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DESIGN OF A HOT PLUME INTERACTION FACILITY AT DLR COLOGNE
Dominik Saile1, Daniel Kirchheck1, Ali G¨
ulhan1, and Daniel Banuti2
1DLR, German Aerospace Center, Institute of Aerodynamics and Flow Technology, 51147 Cologne, Germany
2DLR, German Aerospace Center, Institute of Aerodynamics and Flow Technology, Spacecraft, 37073 G¨
ottingen,
Germany
ABSTRACT
Space transportation systems are exposed to high thermal
and mechanical loads during the ascend in the transonic
flow regime. By now, there are still many uncertainties,
which can not be solved with state of the art computa-
tional fluid dynamic models or experiments with cold jet
flows. A test facility with a high degree of similarity to
flight with respect to the influence of the hot nozzle flow
can contribute to improve the understanding of interac-
tion effects between the hot nozzle flow and the ambient
flow by providing reliable data for validation. The objec-
tive of the paper at hand is to present the work progress
on such a facility. Issues and challenges concerning the
base flows are discussed and potential research areas for
investigations are considered. Relevant conditions during
the ascend of Ariane 5 are used as baseline and appropri-
ate scaling laws are discussed to conclude requirements
for the operational conditions for the existing wind tunnel
Vertical Test Section Cologne (VMK). These operational
conditions are used to develop a concept. After a proof of
concept is given by CFD calculations, details to the sup-
ply system including the operational range are described
and opposed to existing test benches without interaction
capabilities.
Key words: L
A
T
E
X; Space transportation system; hot
plume interaction; GOX/GH2; transonic flow regime;
base flow;.
1. INTRODUCTION
At T+ 137 s into the flight 157 of Ariane 5 ECA,a
pressure drop at the dump cooling outlet appeared,
which is characteristic for a leak in the cooling system
of the Vulcain 2 rocket engine. In the following, a rapid
degradation of the cooling of the nozzle was observed,
which led to a loss of the nozzle’s integrity and finally
to a loss of control over the launcher’s trajectory. This is
what was presented in Paris by the inquiry board in the
press conference with respect to this failure [1, 2]. The
non-exhaustive definition of the design loads, combined
with a combination of various stress factors during flight
was declared to be on one of the most probable cause for
the failure.
The objective of the paper at hand is to describe the on-
going efforts at the DLR Cologne to upgrade the Vertical
Test Section Cologne (VMK) [3, 4] to enable the investi-
gation of the hot flow interaction phenomena. In detail:
The interaction of a hot exhaust jet with the ambient flow
is of interest. For this reason, VMK must be equipped
with a supply facility of gaseous oxygen and gaseous hy-
drogen. In this paper, concept is developed and a proof-
of-concept of the some aspects of the scaling issue is con-
ducted by means of numerical simulations. The concept
includes the test environment like the wind tunnel itself,
the model, the supply system and the measurement equip-
ment.
In the long-term, the objective is to provide a platform in
a field where CFD is not available yet or difficult/costly
to obtain, and thus data for validation is required. The
facility will allow the investigation of isolated flow ef-
fects like for cold flow jets and combined flow effects
with chemical reactions, radiation, particles of various
mass and size, condensation, high temperature or vis-
cosity effects over a wide Mach number range with ap-
propriate measurement techniques with a state of the art
spatial/temporal resolution for propellants being identi-
cal to current space transportation systems or reflecting
certain desired properties. This way, the hot plume test-
ing facility will help to rule out discrepancies and incon-
sistencies between in-flight measurements, wind tunnel
measurements and CFD and improve the understanding
of the flow phenomena of rockets.
The paper is structured as follows: In the next section,
a strong justification is given for such a facility by dis-
cussing the underlying motivation, which is a based on a
short discussion of the numerous issues during the ascend
of a space transportation system. This is the field of fu-
ture measurements and applications. Additionally, some
kind of inventory is given with respect to state-of-the-art
developments in the space transportation field, considera-
tions to scaling issues and the available test environment
and equipment. The next section deals with the meth-
ods, and lists how the scaling laws are applied to find the
operating range of the facility, details to the numerical
simulation, and input parameters to enable a selection of
suitable measurement techniques.
2. MOTIVATION, APPLICATIONS, JUSTIFICA-
TION AND INVENTORY
2.1. Exhaust Plume Induced Challenges on Space
Transportation Systems
The catastrophic event from the introduction is contin-
ued here to elaborate some more details to the loads that
space transportation systems are exposed. The first note
of an abnormality of Ariane 5 operation was T+ 137 s
at the booster separation, which takes place at an altitude
of about 68 km [5, 6] with about 2000 ms1. Despite
this high altitude, it is commonly accepted that the most
challenging environment wrt to mechanical loads is dur-
ing the transonic flight phase. For instance in Ref. [5], it
can be seen that the highest dynamic pressure is reached
at an altitude of 6.8km and a velocity of 487.7ms
1,
which corresponds to Mach 1.56. Ref. [7] presents flight
data and wind tunnel tests regarding the buffet loads on
the Ariane 5 afterbody. It can be seen that the unsteady
loads on the nozzle like the pressure fluctuations, vibra-
tions and actuator loads are strongest in the transonic
flight regime and decrease over the flight time and cor-
responding higher Mach numbers (Ma > 1.5). This
statement is supported in Ref. [8]. The article reports
of flight data that reveal anomalous side loads acting on
the Vulcain 1 nozzle, which where then investigated ex-
perimentally [9] and numerically [10] on tolerable loads.
ESA described the magnitude of the buffet loads as a
major concern [11]. It is attributed to disturbances in-
duced by the three-dimensional flow and the interaction
with the boosters and started a series of activities con-
cerning a possible coupling between the external flow and
the shock-separated flow in the nozzle.
