating rates
induced by the hypersonic flow around satellites dur-
ing re-entry. This is largely due to the fact that satellite
geometries are significantly different from most other re-
entry bodies—they are typified by sharp corners, facets,
and multi-scale structures. These features cause strong
expansions and compressions in the flow around the sat-
ellite, significantly thinning or thickening the boundary-
layer and therefore increasing or decreasing local heating
rates. Beyond the obvious importance of understanding
what the maximum heat flux to a body is, some recent
studies have suggested that satellite fragmentation mecha-
nisms are driven by failure of fasteners and glues rather
than melting of body panels (Soares and Merrifield 2018).
As these components are often located near corners and
edges, fully understanding the heating rates at these loca-
tions is particularly important.
The fundamental roadblock to a better understanding of
the re-entry heating rates to satellites is that there is very
little freely available high-fidelity data, either experimental
or numerical, of hypersonic aerothermal heating to faceted
shapes such as cuboids, plates, or cylinders. Heating rates
to flat-ended cylinders were analysed experimentally and
theoretically in Eaves (1968), Inouye etal. (1968), Klett
(1964), Kuehn (1963), andMatthews and Eaves (1967). In
particular, the work of Matthews and Eaves (1967) iden-
tified that, at certain conditions, a separation bubble can
form immediately downstream of a cylinder’s expansion
edge. Unfortunately, there are no data available for heating
rates under the separation bubble. Nevertheless, the authors
suggested that these heat fluxes, and the separation bubble
formation, are highly dependent on the Reynolds number. A
2D CFD study investigating the effect of Reynolds number
on the hypersonic flow around faceted shapes confirmed that
the formation of such a separation bubble was dependent on
Reynolds number (Rees etal. 2018), and that the presence of
a separation could significantly decrease local heating rates.
This reduction in local heating is especially significant in the
context of satellite demise as it will result in an increased
casualty risk. In recent free-flight experiments of a cube in
a hypersonic flow (Seltner etal. 2019), the authors claimed
the presence of a separation at the leading edge of the cube,
but the resolution of the schlieren was not high enough to
capture it in detail. In addition to cylinders, heating rates
to cuboids have also been studied in the reports of Crosby
and Knox (1980), and Laganelli (1980), who experimentally
measured heat fluxes to a cube in a Mach 8 flow at discrete
locations using thin-foil calorimeters. However, these stud-
ies only report results at one flow condition and at limited
discrete locations on the model surface.
This scope of this work is the experimental study of
the hypersonic flow around a cuboid shape, with empha-
sis placed on leveraging modern measurement techniques
to obtain accurate heat flux measurements, especially near
the corners and edges of the geometry. To achieve these
high-fidelity heat flux measurements, the temperature field
is measured over the entire surface of the wind tunnel model
using InfraRed Thermography (IRT), and the heat flux is
calculated by the solution of a three-dimensional inverse
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Experiments in Fluids (2020) 61:151
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heat conduction problem (3D-IHCP). In this way, the Stan-
ton number,
is calculated at every point on the surface
of the cube, including in the regions closest to the corners
and edges. The use ofIRT is advantageous due to the fact
that it is a non-intrusive measurement technique, with a high
spatial resolution, low response time, and high sensitivity. In
addition to the IRT measurements, the flow structure around
the cube is imaged with schlieren photography to further
study the separation formation described in Matthews and
Eaves (1967), Rees etal. (2018), andSeltner etal. (2019).
2 Flow facility, models, andtest conditions
2.1 High SuperSonic tunnel (HSST)
Experiments have been performed in the University of Man-
chester’s High SuperSonic Tunnel (HSST), based in the
department of Mechanical, Aerospace and Civil Engineer-
ing. HSST is a long-duration blow-down facility with an
electric resistive heater and a swappable nozzle. A schematic
diagram of the wind tunnel is presented in Fig.1, and a table
of the achievable flow conditions with a Mach 5 nozzle are
presented in Table1. Optical access to the working section is
provided by two parallel, rectangular quartz windows which
span the full length of the useful test jet. Infrared access is
afforded via a
75 mm
diameter uncoated germanium win-
dow. Fixed model mounting positions are provided by an
arc-balance sting, which allows the model to be mounted
at angles of attack of
. Model orientation on the sting
is afforded by a keyway and grub-screw arrangement. A
detailed description of the tunnel and its operation can be
found in Erdem (2011), andFisher (2019).
2.2 Schlieren
Schlieren images were acquired through Töpler’s Z-type
schlieren method. Two 12 in.diameter f/7.9 mirrors pass the
light from a Newport optics model 66921 Xenon arc lamp,
typically at 450 W, onto the knife-edge in the cut-off plane
which is then focused though a
500 mm
focal length ach-
romatic doublet lens onto the camera sensor. The images
are captured with a commercial Nikon D5200 24-megapixel
DSLR camera.
2.3 Models, materials, andmounting
Two model geometries are tested: a
length cube
and a
diameter hemisphere, with the hemisphere
model being used to validate the IRT and heat flux calcula-
tion techniques (see Sect.6.1). The models were mounted
to a sting adaptor, which allows the model to fit to the
arc-balance sting (Fig.2). By swapping the sting adap-
tor, the models can be mounted in different roll orienta-
tions, allowing different facets of the cube to be imaged
by the IR camera and schlieren. For IRT measurements,
the cube model is oriented in a rolled
orientation such
that three surfaces of the cube are imaged simultaneously
(see Fig.5a for a sample IR image of a cube model). In this
way, temperature data at a corner of the cube are obtained.
Table 1 HSST characteristic flow conditions with a Mach 5 nozzle
Parameter Min Max
range [K] 320 950
range [kPa] 200 850
Run time [s] 0.5 7.5
Enthalpy [kJ/kg] 19.8 654
Re ×105
[m−1]9.69 226
Test jet diameter [m] 0.152
Test gas Air
Fig. 1 Schematic diagram of the HSST facility
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For schlieren measurements, the cube is mounted in an
unrolled orientation so that all relevant flow structures can
be imaged.
The models are manufactured from MACOR®, a machin-
able glass-ceramic. It was chosen due to its favourable ther-
mal properties, high emissivity, success in previous experi-
mental hypersonic IRT applications (Cardone etal. 2012),
and ease of machining. Measurements of the temperature
variation of MACOR’s thermal properties are known from
Imbriale (2013), and are plotted in Fig 3. The directional
emissivity variation of MACOR has been reported in Car-
done etal. (2012). Imbriale (2013) correlated this data to
find a correlation of the form
with coefficients
, and
The temperature variation of the emissivity of MACOR is
unknown, and is assumed to be negligible in the current
tests. However, measurements of the emissivity temperature
variation ofsimilar ceramic materials such as fused silica
glass (Clayton 1962) suggest that the emissivity variation
of such materials is very small up to temperatures of the
order of
530 K
The sting adaptors are manufactured from
, a 3D
printed simulated polypropylene with a high thermal defor-
mation temperature (Rigur Polyjet 2016).
2.4 Test conditions
The conditions achievable in HSST (Table1) are much less
energetic than real re-entry flows, which generally have
enthalpies on the order of tens of MJ/kg and Mach num-
bers on the order of 25–30 (for re-entry from LEO). Despite
the divergence in total energy between the present tests and
flight conditions, the higher density and lower velocity in
HSST means that the Reynolds numbers achieved in the
wind tunnel are still representative of re-entry Reynolds
numbers, which are typically on the order of
at 80 km.
Furthermore, due to the exponential increase in the density
of the Earth’s atmosphere during the initial stages of re-
entry, the Reynolds number of re-entry flows can increase
rapidly at altitudes around 80 km while the Mach number
only varies weakly.
For these two reasons, the experimental flow conditions
are specifically chosen to investigate the effect of Reynolds
cos (𝜃)
Fig. 2 Experimental set-up in the HSST working section
300 350 400 450 500
Fig. 3 Temperature variation of MACOR thermal properties given by
Corning Inc. and as reported in Imbriale (2013)
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Experiments in Fluids (2020) 61:151
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number on the flow field and surface heat fluxes, while main-
taining a constant Mach number. The tested flow conditions
are presented in Table2. IRT measurements are taken at four
conditions at Mach 5 with nominal
T0=800 K
and nominal
varying from 200 to
800 kPa
. Further schlieren images
are taken at a fifth condition with nominal
T0=350 K
. Due
to the fact that hypersonic flow field behaviour is relatively
weakly dependent on Mach number (known as Mach num-
ber independence), these experimental conditions can still
provide valuable insight into re-entry flow behaviour.
3 Infrared measurements anddata
3.1 Infrared camera andcalibration
The calibration of the infrared camera follows the basic
principles described in Carlomagno and Cardone (2010).