In general, buffeting describes the response of struc-
tural modes to aerodynamic/buffet loads. The relevance
of axis-symmetrically excited loads was pointed out by
Fuchs et al. [12] since they contribute significantly to
the lateral force in the base region. Further, H. Wong
[13] indicates that base-flow buffeting may trigger three-
dimensional flow separation, and thus side loads during
flight, which poses a serious issue for rocket nozzles [14].
As basis for the further descriptions and discussions, the
mean flow topology of the near field at the generic base of
an axis-symmetric generic rocket base is shown in Fig. 1.
For the flow topology and the following considerations
of the paper at hand, Ariane 5 is used as reference with
respect to the physical properties and generic, geometric
dimensions.
The flow topology sketch is based on numerical simula-
tions at Mach 0.7conducted by Deck and Thorigny [15].
The sketch shows the incoming ambient flow along the
vehicle (1) with the subsequent separation of the flow at
the end of the main body (2). The flow is deflected to-
1
2
3
45
67
0.99
0.01
0.01 0.99
Figure 1. Near field flow topology at the base for an
overexpanded nozzle flow in a subsonic external flow
(sketched from Ref. [15]).
wards the nozzle and reattaches - depending on the ambi-
ent conditions, which is mainly the Mach number - on the
nozzle (3) or on the supersonic jet acting as fluidic wall.
The ambient flow and the vehicle/model walls enclose a
highly energetic and unsteady recirculation region (4). In
a time-averaged consideration, the flow features a large
primary vortex (5) and a secondary corner vortex (6). The
transonic flight phase is reached at a very early stage dur-
ing the ascent (see e.g. Ref. [5]), thus the nozzle flow (7)
- at least for the Vulcain 2 engine with a very high expan-
sion ratio - is overexpanded (Ref. [16]). The supersonic
jet is predominantly inviscid and turbulent mixing pro-
cesses takes place at a confined shear layer (7) to the am-
bient flow. The isolines marked with 0.01 and 0.99 depict
the correspondent local velocity with normalized with the
ambient flow. Additionally, the velocity profiles in the re-
circulation region are plotted at different distances down-
stream from the separation. Note that the instantaneous
flow fields do not resemble the averaged flow field and is
highly turbulent.
The emanating shear layer from the edge of the main
body is assumed to be the driving force for the buffet-
ing instability. Kelvin-Helmholtz form at the shear layer
interface between the recirculation region, travel down-
stream and impinge on the Vulcain engine nozzle. Ac-
cording to the hypothesis of Wong et al. [8], this results
in acoustic waves generated at the reattachment location,
which travel upstream and may couple with the separa-
tion from the shoulder. In resonance, the effect may lead
to an amplification of the amplitudes in the shear layer
excitation. Via this feed-back loop, energy might be fed
into recirculation region.
This aero-acoustic feedback loop hypothesis was re-
viewed by Deck and Thorigny [15]. His investigations re-
veal that the axis-symmetric separating/reattaching flow
is governed by large scale coherent motions. According
to experiments run by Depr´
es [17], more than 90% of the
pressure fluctuations can be attributed to the antisymmet-
ric mode m=1, which is associated with helical vortex
structures randomly orientated in the azimuthal direction
as described by Fuchs et al. [12] for a flow past a circu-
lar disk. Further, it is stated by Deck and Thorigny [15]
that shear layer flapping and shear layer shedding are dif-
ferent aspects of the same motion. The first describes
a periodic growth and decay process of the recirculation
bubble in the vertical direction with a characteristic fre-
quency of fLr/U1⇡0.08 (based on the reattachment
length Lr). Further, Ref. [15] finds the secondary corner
vortex growing and decaying with the same frequency.
Shear layer shedding and the movement of the reattach-
ment length are associated with fLr/U1⇡0.2. Ac-
cording to experiments mentioned in Ref. [15], 50% of
the pressure fluctuations are attributed to this characteris-
tic frequency. The onset for such a coherent large scale
motion results from a feedback process between the sepa-
ration point and the maximum intensity backflow location
of the recirculation region, which is found to be origin
of the absolutely unstable flow region. The generation
of pressure waves is attributed to an interaction between
the passage of large-scale and free-stream inrush between
vortices. Recently, Weiss et al. [18] confirmed the hy-
pothesis of an absolute unstable flow region by means of
linear stability analysis coupled with a two-point correla-
tion analysis.
Unsteady loads are only one reason for the justification of
a hot plume testing facility. Ref. [19] lists many applica-
tions where an improved knowledge about rocket exhaust
plumes is of interest like for the (1) design, the develop-
ment and the operation of flight vehicles, to (2) detect
and track flight vehicles, to (3) minimize radio-frequency
interference and to (4) develop generally a better under-
standing for the plume behavior to provide foundations
for simulations, to optimize intrinsic features wrt radio
frequency attenuation, noise et cetera or to investigate the
environmental impact.
Acoustic plume noise has to be added to the list of me-
chanical loads that act on the space transportation sys-
tem. For instance, Saturn V emitted acoustic power of
about 2·108W[19], which corresponds to the power
consumption of 200000 average homes. The angle of at-
tack plays a major role in the dynamics of base region.
The reattachment of the flow on the other hand, either
on the nozzle wall (as shown in Fig. 1) or on the jet act-
ing as fluidic wall has according to the investigations of
Depr´
es et al. [20] no influence on the pressure distribution
and spectral behavior in the base region. A jet tempera-
ture increase from T0= 273 K to 800 K with a nozzle,
which is immersed in the recirculation region, increases
the base pressure by 15% as demonstrated by Zapryagaev
et al. [21]. This in turn results in an overestimation of the
base drag, which holds true to observations in real scale
as shown with in-flight measurements conducted on Ari-
ane 5. This insinuates that the simulation capability with
appropriate jet temperatures is required.
With increasing altitudes and decreasing ambient pres-
sure, the plume diameter increases. For a multiple nozzle
space transportation configuration, the interaction of mul-
tiple plumes influences and distorts the flow pattern, and
the altered heat transfer process may cause a low density
region in the base, which may significantly influence on
the drag. Systems for drag reduction and for the damping
of buffet loads are under investigation.
An overview to heat sources is given in Goethert [22].
It is classified in (1) radiation caused by the exhaust gas
and the rocket nozzle, (2) the recirculation of hot exhaust
gases into the base region without afterburning, the (3)
recirculation of fuel-rich exhaust gases with afterburn-
ing at the base and the interaction of exhausts or multi-
ple nozzle effects. Simmons [23] describes the impact
on the emissive properties if the hot exhaust gases carry
soot particles like from hydrocarbon fuels, aluminum ox-
ide from solid propellants or condensed particles. The
particles flow exhibit a lag of velocity and temperature
and causes a separation in the two-phase flow, which can
alter the plume shape significantly [24]. The slow down
of the typically fuel rich exhaust gases with incompletely
oxidized fuel species in the shear layer of the ambient
air and the jet causes a temperature recovery, which then
promotes afterburning and IR emission. The recircula-
tion region can here act as a flameholder and sustain this
process. As described before, a multiple nozzle config-
uration lead at higher altitudes to the impingement and
interaction of the separate jets, which causes hot gases to
travel even upstream on the space vehicle [22]. The con-
sequences are high temperature areas with the inherent
effects of an increased emission, increased temperatures
in the shock-heated air, the augmentation of hot backflow
and upstream separation. The plume separated region can
act under such circumstances as flameholder [25].
Further, Goethert [22] categorizes the significance of the
different heat sources for a rocket in general. Afterburn-
ing dominates in the low altitude range up to an altitude of
about 4.5km (15000 ft). At medium altitudes at 4.5km
to 15 km (50000 ft), afterburning has a major contribu-
tion to heat loads and a significant part comes from the
plume interference. The first diminishes completely for
altitudes above 15 km, while the latter is described as
major. Additionally, recirculation of hot gases without
afterburning for a multi-nozzle rocket configuration and
radiation reaches a significant heat load level.
The plume trail is of great interest concerning environ-
mental issues or the detectability and trackability of flight
vehicle. The first relates to impact of the plume remains
in the atmosphere and specifically in the stratosphere or
ozone layer. The latter is described with the terminol-
ogy rocket exhaust signatures and involves the subject of
primary and secondary smoke, its classification, plume
microwave and radiation properties. A better understand-
ing understand of the mechanisms of smoke, soot, or
vapor formation and the possibility to control them is
required. One of the objectives is the minimization of
radio-frequency interference and reduction of attenuation
for specific antennas due to the plume. Kinefuchi et al.
[26] reports of these effects for the Japanese M-V rocket
and the European VEGA launch vehicle due to the usage
of solid propulsion systems. In this regard, the rocket ex-
haust plume is seen as near field of the trail and starting
point for further predictions of the trail [24].
Deposition and possible contamination of condensed
species on the space transportation system and its com-
ponents being optical surfaces, windows, solar panels or
radiating heat emission surfaces due to plume impinge-
ment on vehicle is a further aspect of consideration. At
high altitudes, the supersonic plume exhibits a Prandtl-
Meyer expansion angles larger than 90and the subsonic
boundary layer, although low, but not negligible in mass
flow, can even travel further upstream with all associated
issues induced by the exhaust gas.
The current revival of reusable launch vehicles and
SpaceX efforts with the Falcon 9 Reusable Development
Vehicle (F9R Dev) as shown in Ref. [27] opens up a less
investigated field for future investigations. Firing in retro
mode certainly involves many of the above described
challenges, e.g. instabilities as shown in Ref. [28], over a
wide Mach number range in a different fashion.
The aforementioned examples reveal that base region is
an area of uncertainty in the current launch vehicle design
process. In the past, base flow effects have mostly been
examined in experiments with cold supersonic jets, which
have shown to not mirror the base flow effects satisfacto-
rily. The objective of the hot plume interaction facility
is to provide a platform for the examination of base flow
phenomena under realistic flight. Due to the importance
of Ariane 5 wrt the European space transportation sys-
tem, the Vulcain engine, and thus oxygen and hydrogen,
is chosen as baseline for the hot plume facility.
2.2. State of the Art and Future Developments
Various test benches with focus on investigating cryo-
genic injection and combustion phenomena are operated
throughout the world with small scale combustion cham-
bers. The objective of the work at hand is not on these
phenomena, but the parameters of the test benches is used
to get an idea about the classification of the future inter-
action facility. These test benches investigate combustion
chambers, which correspond to applicable chamber with
respect to the size. An exemplary overview about exist-
ing facilities is given in Tab. 1. It contains the Cryogenic
Combustion Laboratory (CCL) of the PennState Univer-
sity, the Mascotte test bench of Onera and the small scale
test facility M3.1 of DLR Lampoldshausen.