The infrared camera used is a FLIR A655SC fitted with a
FOV lens. The detector resolution is
640 ×480
and the frame rate is
50 Hz
. The IR calibration is performed
using a Fluke 9132 portable infrared calibrator. The cali-
brator consists of a quasi-black-body target with
which can be heated up to 500 °C in 0.1 °C increments. The
calibration is performed insitu, that is with the calibrator
placed in the tunnel working section with the camera view-
ing the calibrator through the germanium window. In this
case, the total radiant intensity detected by the camera
can be written as:
are the transmissivities of the atmosphere
and germanium window,
, and
are the radiant
intensities corresponding to a black body at the tempera-
tures of the target, the atmosphere and the optical window
respectively. The emissivities of the target and the window
are denoted
respectively. By substituting Planck’s
law into Eq.2, and assuming that the absorptivity of the
D=𝜏opt𝜏atm 𝜀I
bb +𝜏opt𝜏atm (1𝜀)I
atmosphere is negligible, that is that
𝜏atm =1
, the following
expression is obtained:
, and
refer to the temperatures of the
object, the ambient environment, and the window, respec-
tively, and R, B, and F are coefficients of radiation. Assum-
ing that
is constant with temperature, then this coefficient
multiplies into the calibration coefficient R. Furthermore, if
are constant (which is reasonable for the test
facility in question) then the last term of Eq.3 becomes a
constant C which must be found during the calibration. The
calibration equation, therefore, becomes:
The addition of a constant C to the calibration equation was
proposed by Zaccara etal. (2019) as a way of taking the
camera Non-Uniformity Correction (NUC) into account and
regulate the different gains and zero offsets of each pixelof
the Focal Plane Array. In our case, it simply represents and
corrects for any emission of the germanium window. We are
able to take the camera NUC into account by calibrating the
camera to a NUC-corrected intensity value called the Object
Signal, which is calculated by the FLIR A655SC’s firmware.
During the calibration, the signal of the calibration tar-
get is recorded at 55 evenly spaced data points between
676.5 K
. The Levenberg-Marquardt nonlinear
least squares algorithm is used to calculate the calibra-
tion coefficients R, B, F, and C in Eq.4. The resulting
eBTobj F+𝜏opt(1𝜀)
Table 2 Experimental flow conditions
Reynolds numbers are calculated using the cube length
L=30 mm
Case no. M
Re ×
, kPa
1 5 40.0 208 782 Y
2 5 79.5 424 796 Y
3 5 109 620 831 Y
4 5 148 835 825 Y
5 5 549 810 348 N
300 400 500600 700
1000 Calibration Values
Curve Fit
Fig. 4 Camera calibration curve
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calibration curve is shown in Fig.4, while the calibration
parameters, including the coefficient of determination and
the RMS error are presented in Table3.
The camera is mounted to a Minitec frame fixed to the
floor of the laboratory. It is positioned at a
angle to
the horizontal axis of the model (Fig.2), which allows it
to image three sides of the cube model.
3.2 Image processing
This section describes the image processing algorithm used
to convert the raw IR data acquired by the camera (Fig.5a)
to temperature values suitable for input to the heat flux
calculation. First, the raw data are filtered using a three-
dimensional Savitsky-Golay filter in both space and time.
Following filtering, the IR video is stabilised to remove the
effect of model and sting vibration during tunnel start-up and
shut-down. The image registration algorithm used to stabi-
lise the IR video is the single-step discrete Fourier transform
approach proposed by Guizar-Sicairos etal. (2008), which
has already been used successfully on IR videos (Avallone
etal. 2015). This algorithm calculates the displacement
Table 3 Summary of the
infrared camera calibration Parameter Value
Coefficient of determi-
RMS error 0.1
Fig. 5 Image processing steps for the cube IR data
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between two images to a sub-pixel accuracy by computing
the upsampled cross-correlation between an image and a
reference image by the fast Fourier transform.
Following image registration, the locations of the differ-
ent surfaces of the cube in the IR image must be identified,
and an affine transformation calculated to transform the
perspective view of each of the faces to the square arrays
corresponding to the mesh used in the heat flux calcula-
tion method described in Sect.4. Previously, Cardone etal.
(2012) produced a mapping between a surface mesh and
an IR image by means of an optical calibration of the IR
camera. Due to the simplicity of the geometry considered
in this case, we take a much simpler approach. The edges
of the cube in the IR image (Fig.5b) are identified with a
fuzzy-logic-based edge detection algorithm. Following this,
the most likely location of the cube edges are extracted, and
their intersections are used to define the corners of the cube.
These corner locations are used to define the moving points
of an affine transformation to a square (Fig.5c). During
the calculation of the affine transformation, the IR image is
down-sampled to give a final spatial resolution of the cube
temperature maps of
2 pix/mm
(Fig.5d). The sensitivity of
the heat flux calculation to errors in both the edge detection
algorithm as well as the down-sampling procedure are dis-
cussed in Sect.6.5.
Once the affine transformation has been calculated for
every IR frame, the temperature map on each surface of the
cube can be calculated by applying the infrared calibration
(Eq.4) to each surface. By applying the calibration indepen-
dently to each surface, the variation in emissivity of each
surface (due to the directional emissivity effect) can be taken
into account. To calculate the directional emissivity for each
cube face, the viewing angle
to each surface is calculated
using knowledge of the camera viewing direction
(this is
known from the camera’s orientation in its mounting posi-
tion), and thenormal vector to each of the cube’s faces
(which is known from the cube’s orientation on the sting):
The sensitivity of the computed heat flux values to errors in
the model emissivity, both due to surface finish as well as
directional effects, is discussed in Sect.5.
4 Hypersonic heat ux calculation
Once the temperature history of an object has been meas-
ured, the heat flux to the surface of the object can be cal-
culated by a physical model—the heat conduction equa-
tion—of the heat transfer in the measurement area. Walker
and Scott (1998) identified three different classes of such
1. Analytical techniques
2. Direct numerical techniques
3. Inverse techniques.
The first of Walker and Scott’s three solution classes uses a
theoretical closed form solution to the 1D form of the heat
equation, such as that proposed in Cook and Felderman
(1966) and Kendall and Dixon (1996). If the boundary flux
is piecewise constant and the body material properties are
constant, this is a very simple and quick way of calculating
the heat flux to a body. However, the restrictions on the form
of the boundary conditions, as well as the requirement for
the material properties to be constant are both strong limita-
tions. The second of Walker and Scott’s techniques addresses
the first’s drawbacks by numerically solving the heat conduc-
tion equation using numerical techniques. These could be
a finite difference, finite volume, or finite element method
with implicit or explicit time-stepping. The experimental
temperature measurements are given as Dirichlet boundary
conditions. This allows more flexibility in the solution, as
it permits 2D and 3D conduction to be taken into account,
as well as variable material thermal properties, such as
described in Häberle and Gülhan (2007) and Henckels and
Gruhn (2004). The primary drawback of this technique is
that it involves differentiating the experimental data, thus
magnifying any experimental error or noise.
The third, most sophisticated and robust method is to
solve an inverse heat condution problem (IHCP). Typically,
inverse problems involve the calculation of an object’s
boundary conditions using knowledge of some internal
conditions (Ozisik and Orlande 2000). In the context of
measuring heat fluxes using IRT, the heat flux to the sur-
face is estimated by considering the evolution of the surface
temperature. IHCPs, while offering maximum flexibility
in heat flux calculations (Avallone etal. 2015) also have
their disadvantages, namely their significant complexity and
computational cost. In addition to this, they are ill-posed
problems as their solutions are not unique, meaning their
solutions are extremely sensitive to small changes in input
data (Ozisik and Orlande 2000).
Due to the fact that hypersonic heat fluxes tend to be
extremely high, a common assumption made when calculating
heat fluxes with any of the above techniques is that any trans-
verse heat transfer within the body is small compared to the
convective heat flux to the body. In this case, it is reasonable to
neglect any transverse conduction and only consider the heat
flux normal to the body surface. This assumption gains further
justification when heat flux is being measured to a model with
a low thermal diffusivity. However, in cases where significant
transverse conduction is present, the dimensionality of the heat
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conduction equation must be increased in order to achieve an
accurate solution. Previous applications of a 2D-IHCP in the
context of hypersonic aerothermodynamics include Avallone
etal. (2015) and Zaccara etal. (2019), where a 2D-IHCP solu-
tion was used to calculate the heat transfer caused by turbu-
lent transition on a wedge and cone respectively. Both results
showed very limited dependence on transverse conduction.
Other authors (Nortershauser and Millan 2000; Sousa etal.
2012) have performed 3D-IHCP solutions on problems con-
sidering the heating of small test articles by flames and electric
heaters, but they only considered relatively small computa-
tional domains (approximately 300 cells). In the case of the
present work, any assumptions of 1D or 2D conduction cannot
be made—the conduction in the model near the corners and
edges will be strongly two or three-dimensional and therefore
the 3D-IHCP solution must be solved.