To cope with potential future developments, the facility
is designed to feature methane compatibility and oper-
ability. This is driven by recent efforts throughout the
world towards methane/oxygen engines. Airbus Defence
and Space Defense initiated LOX/Methane studies for
rocket engines of a 350 kN,420 kN and 600 kN thrust
class named ACE-35R, ACE-42R and ACE-60R, respec-
tively, followed by sub-scale and equipment tests. It is
stated that the engine demonstrator ACE-35R could be
ready for test in 2018 Ref. [34]. Firing test of a European
staged combustion engines are presented in Ref. [35]. In
2012, SpaceX announced the development of a methane-
fueled rocket engine comparable in thrust with the F-1
engine of the Saturn V named Raptor. The Russian-based
NPO Energomash company considers the modification of
their RD-169, RD-182, RD-184, RD-185, RD-190 and
RD-192 rocket engines to operate with liquefied natural
instead of kerosene [36]. French/Russian activities are
summarized in Ref. [37]. The future launcher preparatory
program (FLPP) of ESA focuses the future demonstra-
tion work on two propellant combinations, one of which
is LOX/methane [38, 39]. A general overview to the liq-
uid propulsion systems and propellant choices is given
in Ref. [40]. NASA, Aerojet and the German Aerospace
Agency are mentioned to pursue activities in the respec-
tive field. Korea and Japan performed tests, which are
described in Ref. [41, 42], respectively.
2.3. Scaling Considerations
The main parameters governing the base flow of a rocket
were investigated by Goethert and Barnes [43]. Under the
assumption of geometric similarity, it is concluded that
non-viscous and viscous effects have to be considered.
Similarity for the non-viscous effects are satisfied: if (G-
1) the static pressure change case by the a change in flow
direction for full-scale vehicle and model is equal for the
external flow, if (G-2) the static pressure change caused
by a change in the flow direction for full-scale vehicle and
model is met for the exhaust jet, and if (G-3) the ratio of
the static pressure in the jet at the nozzle exit to the static
pressure in the undisturbed external flow for the full-scale
vehicle and the model is equal. The first two aspects re-
late to the pressure sensitivity of the external/jet flow with
respect to the flow direction, which is required to be con-
stant for both hot and cold flow. The non-viscous scaling
rules (G-1 to G-3) were applied in Ref. [44] and have
shown to find a good correlation for the base pressure be-
tween the hot and cold gas nozzle flows.
For the viscous effects, Ref. [43] considers the mixing
process between the ambient flow and the exhaust jet as
dominating and neglects the mixing process in the base
region due to the small velocities, and thus only concerns
the mixing of an essentially parallel flow. Two similarity
parameters are developed: The excess pumping mass (G-
4) parameter and the total pressure at the jet boundary
streamline (G-5). The first is a measure for the entrained
mass of the mixing layer, the second is required to pro-
vide similar discharge conditions.
More recently, Frey [45] suggested a scaling approach
for Ariane 5’s point of maximum buffeting at Mach 0.7.
It is based on shear layer similarity with respect to plume
shape and entrainment arguing that the shear layer is the
only way of interaction between the nozzle flow and the
ambient flow. Consequently shear layer growth must be
equal for flight and wind tunnel experiment. Six param-
eters describing the conditions directly at the nozzle exit
are identified to govern this process: (F-1) The velocity of
the plume, (F-2) the velocity of the surrounding external
flow at the nozzle lip, (F-3) the density ratio between the
external and plume flow, (F-4) the degree of turbulence
of the internal flow, respectively the degree of turbulence
of the boundary layer along the inside of the nozzle, (F-
5) the angle between the shear layer direction and the
nozzle axis, and (F-6) the nozzle lip thickness. The pa-
rameters refer to the location downstream from the nozzle
lip, and respectively downstream from the pressure adap-
tion of the nozzle resulting in a shock or expansion fan.
Further considerations a dedicated to the issue of reach-
ing flight-comparably high jet exit velocities. Sensitivi-
Table 1. Operational parameters of exemplary test benches. The abbreviation (L) and (G) specify the state of the
fuel/oxydizer as liquid and/or gaseous, respectively.
Test bench CCL [29, 30] M3.1 [31, 32] Mascotte [33]
Total mass flow ˙mgs1-400 (L) -
Mass flow oxygen ˙mO2gs1450 (G/L) - 40 400 (L)
Mass flow hydrogen ˙mH2gs1113(G) - 575 (G)
Pressure limit pCC,max MPa 9.6 4 10
ties like the specific heat ratio, the influence of the molar
mass of exhaust gas, gas temperature and mixing ratio
are addressed with the goal to increase the exit veloc-
ity. It is concluded that the highest degree of similarity in
terms of exit velocity is reached either by using helium,
as being used in Ref. [46], or even better by operating a
combustion chamber at a hydrogen rich mixture ratio of
about 0.7. The latter yields at 82% of the Vulcain exit
velocity in comparison to helium or cold air, which is at
56% and 15%, respectively. The low mixture ratio comes
with the benefit for the materials used for the combus-
tion chamber due to combustion chamber temperatures
of about 1000 K.
2.4. Test Environment
In the current state, VMK as shown in (Fig. 2) is a blow-
down type wind tunnel featuring a vertical free test sec-
tion for tests in the subsonic to supersonic range starting
from Mach 0.5up to 3.2. The stagnation temperature
can be adjusted up to 700 K and a stagnation pressure up
to 3.5MPa is possible. As a result, sea level conditions
can be simulated up to a Mach number of 2.96. This
and the operational range of VMK in terms of adjustable
unit Reynolds number range for the various Mach num-
bers is shown in Fig. 3. Three different axis-symmetric
nozzles with an exit diameter of 184 mm,270 mm and
340 mm are available for subsonic tests. For supersonic
tests, VMK facilitates tests with 14 axis-symmetric noz-
zles exhibiting an exit diameter of 150 mm,230 mm and
310 mm.
Further, the test section of VMK is integrated in tower
with a height of 11 m and a ground area of 4m by 4m.
The concrete walls have a thickness 70 cm and all struc-
tural elements like schlieren windows or steel doors are
designed to be explosion-proof. Consequently, VMK is
predestinated for wind tunnel experiments with propel-
lants and explosives, which is shown in the long his-
tory of experiments regarding side jets with solid propel-
lants (e.g. Ref. [47, 48]) or ram jets (Fig. 4) with in-
tegrated and realistic combustion chambers fueled with
various propellants like hydrogen, carbon based propel-
lants or others. Recently, the range of applications for
VMK was extended to investigate hovering re-entry cap-
sules in free flight as examined for the MarcoPolo-R ERC
capsule [49, 50]. The paper at hand summarizes the
ongoing work progress to extend VMK’s capabilities to
Supersonic and Hypersonic Technology Department, Cologne
Description of the Vertical Wind Tunnel Cologne (VMK)
-
1
-
Facility components and flow condition
TheverticalwindtunnelCologneisaconventional
blowdownfacilitysimulatingsub-andsupersonic
flowconditions.Itisoperatedintermittentlyin
accordancewiththestorageprinciple,i.e.com-
pressedairandheatisstoredpriortothetestand
thenreleasedduringthetestrun.Mainfacility
componentsaredescribedinthefollowingsections
(Fig.1).
Fig. 1: VMK facility components.
Beforestorage,airiscompressedbycentrifugal
andpistoncompressors,andalsodriedinorderto
reduceitsdewpoint.Thetotalvolumeofthefive
storagevesselsis1,000cubicmetersandtheir
maximumallowablepressureis60bar.Aquick
reactingcontrolvalveisusedtoadjustthestagna-
tionpressureoftheairsupplytothedesiredflow
condition.
Becausetheaircoolsdownwhenitisaccelerated
inthewindtunnelnozzle(convertingpotentialinto
kineticenergy),thefacilityisequippedwithtwooil
burnersof2,000kWtotalcapacityinordertopre-
heattheairtothedesiredtemperaturelevel.
Thereby,astagnationtemperatureofupto477°C
canbeadjusted,i.e.statictemperaturesof15°C
canbereachedforMachnumbersuptoMach2.8.
Togeneratedifferentflowconditionsthreesub-
sonicnozzlesof184,270and340mmdiameter
and14supersonicnozzlesof150,230and
310mmdiameterintheMachnumberrangefrom
1.57to3.23areavailable(Fig.2).
100 200 300
0
0.5
1
1.5
2
2.5
3
3.5
4
Unit Reynoldsnumber [106/1m]
Machnumber
100 200 300
0
0.5
1
1.5
2
2.5
3
3.5
4
5020
10
Max. supersonic Mach number
Min. subsonic Mach number
Min. supersonic Mach number
Max. subsonic Mach number
pt=35bar
Tt=750 K
pt=35bar
Tt=288K
ps=1,0132bar
Tt=288 K
Sea level cond.:
ps=1,0132bar
Ts=288K
ps=1,0132bar
Tt=750K
Fig. 2: Reynolds number range of VMK facility.
Theoperatingrangeofthe“subsonic”nozzlesis
limitedontheonehandbyreachingsoundvelocity
atthethroatandontheotherhandbythecontrol
rangeofthequickreactionvalve.Tillnow,Mach
numbersfrom0.5upto0.95atverygoodPitot
pressurehomogeneitywereachieved(Fig.3).
Fig. 3: Pitot profiles at different locations down-
stream the nozzle exit at Mach number 0.8.
Atthethroatofthe“supersonic”nozzlessound
velocityisreached.Forthiscase,theoperation
pressureisnormallyadjustedsothatthestatic
pressureinthejetisadaptedtothepressureinthe
testsection,i.e.thestreamlinesleavethenozzle
exitparalleltothemeanflowdirection.Ifthestag-
nationpressureisfurtherdecreased,theflowcon-
strictsuntilitfinallybreaksdown.Forpre-heated
air,runsusuallylastbetween30sand60s,while
severalminutesoftesttimecanbereachedfor
coldflow.
Figure 2. VMK facility components.
Supersonic and Hypersonic Technology Department, Cologne
Description of the Vertical Wind Tunnel Cologne (VMK)
-
1
-
Facility components and flow condition
TheverticalwindtunnelCologneisaconventional
blowdownfacilitysimulatingsub-andsupersonic
flowconditions.Itisoperatedintermittentlyin
accordancewiththestorageprinciple,i.e.com-
pressedairandheatisstoredpriortothetestand
thenreleasedduringthetestrun.Mainfacility
componentsaredescribedinthefollowingsections
(Fig.1).
Fig. 1: VMK facility components.
Beforestorage,airiscompressedbycentrifugal
andpistoncompressors,andalsodriedinorderto
reduceitsdewpoint.Thetotalvolumeofthefive
storagevesselsis1,000cubicmetersandtheir
maximumallowablepressureis60bar.Aquick
reactingcontrolvalveisusedtoadjustthestagna-
tionpressureoftheairsupplytothedesiredflow
condition.
Becausetheaircoolsdownwhenitisaccelerated
inthewindtunnelnozzle(convertingpotentialinto
kineticenergy),thefacilityisequippedwithtwooil
burnersof2,000kWtotalcapacityinordertopre-
heattheairtothedesiredtemperaturelevel.
Thereby,astagnationtemperatureofupto477°C
canbeadjusted,i.e.statictemperaturesof15°C
canbereachedforMachnumbersuptoMach2.8.
Togeneratedifferentflowconditionsthreesub-
sonicnozzlesof184,270and340mmdiameter
and14supersonicnozzlesof150,230and
310mmdiameterintheMachnumberrangefrom
1.57to3.23areavailable(Fig.2).
100 200 300
0
0.5
1
1.5
2
2.5
3
3.5
4
Unit Reynolds number[106/1m]
Machnumber
100 200 300
0
0.5
1
1.5
2
2.5
3
3.5
4
5020
10
Max. supersonic Mach number
Min. subsonic Mach number
Min. supersonic Mach number
Max. subsonic Mach number
pt=35bar
Tt=750K
pt=35bar
Tt=288K
ps=1,0132bar
Tt=288K
Sea level cond.:
ps=1,0132bar
Ts=288K
ps=1,0132bar
Tt=750K
Fig. 2: Reynolds number range of VMK facility.