In the remainder of this section, we briefly describe a
method of solving the IHCP on the cuboidal domain by the
conjugate gradient method with adjoint and sensitivity prob-
lems, a commonly used IHCP solution methodology (Huang
and Wang 1999; Imbriale 2013). In the summary below, we
follow the derivation of Ozisik and Orlande (2000).
4.1 Denition ofthedirect problem
We start by introducing the direct problem. Consider a
cuboid domain
with surfaces
S=S1, ..., S6
. The heat equa-
tion on this domain, with Neumann boundary conditions can
be written as:
, and k are the material density, specific heat
capacity, and thermal conductivity respectively. The material
temperature is T, q is the conductive heat flux, and
and t
are the space and time variables.
4.2 Denition oftheinverse problem
The inverse problem can be described as follows: find the
value of q(S,t) which gives the known temperature evolu-
tion at each measurement point
on S (the details of the
computational mesh are given in Sect.4.6). In practise, this
is every pixel of the IR image. Start by defining the cost
functional for this problem:
is the solution to 6 for q(S,t) at each measurement
point, and M is the total number of measurement points. The
𝜕n=q(𝐱,t),𝐱S,0 <t<tf
f(q(S,t)) = tf
Tm(q(S,t)) − Ym
conjugate gradient method (CGM) attempts to iteratively
construct a value of q(S,t) of the form:
where the subscript n denotes the iteration count.
4.3 Calculation ofthesearch step size
andthesensitivity problem
The step size
is taken to be the step size by which the cost
reaches a minimum in the direction
. The
expression for
can therefore be found by minimisingEq.7,
is defined as the directional derivative of the tem-
perature T in the direction of q. To find an expression for the
evolution of
, it is assumed that a perturbation of
in Eq.6 causes a perturbation
in the solution. Sub-
stituting these in the direct problem, subtracting the original
direct problem from the resulting equations and ignoring 2nd
order terms yields the sensitivity problem:
4.4 Calculation oftheconjugate direction
andtheadjoint problem
The conjugate direction,
can be calculated by the equation
is the gradient of the cost functional and
called the conjugation coefficient, calculatedusing the
Fletcher-Reeves formula:
The only unknown quantity in Eqs.11 and 12 is the gradient
of the cost functional,
. The expression for the evolution
is known as the adjoint equation:
𝛥T(𝐱,t)=0, 𝐱𝛺,t=0
𝜕n=𝛥q(𝐱,t),𝐱S,0 <t<tf
𝜕t= −∇(k(T)∇𝜆),𝐱𝛺
𝜆(𝐱,t)=0, 𝐱𝛺,t=tf
𝜕n=2(TY),𝐱S,0 <t<tf
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. The derivation of the adjoint problem follows
a process similar to the one used to derive the sensitivity
problem. The details can be found in Ozisik and Orlande
Of note is the fact that, due to the nature of the final value
problem (Eq.13), the gradient of the cost functional at the
final time is zero, that is,
and, therefore,
the conjugate direction will always be zero at the final time.
To overcome this singularity, the gradient at the final time
is modified as follows:
is the time step used to solve the heat equation. In
this way, the effect of the singularity is reduced. To further
reduce the effect of the singularity, the IHCP solution cal-
culation is extended beyond the end of the tunnel run by 50
time steps (during this time the heat flux to the cube is zero).
4.5 Stopping criterion
The stopping criterion of the CGM can be defined either by
a tolerance criterion, or when the algorithm reaches a mini-
mum, that is, when there is negligible change in the solu-
tion after a direction re-set. In other words, if there are no
measurement errors, the stopping criterion can be defined as
𝜀 << 1
. In practice, the temperature measurement
error will place a constraint on how small
can become.
Following Huang and Wang (1999)and Ozisik and Orlande
(2000), the temperature measurement residuals will be
approximately equal to the standard deviation of the tem-
perature measurement errors, that is:
in Eq.15 can be expressed by:
which gives the appropriate value of
for the current
4.6 Algorithm
In summary, the CGM for solving the IHCP can be described
in the following steps at each iteration n:
1. Solve Eq.6 with
as the boundary conditions.
2. Check the stopping criterion (Eq.15). If satisfied, exit
the solution.
3. Solve the adjoint problem (Eq.13) to obtain
4. Calculate the conjugation coefficient
by the Fletcher-
Reeves formula (Eq.12).
5. Calculate the conjugate direction
6. Solve the sensitivity problem (Eq.10) with
7. Calculate the step length
8. Update the solution to obtain
Previous applications of the CGM (such as Imbriale 2013)
to hypersonic problems did not take into account the tem-
perature variation of the material thermal properties k(T)
. To do this, we calculate the values of these
properties at every spatial point and time step during the
solution of the direct problem (step 1 in Sect.4.6). These
values are then used at the corresponding locations and
times in the solution of the adjoint and sensitivity prob-
lem on the same domain. In this sense, k and
in Eqs.10
and 13 could be written as functions of space
and time t
rather than temperature T. These values are then updated
at each iteration during the solution of the direct problem.
For this work, the algorithm has been implemented
in Matlab. The direct, sensitivity, and adjoint problems
are solved with the same forward-time central-space
(FTCS) finite differencing scheme on an equally spaced,
structured grid with
dx =0.5 mm
. To reduce the compu-
tational expense of the IHCP, the flow around the cube
is considered symmetrical and the IHCP is only solved
over one quarter of the cube, that is
0.5 zL0.5
(see Fig.5d for coordinates), giving a final
grid size of
54 ×103
points. The boundary conditions at
the three surfaces of the domain where temperatures are
unknown are considered to be adiabatic. This is justified
as the transverse conduction normal to the boundaries in
these regions is likely to be small.
Once the conductive heat flux q has been evaluated,
the modified Stanton number
is calculated, defined as
is the total radiative heat flux away from a surface,
given by the Stefan-Boltzmann equation,
is the convec-
tive heat flux to the surface, and
is the total enthalpy of
the free-stream flow, given by:
In the remainder of this work, any reference to Stanton num-
ber refers to the modified Stanton number as defined above.
The wall and free-stream enthalpy values
are cal-
culated using the HOT thermal database package for Matlab
and Octave (Martin 2019).
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5 Error sensitivity analysis
To validate the IHCP approach to calculating heat fluxes
outlined in Sect.4, as well as to estimate the errors asso-
ciated with the calculations, it is important to investigate
both the uncertainty inherent to the IHCP solution, as well
as the sensitivity of the solution methodology to errors in
the input data. To conduct such an analysis, it is necessary
to use a synthetic (rather than experimental) dataset. In
this way the precise difference between a true
(which is known for the synthetic dataset) and the IHCP-
calculated value can be found.
For the sensitivity analysis in this study, the synthetic
heat fluxes applied to the cube geometry are chosen such
that they approximate the expected experimental values.
On the stagnation surface of the cube, the heat fluxes are
given by a modified version of the expression given by
Klett (1964) for the heat flux to a flat-ended cylinder:
where L is the cube length, s and z are the coordinates shown
in Fig.5d, and
. The conductive heat flux value at
the stagnation point is chosen as
On the streamwise surfaces of the cube, the heat flux
distribution is given by the Eckert reference temperature
method for a flat plate hypersonic boundary layer. To
include temporal variation in the sample data, the heat
flux values are multiplied by a factor of
Once the temperature maps on each surface have been
calculated by a solution of the direct problem (Eq.6), the
temperatures are passed backwards through the IR cali-
bration to obtain equivalent Object Signal values. This
makes it possible to investigate how various errors (such
as in emissivity and ambient temperature measurement)
propagate through the calibration procedure.
When calculating the Stanton number values, the syn-
thetic values of
, and
are found from typical
HSST total temperature and pressure data for a run at
nominal conditions of
T0=800 K
P0=800 kPa
Errors are assumed to enter the Stanton number calcula-
tion from the following sources:
1. Error inherent in the IHCP solver.
2. Error in the ambient temperature measurement.
3. Error in the camera calibration.
4. Error in assumed emissivity values.
5. Error in the initial temperature distribution over the
6. Error in the value of thermal conductivity.
7. Error in the value of specific heat capacity.
8. Error in measurement of the wind tunnel total tempera-
9. Error in measurement of the wind tunnel total pressure.
To investigate the sensitivity of the data processing proce-
dure to each of these error sources, the inputs to the IHCP are
perturbed by an amount approximately equal to the estimated
measurement error
. The resulting error in the Stanton num-
is then characterised using the normalized root mean
squared error:
is the true value of the Stanton number, given
by the synthetic data, and
are the time indices
where a steady-state Stanton number is achieved. A sum-
mary of the error sources, their assumed error (and justi-
fication), and the resulting values of
can be found in
Traditionally, the global accuracy of the calculated Stanton
numbers could be estimated as the root sum of the squares of
values. However, by this method, the error associ-
ated with the IHCP methodology will also be present in each
approximation of the partial derivatives, and the total error will
contain incidences of the IHCP error. To correct for this, we
write the sum of squares error equationas
M(tend tstart )
m=1(CH(𝐱m,t)− ̂
Table 4 Error sources and sensitivity
Error source Variation
IHCP solution 1.5
Ambient temperature measurement
Tamb [K]
1.7% 1.2
Infrared calibration
K 1.2
Model initial temperature distribution
2.7% 1.2
Material emissivity
8.0% 8.3
Material thermal conductivity
k[J/kg K]
6.0% 1.2
Material specific heat capacity
cp[J/kg K]
2.0% 1.2
Free-stream total temperature
2.5% 7.2
Free-stream total pressure
4.2% 4.4
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so that only one incidence of
is included in the final
error value. This gives a global error in
of 12%, which is
dominated by the errors due to the material emissivity (see
Table4). It is important to note, however, that this value is
not constant across the entire computational domain. Fig-
ure6 shows how the error value changes across the com-
putational domain. In this case, we define a time average
For the flat faces of the model, the error is approximately
9%, below the global error of 12%, but in regions of strong
multi-dimensional conduction the error can increase up to
15%. It is notable that the error is much higher near the
edges and corners of the cube, where three-dimensional con-
duction is strongest.
The effect of the internal conduction and the magnitude
of the error in the regions where the conduction is strongest
could bedirectly experimentally assessed by placing ther-
mocouples internally in the model. Unfortunately, the effect
of installing such instrumentation would significantly com-
plicate the infrared measurements and the IHCP solution.
Due to the extremely low thermal diffusivity of MACOR,
the thermocouples would have to be placed very near the
surface of the cube. The placement of these instruments
(tend tstart )
(CH(𝐱m,t)− ̂
would affect the distribution of the cube material as well
as the surface temperature of the cube. These factors would
have to be corrected for in the solution of the IHCP which
would require a much more complex mesh.
6 Results
6.1 Heat ux measurement validation
withahemisphere model
Taking IRT measurements of temperature histories and cal-
culating heat fluxes with an IHCP solution are both complex
processes with many possible sources of error in both exper-
imental set-up (including IR calibration) and data reduction
algorithms. Therefore, to validate the combined experimen-
tal set-up and the IHCP solver, they are used to calculate
the experimental stagnation point heat flux to a hemisphere.
This is a particularly useful geometry with which to validate
the experimental set-up as the stagnation point heat flux on
a hemisphere is well characterised, and can be accurately
calculated using a number of different methods.
For this application, we obtain a theoretical value of the
heat flux at the stagnation point of a sphere by solving the
self similar form of the boundary layer equations (for their
derivation see Anderson (2006)):
, and
. In these relations,
are the Lees-Dor-
onitsyn variables, subscripts
refer to flow variables at
the edge of the boundary-layer, and Pr is the flow Prandtl
number. The reader is referred to Anderson (2006) for more
details about these equations as well their derivations. The
equations are solved using the tridiagonal solution method
described in Blottner (1979) and implemented in Adams
(2002). This is preferred to using an existing correlation
(such as the Fay and Riddell (1958) or Sutton and Graves
(1971) correlations) due to the fact that the free-stream con-
ditions used for these experiments lie outside of the range of
validity of these correlations.
These theoretical heat flux values are compared to experi-
mental values at the stagnation point of a
hemisphere in a hypersonic flow with
T0=805 K
P0=200 kPa
(which corresponds to a free-stream
Reynolds number of
Rem=1.23 ×106
). The experimen-
tal heat flux values are calculated using both the IHCP solu-
tion described above, as well as the Cook and Felderman
(1966) equation, a Class 1 method using the classification
Cf ��
+�� =
Pr g
Fig. 6 Contours of the error in Stanton number
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system described in Sect.4. The IHCP is solved assuming
1D conduction only (i.e. neglecting any transverse con-
duction in the hemisphere model) at the stagnation point.
A plot of experimental Stanton numbers at the stagnation
point of the hemisphere during the run of the HSST is pre-
sented in Fig.7, alongside the theoretical value. The time-
averaged 1D-IHCP value during the wind-tunnel steady
, while the Cook and Fel-
derman value is
, slightly higher likely due
to its assumption of constant material thermal properties.
The boundary-layer self similar solution gives a value of
, 6% smaller than the IHCP solution, and 8%
smaller than the Cook and Felderman value. How much the
difference between the theoretical and experimental heat flux
values is due to the IRT measurement technique and heat
flux calculation method, or due to errors in the free-stream
flow parameter estimation could be ascertained by perform-
ing an additional test with a conventional heat flux probe,
instrumented with a thin-film heat flux gauge. This was not
performed in the present experimental campaign.
The RMS error of the IHCP solution is 1.7%, while
for the Cook and Felderman solution it is 2.2%. These are
slightly lower than the values of 3–4% calculated by Aval-
lone etal. (2013), likely due to the slightly higher noise
𝜎=0.8 K
) in the temperature histories gathered by Aval-
lone etal.
The results described in this section, in combination with
the error sensitivity analysis described in Sect.5 gives fur-
ther confidence in the cube results discussed below.
6.2 Cuboid schlieren results
Schlieren photographs of the cube model at high (
549 ×103
and low (
40.0 ×103
) Reynolds numbers are presented in
Fig.8. The schlieren appears to confirm the appearance of
different flow structures with increasing Reynolds number as
predicted by Rees etal. (2018). Most notable is the appear-
ance of an apparent separation shock at the windward edge
of the cube at high Reynolds numbers (Fig.9) which is not
present at lower Reynolds numbers. This pattern is similar to
that imaged by Matthews and Eaves (1967) around a cylin-
der, and supports the conclusion made in Rees etal. (2018)
that a separation bubble can form on the sides of a cube at
hypersonic speeds, as Reynolds number is increased, even if
it was not present at lower Reynolds values. The free-stream
total temperature was lowered significantly to achieve the
high Reynolds number condition, and therefore it was not
possible to take any high-quality IRT data at this condition.
Cook and Felderman (1966)
Current 1D IHCP
Theoretical value
Fig. 7 Hemisphere stagnation point Stanton number evolution during
a tunnel run
Fig. 8 Schlieren images for a cuboid at
at high and low Reyn-
olds numbers
Fig. 9 Labelled schlieren at
Re =549 ×103
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As a result the separation bubble’s effect on the heat flux
could not be quantified.
6.3 Heat uxes tothecube
The solutions to the 3D-IHCP for the cube model at the
four-different free-stream conditions are plotted in Fig.10.
These are time-averaged Stanton numbers calculated by
averaging overthe steady-state portionof the tunnel run time
s). The trends in the contours are as expected,
with heat flux increasing due to the thinning boundary layer
as the edges and corners are approached. Notably, in addi-
tion to the increases of heat flux at the edges of the cube,
there are also wedge-shaped regions of high
along the
streamwise edges of the cube. These appear to show some
Reynolds number dependence, getting wider as the Reyn-
olds number is increased. Figures11 and 12 show plots of
Stanton number normalised by the stagnation point Stan-
ton number in both the streamwise (at the centreline of the
) and spanwise (at
) directions.
The identification of the stagnation point Stanton number
is discussed in Sect.7.1. Figure11 shows that the wedges
of increased Stanton number along the streamwise edges
of the cube are bounded either side by a region of slightly
decreased Stanton number. Furthermore, Fig.12 shows that
on the side face of the cube, the Stanton number tends to
reach a maximum at
The solutions to the 1D-IHCP (neglecting any transverse
conduction) for the cube model at the four free-stream con-
ditions are presented in Fig.13. Due to the nature of the
1D solution, the 1D-IHCP can be solved across the whole
measurement surface of the cube (no symmetry assumption
is necessary) and so results for the entire cube (rather than
a quarter-cube as in Fig.10) are shown. These results show
broadly similar behaviour to the 3D-IHCP solutions, albeit
with less-defined changes in heat flux near the corners and
edges of the model.
One thing which is notable in the 1D-IHCP contours
that is not immediately obvious in the 3D-IHCP contours
is the presence of regions of lower Stanton number on the
stagnation surface of the cube. These regions correspond
to the low-temperature regions visible in the raw IR image
(Fig.5a). These regions are off-centre of the stagnation sur-
face, and the relative strength of the reduction in heat flux
appears to increase with Reynolds number. These regions
are also present in the 3D-IHCP solution, however, due to
the noisier nature of the 3D solution, theyare not as clear.