Theoperatingrangeofthe“subsonic”nozzlesis
limitedontheonehandbyreachingsoundvelocity
atthethroatandontheotherhandbythecontrol
rangeofthequickreactionvalve.Tillnow,Mach
numbersfrom0.5upto0.95atverygoodPitot
pressurehomogeneitywereachieved(Fig.3).
Fig. 3: Pitot profiles at different locations down-
stream the nozzle exit at Mach number 0.8.
Atthethroatofthe“supersonic”nozzlessound
velocityisreached.Forthiscase,theoperation
pressureisnormallyadjustedsothatthestatic
pressureinthejetisadaptedtothepressureinthe
testsection,i.e.thestreamlinesleavethenozzle
exitparalleltothemeanflowdirection.Ifthestag-
nationpressureisfurtherdecreased,theflowcon-
strictsuntilitfinallybreaksdown.Forpre-heated
air,runsusuallylastbetween30sand60s,while
severalminutesoftesttimecanbereachedfor
coldflow.
Figure 3. Reynolds number range of VMK facility.
Figure 4. Exemplary propulsion test of a RAM-jet config-
uration in VMK.
perform hot plume tests with gaseous hydrogen/gaseous
oxygen combustion for the simulation of thermal and me-
chanical loads on the base of space transportation ve-
hicles during the ascent with focus on the most critical
part, which is transonic flow regime. As reference for
planned hot plume interaction facility in Cologne serves
the main stage of Ariane 5 with its Vulcain engine since it
is the main European space transportation system. In the
past, experiments were performed with the solid propel-
lant combination based on HTPB/AP/Al as used in the
EAP boosters of Ariane 5. In the future, this upgrade
empowers experiments in VMK to study generic single
plume effects with various propellants and complete Ar-
iane 6-like configurations with plume-plume interaction
and the resulting effects on the structure. Additionally,
the methane compatibility of the facility covers possible
future developments as discussed in industry and various
international space agencies. Consequently, base flow ex-
periments can be run for all current and possible future
members of the European space launcher family. This in-
cludes Vega,Ariane 5,Ariane 6 and most FLPP studies.
The upgrade of the hot plume testing facility can be
divided in several steps. Step 1: The realization of
the hydrogen/oxygen supply system is scheduled for
2015/2016. Subsequently, experiments can directly be
conducted after the manufacturing of a combustion cham-
ber and integration in a already partially existing wind
tunnel model from investigations conducted by Emunds
and H. Riedel [51]. The targeted diameter for the rep-
resentation of the main stage is in correlation with these
previous experiments at 67 mm for a subsonic wind tun-
nel nozzle with 340 mm.
As a mid-term plan (step 2), it is planned to equip VMK
with a larger subsonic wind tunnel nozzle of the size
of 600 mm. This makes the 1:1integration of the
PennState combustion chamber with an external diame-
ter of 148 mm in the wind tunnel model seem reason-
able. The PennState burner was selected for several rea-
sons. Among others since it is run with gaseous oxygen
and gaseous hydrogen up to the stoichiometric mixture
ratio, since it is considered to be a benchmark test case,
and thus well experimentally [29] and numerically [52]
investigated and documented. For instance, details to the
dimensions of the combustion chamber and injector are
given in Ref. [30]. The CFD in Ch. 4 is based on this
upgrade step 2.
The realization of step 3 falls also in the range of mid-
term planning. The department is currently studying op-
portunities to reduce the ambient pressure in the measure-
ment section. This aspect is advantageous with respect to
the adjustment of the pressure ratio between the external
and nozzle flow and basically allow altitude simulation
like during the ascent of space transportation systems.
In the long term (step 4), the principle of the upstream
support concept will also be applied for supersonic test-
ing. This can be done with appropriately designed wind
tunnel nozzles as shown for example in Ref. [53, 54]. In-
dependently from step 4, classical strut supported con-
figuration can easily be test with the existing supersonic
wind tunnel nozzles of VMK.
The developments and experiences for VMK will be
used as blueprint to establish hot plume testing capabili-
ties in Trisonic Test Section Cologne (TMK) [55, 56] in
Cologne. This facility is equipped with a transonic test
section with perforated walls for an improved flow qual-
ity. Perturbations coming from free jet shear layer of an
open test section as described by Weiss and Deck [57] can
be avoided this way. Further advantages lay in the pos-
sibility to adjust the static pressure in the measurement
section with little modification for altitude simulations,
the possibility to change the angle of attack/yaw and ex-
periments can be conducted continuously from Mach 0.5
to 5.7. The comparison of experiments in both facilities
provide insights into the influence of the support.
3. METHODS
A justification for a hot plume facility was given in
Ch. 2. The subsequent question to clarify is how to es-
tablish similarity to flight. An introduction to the scal-
ing issue was given in Ch. 2.3 based on the work by
Ref. Goethert and Barnes [43] and Frey [45]. The two
scaling approaches are opposed to each other and eval-
uated in Ch. 3.1 to find appropriate parameters that re-
flect similarity to Ariane 5 over the complete transonic
Mach number range. For this task, it is necessary to
make realistic assumptions, which is based on existing
hardware. The PennState burner was chosen as refer-
ence for the hot gas generation. It is run at maximum
chamber pressure of 6.89 MPa. For comparison: Vul-
cain 2 operates at ⇠12 MPa. The wind tunnel VMK in
Cologne is taken as a given second prescription. Note that
the static pressure at the exit is always the ambient pres-
sure due to the open test section environment. In Ch. re-
fchap:TestEnvironment, it is made use of the scaling as-
sessment to define the reference conditions for the future
facility. The DLR TAU code and the applied method is
described in Ch. 3.3.