6.4 Eects ofthree‑dimensional conduction
The transverse (conductive) heat transfer within the model
is time-dependent. Asdifferent parts of the model heat up
at different rates the temperature gradients (and therefore
internal conduction rates) willvary with time. A compre-
hensive analysis of the effect of internal conductionon the
accuracy of the 1D versus the 3D-IHCP solution would take
this variation into account. However, for the purposes of this
analysis, we will simply compare the time-averaged differ-
ence between the
values obtained with the 3D and 1D
conduction assumptions. We define contours of the differ-
ence due to dimensionality as
which are plotted in Fig.14. The results show that, although
the effect of high-dimensional conduction is unimportant
on the stagnation surface and large parts of the side-faces of
the cube, failing to account for 3D conduction in the regions
near the corners and edges of the cube can result in errors
up to 500%. These errors are likely lower at the start of the
tunnel run time and higher towards the end. Furthermore,
there are regions of significant high-dimensional conduction
around the edges of the hot wedges described above. Despite
these errors associated with the 1D-IHCP, the 3D-IHCP
solutions also have some limitations. First, the 3D results
are noisier than the 1D results. Thisnoise is captured by the
error contoursshown in Fig.6. More importantly, the 3D
solutions rely on an accurate image-to-IHCP input perspec-
tive transformation. If the location of the edge of a cube is
assumed to be even slightly wrong the
calculation can be
significantly affected.
6.5 Sensitivity toimage processing andquality
To investigate the sensitivity of the 3D-IHCP solution to
errors in the identification of theedge locations of the cube,
seven analyses of solution’s sensitivity to the edge location
were performed. For this analysis, the four corners defining
the cube’s stagnation surface (see Fig.5c) were moved up
and down by 3 pixels (in the raw IR image), to give a total
of 6 different edge locations, in addition to the baseline loca-
tion used in the main results. Furthermore, to investigate
the effect of the down-sampling of the IR image, another
analysis was performed wherethe image was processed with
minimal down-sampling, resulting in a solution with a spa-
tial resolution of 3.7 pix/mm rather than 2.0 pix/mm.
The results of these sensitivity analyses on the Case
4 (high Re) results can be seen by examining the cube
centreline Stanton numbers (Fig.15). As the corners get
moved up the IR image (that is, away from the stagnation
surface), there a region of non-physical negative Stanton
number that appears after moving the edge location only
2 pixels. These negative Stanton numbers were reported in
Rees etal. (2019), but were not identified asbeing caused
by errors in the image processing method. As the corners
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get moved down the IR image, that is towards the stagnation
surface, the increase in Stanton number across the cube edge
becomes non-monotonic after a movement of 3 pixels. The
non-physicality of this behaviour suggests that the current
image processing algorithm can locate the edge of the cube
in the IR image to within
pixels. However, Fig.15c
suggests that the spatial resolution of the image, as well
as the down-sampling during the image processing could
alsoaffect the accuracy of the edge location. Although the 2
pix/mm and 3.7 pix/mm lines in Fig.15c look very similar,
Fig. 10 Stanton number contours calculated using a 3D-IHCP solution for a cube at Mach 5 and at four different Reynolds numbers
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we note that the extra noise present in the high-resolution
solution may be causing a non-monotonic change in heat
flux through the cube edge at
. Finally, we note that
the down-sampling does not otherwise affect the 3D-IHCP
solution, suggesting that down-sampling process does not
significantly affect the results.
Obviously, accurate identification of the cube’s edge
locations in the infrared image is a crucial aspect of obtain-
ing accurate heat fluxes using the 3D-IHCP solution. Edge
detection algorithms generally rely on the assumption that
edges are locations of high gradient (even if they may
not place the edge at the region of highest gradient). This
assumption is not true in the case of the windward edge
) of the cube. Therefore, the localisation of this
edge of the cube, as well as the streamwise edge at
relies on the assumption that the edge is located at the region
of highest temperature, which is not justified a priori. This
problem could be mitigated by performing an optical cali-
bration of the camera and then fitting the IHCP cube mesh
mapped to its surface (as discussed in Sect.3.2 and by Car-
done etal. (2012)).
7 Discussion
7.1 Stagnation surface behaviour
The most striking result of the stagnation surface contours in
Figs.10 and 13 is the presence of the regions of lower Stan-
ton number. The variation in the Stanton number of the pre-
sent experimental data due to these ‘cold spots’ is up to 30%.
The cold spot location and strength also appears to show a
Reynolds number dependence, getting stronger and further
off-centre as the tunnel total pressure (and therefore Reyn-
olds number) increases. Although this behaviour perhaps
indicates that the cause of these regions is non-uniformity of
the free-stream flow, previous studies of the flow uniformity
of HSST (Erdem 2011; Fisher 2019) suggest that any non-
uniformity across the test jet is negligible. Alternatively, an
imperfection in the model surface conditions could cause
such a behaviour. However, the tests were performed with
two different cube models, ruling out this explanation.
Instead, we propose a different explanation: that these
regions are in fact the stagnation points of the cube, which
have been moved off-centre by the flex of the tunnel mount-
ing sting. The stagnation point on the windward surface of
the cube should manifest itself in the Stanton number con-
tours as such a cold spot. This is due to the fact that as the
flow accelerates away from the stagnation point on the wind-
ward surface, the boundary-layer will thin, increasing the
heat flux, making the stagnation point the region of lowest
heating on the stagnation surface. If the model is perfectly
aligned in the free-stream flow direction, the stagnation
point would be at the geometric centre of the stagnation
surface. However, due to the planar nature of the stagnation
surface of a cube (in other words, the radius of curvature
is infinite), the location of the stagnation point is likely to
be highly sensitive to the cube attitude—a small change in
the cube attitude may result in a significant change in the
stagnation point location. The flexure of the sting during a
run will change the cube attitude slightly, causing the stag-
nation point to move. As the Reynolds number of the flow,
and therefore the aerodynamic force on the cube increases
with the increase in flow total pressure, the sting flex will
get larger, making the change in stagnation point location
and strength stronger. This off-set stagnation point would
explain why the plots in Fig.12 do not collapse exactly, and
Re = 40.0 10 3
Re = 79.5 10 3
Re = 109 10 3
Re = 148 10 3
Fig. 11 Spanwise Stanton number profiles at
0 0.5 1 1.5
1.5 Re = 40.0 103
Re = 79.5 103
Re = 109 103
Re = 148 103
Fig. 12 Streamwise Stanton number profiles at
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why the Stanton number at the geometric centre of the cube
does not correspond to the stagnation point Stanton number.
To investigate the stagnation point Stanton numbers meas-
ured in these experiments, they are compared to correlations
for the stagnation point heating to a flat surface at similar flow
conditions. Previous studies (Matthews and Eaves 1967; Trim-
mer 1968) have measured the stagnation point heat flux to a
flat-ended cylinder and compared it to the stagnation point
heating to a hemisphere. These studies related the relative
heating rates between these two geometries to the effective
stagnation point velocity gradient at the edge of the boundary
We know from many stagnation point correlations that the
Stanton number at a stagnation point is directly proportional
to the square root of
Assuming that the only difference between an equivalent
stagnation point flow on a round surface and a flat face is
the velocity gradient, then:
is the stagnation point Stanton number to a flat
face and
is the stagnation point Stanton number to an
equivalent spherical geometry. It should therefore be pos-
sible, using knowledge of the difference in velocity gradi-
ent between the two geometries, to find a correspondence
Fig. 13 Stanton number contours calculated using a 1D-IHCP solution for a cube at Mach 5 and at four different Reynolds numbers
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between the stagnation point heat flux to a flat surface and
an equivalent sphere.
Trimmer (1968) and Matthews and Eaves (1967) per-
formed experiments to measure the velocity gradient to
an end-on cylinder in hypersonic flow. Cylinders of vari-
ous bluntness were tested, from a hemispherically-capped
cylinder to a flat-ended cylinder. By comparing the non-
dimensionalised velocity gradients measured in these two
studies, the constant of proportionality in Eq.28 is found
to be 0.54 (from the Matthews and Eaves data) and 0.57
(from the Trimmer data). Klett (1964) found similar results,
suggesting a proportionality constant of 0.5, although did
not report what data this value was based on. The Reynolds
number variation of these results (Klett, Trimmer, and Mat-
thews and Eaves) are presented in Fig.16, as well as the
current experimental data for a cube. The combined data
show significant scatter, although the current experimental
data are generally lower than the earlier data. The average
ratio of the stagnation point Stanton numbers measured
on the cube geometries to an equivalent hemisphere value
) (found using the same self-similar solu-
tion used in Sect.6.1) is found to be 0.44 on average across
conditions. This is not surprising as the cube will naturally
have an effective nose radius slightly larger than L/2 due
to the presence of the corners. In this sense, a better way
of defining the effective nose radius of a cube would be to
use the diagonal distance across the corners:
Using this definition, the average constant of proportionality
for the current experimental data becomes 0.52, closer to the
values reported in the earlier references.
It should be remarked that the previous experimental data
were collected using thin-film heat flux gauges at discrete
locations on the model stagnation surface. Therefore, it
would not have been possible for the authors to capture the
variation in heat flux occurring in the current experimental
data. It is therefore possible that the stagnation point heat
fluxes measured in the earlier data may be larger than the
true stagnation values.