3.1. Scaling Considerations
The non-viscous scaling parameter of Goethert and
Barnes [43] referring to the static pressure change of the
nozzle supersonic model flow (G-2) can be expressed as
a function of Mach number and heat capacity ratio. For
hot gas tests with comparable mixture ratios to flight, it
essentially reflects Mach number similarity for the nozzle
flow if the latter is matched. But, this similarity param-
eter is difficult to keep since, on the one hand, the com-
bustion chamber pressure must be large enough to avoid
the violation of the pressure ratio (G-3) and to avoid a
significant alteration of the plume pattern or even flow
separation in the model nozzle. On the other hand, the
maximum chamber pressure is limited in order to handle
the heat flux especially at throat of the nozzle for small
burners. Consequently, the exit Mach number, ergo the
expansion ratio, is adapted to match the pressure ratio
condition G-3. The pressure change condition G-1 for
the ambient flow is imposed by keeping the Mach num-
ber M1,Exp =M1,Ariane similar. This can easily be
set for wind tunnels by adjusting the stagnation pressure
(for a subsonic nozzle).
The viscous scaling parameters require the displacement
streamline, which is modeled as function of the jet’s
Mach number M1. Further input parameters for the ex-
cess pumping mass (G-4) and the total pressure head (G-
5) are only the density and velocity of both the jet and the
ambient flow.
The similarity parameter of Ref. [45] include parameters,
which are dependent on the specific behavior or geometry
of the wind tunnel model. Thus, the degree of turbulence
(F-4) and the nozzle lip thickness (F-6) are not taken into
account for the design at the current state. The other pa-
rameters find their equivalence. The velocity condition
F-1 and F-2 are comparable to the Mach number similar-
ity M-1 and M-2, respectively, and the angle of the shear
layer direction (F-5) is driven by the pressure ratio (G-3).
For the design of the facility, the intention is not to find
a wind tunnel model with high degree of similarity to the
wake flow of Ariane 5 with a similar exhaust plume. The
idea is rather to show the potential of the hot plume sys-
tem in combination with VMK over a wide Mach range
of the transonic flow regime.
The methodology to judge similarity is to determine the
nozzle exit and ambient properties (p,⇢,v,⇢v,T) dur-
ing the ascent of Ariane 5 and set them in relation to the
properties of the nozzle exit and ambient properties of the
planned experiments. The ratio of both is equal to 1for
all properties if the conditions are identical.
In more detail, the idealized nozzle exit parameters for
Ariane 5 and PennState burner plus defined nozzle were
calculated with Rocket Propulsion Analysis [58]. The Ar-
iane 5E H173 version as given in Ref. [59] was taken as
reference. The trajectory and the corresponding parame-
ters during the ascent was calculated with the trajectory
simulation program (Ref. [60]) for Ariane 5. It returns,
among other parameters, the velocity as function of alti-
tude. By means of the atmospheric data, in the case here
the U.S. standard atmopshere [61], the static gas prop-
erties and the velocity were converted and linked to the
Mach number, which are of concern for the experiments.
The static gas properties of the wind tunnel are defined
for all Mach number due the pressure adaptation at the
nozzle exit for the subsonic case of concern. As it can be
seen in Fig. 3, the temperature is variable. For subsonic
tests, this is unusual, thus the reservoir pressure is set to
288 K.
3.2. Test Environment
In order to receive an estimated maximum and minimum
mass flux for the hot plume testing facility, a combus-
tion chamber had to be found or designed, which can be
integrated in a wind tunnel model. A literature research
revealed that the PennState burner is suitable combustion
chamber to continue the study. It corresponds to the di-
mensional limitations, is well documented, features Vul-
cain 2-like combustion chamber condition and a co-axial
injector, which is of relevance for rocket engine applica-
tion. More to the reasoning is given in Ref. [62].
The outer diameter of the VMK wind tunnel model is
approximated as follows: The throat diameter of the
PennState burner, the expansion rate of the Vulcain 2 en-
gine and the Ariane 5-like diameter ratio between the di-
ameter of the main body and the diameter of the nozzle of
0.4is used to calculate the inner nozzle exit diameter and
the diameter of the main body. It accounts to 61.2mm
and 158 mm (including a nozzle thickness of 2mm), re-
spectively. The outer diameter of the PennState burner
is in a comparable range (148 mm). Note that this is an
imaginary model and this approach is used to determine
the maximal outer diameter. Such a configuration would
require a very high combustion chamber pressure to avoid
nozzle flow separation. This in turn causes very high heat
loads on the model.
Three reference conditions (RC 0 to RC 2) are defined
to cover the operating capabilities of the facility. Refer-
ence conditions RC 0 defines a ’cold’ and low pressure
condition, RC 1 is a ’must’ condition to cover the capa-
bilities of the PennState burner, and RC 2 is based on a
maximum, ’nice-to-have’ scenario. No attention is paid
to the question if the heat loads of RC 2 condition can ac-
tually be managed for this exact configuration. The idea
is to enable experiments like this with the hot plume in-
teraction facility and to cover a broader range of potential
experiments with to date undefined configurations.