7.2 O‑stagnation behaviour
In the off-stagnation regions, that is the side faces of the
cube, the most notable flow feature are the regions of
high heat flux, or wedges, which appear to emanate from
the windward corners of the cube. Such flow phenom-
ena have already been observed in CFD simulations of
hypersonic flow around faceted shapes, such as in Gülhan
etal. (2016), but have not been studied experimentally.
The angles of these wedges appear to show dependence
on Reynolds number, suggesting that they are a viscous,
rather than inviscid flow effect. This is in contrast to simi-
lar results presented in Rees etal. (2019), which indicated
Fig. 14 Percent errors due to inappropriate conduction assumption
Fig. 15 Centreline Stanton number profiles for the Edge Sensitivity analyses for Case 4
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no Reynolds number dependence, leading to the oppo-
site conclusion. The likely reason for this discrepancy is
that the results in Rees etal. (2019) only considered a
2D-IHCP in the spanwise direction, which resulted in less
spatially accurate wedge shapes. The shapes of the wedges
in the high and low Reynolds number cases are compared
in Fig.17. These profiles were obtained by thresholding
the Stanton number contours on the side faces so that only
regions where the heat flux was higher than
of the
average heat flux on the surface were visible. The edges of
these thresholded images were then extracted to produce
the profile. The wedge spreading angles for all the tested
conditions are reported in Table5. The heat fluxes under
the wedges can be significant, with regions of peak heat-
ing reaching up to
of the stagnation point heat flux
value. The average values of the increased Stanton number
caused by these wedges,
(defined as the average value
of the Stanton number bound by the contours in Fig.17)
are reported in Table5.
Notably, the regions of high heat flux wedges are
bounded on either side by a similar region of lower heat
flux (see Fig.11). Such a pattern is often seen in the pres-
ence of vortical structures, suggesting that the wedges are
generated by vortices being shed by the corner of the cube.
The Reynolds number dependence of the wedge angle is
further evidence that these wedges are a viscous flow
Away from these wedges, the heat flux on the side faces
of the cube is much lower than anywhere else. The aver-
age values of the Stanton number along the centreline of
the cube,
(that is along
) are
reported in Table5, and are between 15 and 18% of the
stagnation point value.
7.3 Comparison tosatellite demise heating models
Object-oriented satellite demise tools such as DRAMA
(Martin etal. 2005) prescribe heat fluxes to simple shapes
such as cuboids and cylinders through the use of a heat-
ingshape factor
, defined as:
is the space-averaged heat flux to the object:
is the stagnation point heat flux to a sphere with
radius equivalent to the effective nose radius of the object
being considered. By convention, the equivalent nose radius
is taken to be
, rather than
as sug-
gested above. In this way, the heat load to an object can be
calculated by finding
, which is trivial to calculate using
any number of correlations [such as Sutton and Graves
(1971) or Fay and Riddell (1958)], and multiplying by the
shape factor (which is stored in a library containing fac-
tors for multiple different primitive shapes at many different
orientations and attitudes). The shape factors for the experi-
mental data are presented in Table5. Although the shape
factors show very little dependence on Reynolds number,
they hide the strong spatial variations in heat flux which
exist in reality. The fact that the highest heat fluxes occur
near the edges of the cube means that these are the regions
sh =
0.4 0.6 0.8 11.2 1.4 1.6
Current experimental data
Klett (1964)
Matthews and Eaves (1967)
Trimmer (1968)
Fig. 16 Comparison of flat-surface stagnation point values for differ-
ent datasets
Fig. 17 Experimental Wedge shapes at high and low Reynolds num-
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Experiments in Fluids (2020) 61:151
1 3
Page 19 of 22 151
which are likely to fail or melt first during destructive re-
entry. As a result, the fragmentation and demise of a satellite
is most likely to be driven by the heat fluxes in these regions.
To take the important spatial variations of heat flux into
account, we propose that shape factors
should instead
be presented as edge-specific shape factors, where the aver-
age in Eq.30 only includes the cube regions near the edges.
There are two ways of doing this. Either all of the edges of
the cube can be included in the average, or the edges can
be separated into three categories: windward (
streamwise (
), and leeward (
, the data
for which are not available from the current experiments),
and then three different heat flux averages
and shape fac-
tors can be calculated. To illustrate this, shape factors for
all these different definitions have been plotted in Fig.18.
In this case, the definition of the ‘edge’ is the region of the
cube surface within 0.1L of the spatial discontinuity. Only
this region is used in the calculation of Eq.30.This defi-
nition is slightly arbitrary and the factor of 0.1 could be
increased or decreased depending on the context. For exam-
ple, if it is known that the epoxy joints holding the satellite
panel together have a length of
would be the most
appropriate factor to use for edge definition.
The above results are presented with an implicit assump-
tion that the current methodologies used for modelling
satellite demise (broadly outlined in the Introduction) are
the most appropriate way of modelling demise. We implic-
itly assume that fragmentation is driven by melting or other
catastrophic failure in regions of maximum heat flux. While
this is a seemingly intuitive model, any model of satellite
demise is extremely challenging to validate experimentally.
Due to the complexity of the demise process, it is possible
that demise is driven by other failure modes, for example
delamination of the aluminium honeycomb sandwich pan-
els. Alternatively, studies of Thermal Protection System
(TPS) materials have looked at possible TPS failure mecha-
nisms which are initiated by damage to the TPS caused by
micrometeoroid impacts (Agrawal etal. 2013). Such micro-
damages could very well be a nucleation point for satellite
demise and failure. If it can be shown that satellite frag-
mentation is driven by a phenomenon other than melting at
corners and edges then the results presented in the current
study must be re-interpreted with that in mind.
Finally, as briefly discussed in Sect.2.4, these results are
based on a free-stream flow with a fraction of the enthalpy
of a real re-entry flow. To further confirm the applicability of
these results to real re-entry flows, the current results should
be compared against and supplemented with additional data
from CFD, high energy (shock tube) experiments, or even
flight data. The higher free-stream enthalpies considered in
these datasets will cause much higher absolute values of
heat flux, and possibly different heating patterns which could
affect the values of the different shape factors calculated
8 Conclusions
The Mach 5 flow-field around a cube has been studied exper-
imentally, using both schlieren photography to visualize the
flow-field as well as infrared thermography and an IHCP
data reduction method to measure the surface heat fluxes to
the cube. The schlieren images revealed the presence of a
separation bubble on the side surface of the cube at certain
Reynolds numbers. This structure was imaged at a Reynolds
number of
549 ×103
. Unfortunately, due to the low free-
stream total temperature required to achieve this Reynolds
number, it was not possible to take any high-quality IRT
data, and therefore the separation bubble’s effect on the heat
flux could not be quantified.
Table 5 Summary of Stanton
number patterns Case no. Re
Wedge angle
1 40.0 11 1.7 4.4
2 79.5 8.1 1.4 3.5
3 109 6.4 1.2 2.9
4 148 5.7 1.0 2.5
0.4 0.6 0.8 1 1.2 1.4 1.6
10 5
0.5 All Edges
Windward Edge (s/L = 0.5)
Streamwise Edge (z/L = 0.0)
Whole Cube
Fig. 18 The evolution of different shape factors with Reynolds num-
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Experiments in Fluids (2020) 61:151
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151 Page 20 of 22
A detailed error sensitivity analysis of the inverse heat
conduction method for heat flux calculation used showed
that the Stanton numbers calculated by the data reduction
method have an error of 12%, which is dominated by errors
in the material emissivity. These errors appear to be highest
near the regions of the cube where internal transverse con-
duction in the cube is strongest, such as corners and edges.
On the stagnation surface of the cube, the heat flux meas-
urement results show broad agreement with existing data
for stagnation point heating to flat axisymmetric surfaces.
However, experimental heat flux contours to the stagna-
tion surface of the cube shows the presence of distinct off-
centred regions of lower heat flux, which we have termed
‘cold spots’. The Stanton number in these regions can be as
much as 30% lower than the Stanton number to the geomet-
ric centre of the cube. Furthermore, their strength appears
to show some dependence on Reynolds number. We deduce
that the cause of these cold spots is due to sting flex during
the tunnel run, causing a misalignment of the stagnation
point. To confirm this hypothesis, a valuable future study
would be to perform an oil flow visualisation of the flow
around the cube. Such an experiment would also confirm the
presence of a separation bubble on the side face of the cube
at higher Reynolds numbers. We propose that the stagna-
tion point heat flux to a cube can be estimated by calculat-
ing the stagnation point heat flux to a sphere with a nose
radius of
, and multiplying the resulting value
by 0.52, a coefficient very similar to those used to estimate
the heat flux to a flat-ended cylinder. However, the stagna-
tion point heat flux may be as much as 30% lower compared
to other regions on the stagnation surface. The most notable
flow feature on the off-stagnation (side) faces of the cube
are wedge-shaped regions of increased heat flux emanating
from the windward corners of the cube. The heat flux under
these wedges can be very high, with regions of peak heating
reaching stagnation point values. The spreading angle of
these wedges show Reynolds number dependence and we
therefore attribute them to the presence of vortical struc-
tures that are shed from the corners of the cube. Using the
experimentally measured heat fluxes to the surface of the
cube, we calculated different shape-factors describing the
average heating to the cube as a whole as well as edges and
corners, where heating is highest.
Finally, we draw attention to what we believe to be
the two most important weaknesses of the current study.
Firstly, as discussed extensively in Sect.2.4 the flow con-
ditions considered in this study are very low-energy when
compared to true flight conditions. Taking equivalent
experimental infrared measurements at flows correspond-
ing to flight conditions would be challenging, as the infra-
red data capturing would need to take into the account
the transmissivity of the reacting shock layer, the strongly
varying material properties (including emissivity), and the
very short test times. Preliminary arcjet and plasmatron
studies of the re-entry flows around CubeSats (Masutti
etal. 2018) have identified regions of high heating and
temperature using uncalibrated IR measurements. Rep-
etitions of these high-enthalpy studies using carefully
calibrated IR measurements would provide valuable data
with which the current measurements could be extrapo-
lated to flight conditions. The second weakness of the cur-
rent study is that it only considers one orientation of the
model with respect to the free-stream. In reality, a satellite
during re-entry will be tumbling rather than maintaining
one attitude, constantly changing the heat flux distribution
over the geometry. The time scale of the satellite tumbling
motion is much larger than the time-scale of the hyper-
sonic flow. As a result, when calculating heat fluxes for a
tumbling geometry, only steady state flow-fields need to be
considered. Even with this simplification, it is impractical
to gather experimental data at a sufficiently fine-grained
range of attitude orientations. Therefore, future research
whichconsiders the effect of tumbling on the heat fluxes
to a satellite would have to be largely CFD based, with
judiciously chosen conditions and orientations at which
validation experiments can be performed.
Acknowledgements TWR would like to acknowledge the financial
and technical support of the European Space Agency (under NPI 480-
2015), Fluid Gravity Engineering, and the Engineering and Physical
Sciences Research Council through the Imperial College Centre for
Doctoral Training in Fluid Dynamics Across Scales(EP/L016230/1).
This work was further supported financially by the award of a UK
Fluids Network Short Research Visit grant. The authors would also
like to thank the technicians at the Imperial College Department of
Aeronautics for manufacturing the models used in these experiments.
Compliance with ethical standards
Conflict of interest The authors declare that they have no conflicts of
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ThomasW.Rees1 · TomB.Fisher2· PaulJ.K.Bruce1· JimA.Merrield3· MarkK.Quinn2
* Thomas W. Rees
1 Imperial College London, Exhibition Road,
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2 University ofManchester, Oxford Road,
ManchesterM139PL, UK
3 Fluid Gravity Engineering Ltd., 1 West Street,
EmsworthPO107DX, UK
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... For satellites in Geostationary Earth Orbit (GEO), their high altitudes mean the most cost-effective disposal method is to raise their orbits to a so-called 'graveyard orbit', which places them well away from other commonly used orbits. For satellites in Low 15 Earth Orbit (LEO), such re-orbiting manoeuvres are far too expensive in terms of propellant use and the associated costs of increased satellite mass, and alternative disposal methods are necessary. This usually means de-orbiting the satellite so that it re-enters the Earth's atmosphere. ...
... The heating rates to cuboids were first studied in the reports of Crosby & Knox [13], and 60 Laganelli [14], who experimentally measured heat fluxes to a cube in a Mach 8 flow at discrete locations using thin-foil calorimeters. More recently, measurements of the hypersonic heat fluxes to a cube were made using InfraRed Thermography (IRT) [15] so that the heat fluxes across the entire surface of the cube could be measured, rather than just at discrete points on the geometry. These IRT measurements identified regions of increased heating along the streamwise edges of the cube, which were attributed to the presence of vortical structures emanating from the cube's sharp corners. ...
... Using both numerical and experimental methods, this work will investigate the effect of small angles of incidence on the hypersonic flow-field around a cuboid. The numerical simulations will provide a method to visualize the 'hot wedges' identified in [15] which significantly increase the heat fluxes to certain 100 regions of the cube when it is at zero incidence, as well as a method to investigate the effect of incidence on these flow structures. The experimental flows will supplement the numerical results by providing valuable validation data. ...
Full-text available
In order to improve predictions of the on-ground casualty risk associated with the uncontrolled atmospheric re-entry of satellites from Low Earth Orbit, there is significant research interest in the development of engineering models of hypersonic heating rates to faceted shapes. A key part of developing such models is generating accurate datasets of the heat fluxes experienced by faceted shapes at various orientations in hypersonic flows. In this work, we use wind tunnel experiments and CFD simulations to study the hypersonic flow around a cuboid geometry at incidence in a Mach 5 flow at Reynolds numbers of 79.5e3, 109e3 and 148e3. The wind tunnel data are obtained in the University of Manchester's High SuperSonic Tunnel and consist of schlieren images and temperature histories collected using infrared thermography. These temperature histories are used to calculate experimental heat fluxes by solving a three-dimensional inverse heat conduction problem. CFD simulations around the same geometry at equivalent free-stream conditions are calculated with the DLR-TAU code. The experimental and CFD results show good agreement both in terms of heat fluxes as well as flow structure. Notable flow structures include wedge-shaped regions of high heat flux which emanate from the windward corners of the cube. Analysis of numerical Q-criterion contours show that these high heat flux regions are caused by vortex structures generated by the expansion at the cube corner. Analysis of the numerical skin friction coefficient shows that even at incidence there is no breakaway separation from the expansion edges of the cube and the flow remains attached throughout. We show that although there is little change in the average heat flux experienced by a cube at 5 deg incidence to the free-stream compared to one at 0 deg incidence, there are significant changes in the heat flux contours over the cubes at these two incidences. Finally, we calculate a number of heating shape factors which can easily be implemented in satellite re-entry and demise prediction analysis tools.
... Comparisons of experimental and numerical Stanton number along the geometrical centrelines of a cube at 5 incidence, in a Mach 5 flow and at a range of Reynolds numbers. Note that at each Reynolds number condition, the experimental data combines results from two different experiments[16]. ...
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Redevelopment of 2D forebody Pathfinder aerothermal study conducted in thermal non-equilibrium flow field with fully catalytic walls employed to promote maximum surface catalysis. Three flow configurations where convective laminar and turbulent heating profiles investigated. Isothermal cold wall condition employed for all flow configuration and radiative equilibrium wall condition for peak turbulent convective heating run6596* only. Chemical kinetics model assessment shown good agreement of CO2 dissociation and recombination of secondary species in the mixture. Shock wave topology propagates with varying flow conditions where hot shock layer moves closer to stagnation point in thermal non-equilibrium setting promoting large peak transitional-rotational and vibrational modes. This redevelopment study is taken in place in hopes of covering gaps where evolution of dissociation and recombination is obtained over set of flow conditions, temperature modes investigated and heat fluxes for varying wall conditions explored and validated with existing work. laying ground works of hypersonic aerothermal analysis for faceted heating profiles for Hypersonic foldable Aeroshell for THermal protection using ORigami (HATHOR) concept. Strong agreement in peak convective heating where heat flux of laminar and turbulent was 127.23 W/cm 2 and 128.4 W/cm 2 respectively which is within 0.5% percentage error of viscous shear layer VSL code employed for similar peak heating condition. Thorough validation process has been carried out to ensure accuracy is met with no large sources of numerical issues and errors. Discrepancies were with in margins of 10-20% which is plausible, greatest source of errors come from convective turbulent heating profile which presents in errors high as 66%, at peak heating the error estimation for peak heating is reduced to 24.4% which is promising given chaotic nature of turbulence flows.
... Aerodynamics of simple-shaped bodies in high-speed flows is generating renewed interest in the field of atmospheric entry of meteoroids, space debris and separating components of launch systems. Especially in the past decade, the characterization of flowfields around single (Lee et al. 2017;Seltner et al. 2019;Rees et al. 2020;Grossir et al. 2020) and multiple bodies (Laurence et al. 2012;Marwege et al. 2018;Park and Park 2020;Register et al. 2020) has gained importance with respect to fragmentation, demise and separation behaviors of these space objects undergoing atmospheric entry. The determination of their flight trajectories helps to predict the impact area of uncontrolled entering spacecrafts with the objective to protect persons and properties from harm. ...
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The influence of the flight attitude on aerodynamic coefficients and static stability of cylindrical bodies in hypersonic flows is of interest in understanding the re/entry of space debris, meteoroid fragments, launch-vehicle stages and other rotating objects. Experiments were therefore carried out in the hypersonic wind tunnel H2K at the German Aerospace Center (DLR) in Cologne. A free-flight technique was employed in H2K, which enables a continuous rotation of the cylinder without any sting interferences in a broad angular range from 0 $$^{\circ }$$ ∘ to 90 $$^{\circ }$$ ∘ . A high-speed stereo-tracking technique measured the model motion during free-flight and high-speed schlieren provided documentation of the flow topology. Aerodynamic coefficients were determined in careful post-processing, based on the measured 6-degrees-of-freedom (6DoF) motion data. Numerical simulations by NASA’s flow solvers Cart3D and US3D were performed for comparison purposes. As a result, the experimental and numerical data show a good agreement. The inclination of the cylinder strongly effects both the flowfield and aerodynamic loads. Experiments and simulations with concave cylinders showed marked difference in aerodynamic behavior due to the presence of a shock–shock interaction (SSI) near the middle of the model. Graphic abstract
... The assumption of D a (x) being a subset of D(x) is due to the fact that the actuation can be partially compromised. We assume that the distributed parameter plant exhibits stable behavior under the control system, and hence, α < 0. The PDE (1)-(2) represents systems involving diffusion (captured by u xx ) and reaction (captured by αu), for example, thermal systems [24][25][26]. ...
Security of Distributed Parameter Cyber-Physical Systems (DPCPSs) is of critical importance in the face of cyber-attack threats. Although security aspects of Cyber-Physical Systems (CPSs) modelled by Ordinary differential Equations (ODEs) have been extensively explored during the past decade, security of DPCPSs has not received its due attention despite its safety-critical nature. In this work, we explore the security aspects of DPCPSs from a system theoretic viewpoint. Specifically, we focus on DPCPSs modelled by linear parabolic Partial Differential Equations (PDEs) subject to cyber-attacks in actuation channel. First, we explore the detectability of such attacks and derive conditions for stealthy attacks. Next, we develop a design framework for cyber-attack detection algorithms based on output injection observers. Such attack detection algorithms explicitly consider stability, robustness and attack sensitivity in their design. Finally, theoretical analysis and simulation studies are performed to illustrate the effectiveness of the proposed approach.
... It is difficult to identify if this discrepancy is caused by an error in the CFD or experimental results. As the boundary-layer passes around the expansion corner it undergoes an extremely strong expansion followed by a re-compression [141,86]. This is a region of the flow where viscosity is important and as a result, different viscous flux schemes used in the CFD solution could affect the numerical results in this region. ...
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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.
This work focuses on the mathematical combination of magnetogasdynamics (MGD) and the Dunn and Kang chemical kinetic model with a multi-component Harten–Lax–van Leer Contact (HLLC) local Riemann solver using high-order interpolation schemes. Due to the nature of nearly hypersonic, unsteady, viscous, reacting MGD flows, the gas dissociates which is taken into account with 11 species and 26 chemical reactions along with the temperature dependent transport properties through the Wilke model. In the present work, the electromagnetic effects are considered by solving a Poisson equation simultaneously with the chemical reaction equations. The non-linear convective terms are treated numerically with third-, fifth- and seventh-order Weighted Essentially Non-Oscillatory (WENO) interpolation schemes along with second-, fourth- and sixth-order central finite difference schemes for the viscous terms. The combination of these numerical and physical models are analysed for stability and validated for benchmark problems such as a) Hartmann flow as MGD validation test case and b) the shock-tube problem for testing the chemical reaction model in which case the reaction rate controlling temperature suggested by Park has also been taken into account. Furthermore, a complex physical problem of an external aerodynamic flow over the cube for the Mach number varying from 1.25 to 3 have also been investigated, because a limited number of studies is available for understanding the flow behaviour of sharp cornered blunt bodies such as cubes. The knowledge of shock structures of these objects has become imperative to predict the atmospheric re-entry trajectories of de-orbiting CubeSats, space debris or fragments of launch vehicles. The main physical finding is that the shock structures are observed to be much more resolved compared to previous works, especially in the wake region. The knowledge gap is bridged by detailing the shock structures and showing the dependence of the shock standoff distance on total enthalpy of the flow.
Computational Fluid Dynamic of a Aerothermodynamics over a hot radiant blunt body the main objective of the project is to determine the pressure, velocity and temperature changes across the oblique shock wave over a wedge of angle at 14.23 an oblique shock wave is captured at Mach number 2 to read the temperature dependences, when an external energy is induced on it. The simulations are run to understand the pressure and temperature acting on the re-entry vehicle and to study the temperature dependencies on the oblique shock wave. The shock wave formed in the supersonic and hypersonic speeds are considered high-density air currents. So here we investigate what happens when temperature of the shock wave increases. To solve this we are using numerical methods like governing equation (continuity equation, momentum equation and energy equation) and ideal gas law equation. The geometry is modelled in the Design Modeller of ANSYS Workbench, meshing is carried out in Meshing tool of workbench and solving and post processing are carried out in FLUENT packages of ANSYS. The pressure and velocity differences are carried out in the three temperatures like 200 K, 300 K and 500 K. The velocity, pressure are dropped across the oblique shock wave depending on the surface temperature of the wedge, but the temperature rise across the shock is observed due to momentum change. There was no noticeable change occurred in the Mach number due to temperature change, this is because the heat transfer and radiation is turned off. Even if the heat transfer and radiation is turned on the solver encountered divergences in the solution and reported error without producing the results.
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In order to better inform hypersonic heating predictions to faceted shapes such as satellites, the hypersonic heat flux to a cube is measured experimentally using infrared thermogra-phy. Stanton numbers across the entire surface of the cube are calculated, which involves the solution of a 2-D inverse heat conduction problem. As well as revealing notable 3D flow phenomena, the results show excellent agreement with Klett's model for stagnation point heat flux to a flat surface. They also suggest that many current demise heating models struggle to accurately model heating to the sides and edges of cubes. The results also reveal the importance of taking model temperature changes as well as 2D and 3D conduction into account when calculating heat fluxes.
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The measurement of the convective wall heat flux in hypersonic flows may be particularly challenging in the presence of high-temperature gradients and when using high-thermal-conductivity materials. In this case, the solution of multidimensional problems is necessary, but it considerably increases the computational cost. In this paper, a low-computational-cost inverse data reduction technique is presented. It uses a recursive least-squares approach in combination with the trust-region-reflective algorithm as optimization procedure. The computational cost is reduced by performing the discrete Fourier transform on the discrete convective heat flux function and by identifying the most relevant coefficients as objects of the optimization algorithm. In the paper, the technique is validated by means of both synthetic data, built in order to reproduce physical conditions, and experimental data, carried out in the Hypersonic Test Facility Delft at Mach 7.5 on two wind tunnel models having different thermal properties.
Aerodynamic coefficients of a cube depending on a broad-range attitude change have been measured using the free-flight technique in a hypersonic flowfield. Experiments were performed therefore in the hypersonic wind tunnel H2K at the German Aerospace Center (DLR) in Cologne. The free-flight technique in H2K has allowed achieving a continuous rotation of the cube without any sting interferences in a broad angular range of 90°. This motion of the model during the free flight has been acquired using a nonintrusive high-speed stereo tracking system. A marker-based tracking algorithm has been applied to reconstruct the three-dimensional flight trajectories and attitudes to determine the resulting forces and moments. By using high-speed schlieren photography, the unsteady flow structures have been recorded. Pitch-angle-dependent aerodynamic coefficients of rotating cubes have been observed.
Spacecraft are typified by complex geometries meaning that predictive tools designed to assess entry, break up and ground casualty risk are not naturally suited to high fidelity modelling treatments (e.g. the CFD and FE analysis which are prevalent in the assessment of entry vehicle design and performance). Simplifying assumptions are inevitable and the consequences of these simplifying assumptions need to be investigated and quantified. The present paper is concerned with the effect of these simplifying assumptions on ground casualty risk. Specifically we report on work concerning: (i) a discussion and appraisal of some current aeroheating models as implemented in well-known tools (ii) proposed potential improvements to existing models and (iii) novel approaches to aeroheating engineering modelling. These topics are investigated in the framework of the recently developed Spacecraft Aerothermal Model tool (SAM) which can be configured to calculate ground casualty risk using varying degrees of aerothermal and break up model complexity. This allows us to investigate the sensitivity of ground casualty risk to simplifying assumptions. Parameter studies performed so far have highlighted the sensitivity of ground casualty risk to the treatment of fragment aeroheating. SAM has the option to calculate the aeroheating to fragments taking the component size, orientation and shape into account. This goes beyond simple panel inclination methodologies in common use, but is nonetheless based on established engineering correlations. As far as the authors are aware, such a methodology is not currently used by any of the well-known spacecraft breakup tools. The consequence of this novel treatment of fragment heating is the main topic of the current paper.