In detail, RC 0 is a hydrogen-rich condition with a mix-
ture ratio of 0.7at a chamber pressure of 2.07 MPa for a
nozzle throat diameter of 8mm. Experiments with con-
dition RC 0 reach very high exit velocities, meaning high
similarity in that respect, at moderate demands on the
materials. According to Ref. [45], the exit velocity can
go up to umax = 3467.1ms
1for expansion in vac-
uum, which corresponds to 82.0% of the Ariane 5’s exit
velocity. Special precautions for heat management can
be circumvented and thus, a wind tunnel model featuring
such a combustion conditions can easily be integrated in
TMK. Condition RC 1 is specified by a chamber pressure
of 6.89 MPa, a mixture ratio of 6.0and a nozzle throat
diameter of 8.0. An expansion ratio of 20 is an addi-
tional requirement to avoid nozzle flow separation during
the experiments (for flow separation see e.g. Ref. [14]).
Condition RC 2 is motivated by two different drivers. The
first driver is the desire to conduct experiments in a pres-
sure range, which is comparable to Ariane 5, meaning up
to ⇠12 MPa. The second driver corresponds to the re-
quirement to test RC 1 condition for a nozzle diameter
that allows similarity with respect to the outer geometric
scaling between Ariane 5 and the wind tunnel model. The
outer diameter of the VMK model (158 mm) leads with
geometric outer scaling and an expansion ratio of ⇠20
to a nozzle throat diameter of about 13.7mm. The result-
ing mass flows of these reference conditions are listed in
Tab. 2. Obviously, if the mass flux is held constant, RC 2
defines the upper limit of the operating range.
The aforementioned reference conditions are valid for the
operational mode. Additionally, ignition conditions IC
must be specified for the ignition mode. The idea is to
ignite the combustion chamber through a nozzle with a
flare system. The ignition mode targets at low pressures
and low a mixture ratio to avoid a pressure overshoot and
to keep the heat loads low in that phase. To be on the
secure side with the ignition through the nozzle, it is in-
tended to stay below conditions that lead to a choked noz-
zle flow. In discussion with industry, the minimum mass
flow for the supply system (see Ch. 4) was determined
and documented (Tab. 2). This minimum mass flow is
considered as sufficient. Ignition with a mixture ratio of
2and a combustion chamber pressure of 0.16MPa can be
realized with nozzles exhibiting a throat diameter from
4.2mm to almost 24.5mm.
This statement is supported by literature on ignition.
Hasegawa et al. [63] successfully conducted laser igni-
tion experiments for comparable parameters and condi-
tions, meaning a mixture ratio O/F =2, combustion
chamber pressures ranging from 0.15 MPa to 0.35 MPa
and a throat diameter of dth =3mm. The calculated
hydrogen mass flow ranges from 0.14 gs1to 0.33 gs1.
An anchored flame was observed by Schmidt et al. [64]
for a nozzle with a throat diameter of 4mm, a mixture
ratio of about 2and a hydrogen mass flow of 0.58 gs1.
3.3. Computational Fluid Dynamics
Computations are carried out using the DLR TAU Code.
It is discussed in detail in the literature see e.g. Ref. [65,
66]. TAU is a hybrid grid, finite volume second order ac-
curacy flow solver. It has been validated for a variety of
steady and unsteady flow cases, ranging from sub to hy-
personic Mach numbers Ref. [67, 68].
As the investigated geometries are essentially axis-
symmetric in this work, all computations are carried out
in two dimensions under assumption of axisymmetry. For
this initial study, the involved gases are treated as a mix-
ture of non-reacting ideal gases. Species accounted for
are H2, H, O2, O, H2O, OH, HO2, H2O2, and N2. The
reservoir composition inside the combustion chamber is
computed with NASA’s “Chemical Equilibrium and Ap-
plications” (CEA) code by Gordon and McBride [67, 68].
The combustion chamber is simulated to operate with a
pressure of 11.9MPa at 3669 K and species correspond-
ing to a mixture ratio of 7.1. The flow is released through
a Vulcain 2 nozzle (scaled to fit on the wind tunnel model
concept) in the ambient Mach 0.86. The stagnation pres-
sure and temperature of the facility is set to 0.164 MPa
and 288 K, respectively. Since an open test section is sim-
ulated, the ambient pressure is initiated with 0.102 MPa.
3.4. Measurement Techniques
The requirements for a suitable selection of measurement
techniques for the proposed concept are listed in the
following. In detail, the test section diameter and length
is according to the planned dimensions of VMK at
600 mm and a run time of the facility between 2s and
50 s is assumed. In the transonic range, the free stream
velocity and Mach number is in the range of 160 ms1
to 360 ms1and 0.5to 1.2, respectively, at a free stream
pressure of about 1 bar. The diameter of wind tunnel
model is in the range of 50 mm to 160 mm, while the
wind tunnel nozzle diameter is about 40% of the model
diameter. The exhaust jet velocity and Mach number was
specified to be between 3000 ms1and 4400 ms1at an
exhaust jet static temperature of 340 K to 2000 K.
4. RESULTS AND DISCUSSION
4.1. Scaling considerations
Reynolds number similarity is demanded for most simu-
lations in aerodynamics. Fig. 5 depicts the unit Reynolds
number ReUof a representative ascend trajectory of Ar-
iane 5 ECA as function of the ambient Mach number
Ma1(Ref. [60]). The altitude His given additionally
to explain the influence of the atmosphere. It can be
seen that the Reynolds number increases steadily due to
the increasing velocity of the vehicle. The maximum is
reached at Ma1=0.82 at an altitude of 4.5km, which
Table 2. Reference conditions for operational (RC0-RC2) and ignition mode (IC).
Reference condition RC 0 RC 1 RC 2 IC
Combustion chamber pressure pCC MPa 2.07 6.89 11.5(15) -
Mixture ratio O/F 0.7 6.0 6.0-
Total mass flow ˙mgs189.2 150.3 440.8-
Mass flow oxygen ˙mO2gs152.5 128.8 377.8 0.6-20
Mass flow hydrogen ˙mH2gs136.7 21.5 63.0 0.3-10
0 0.5 1 1.5 2
0
2.5
5
7.5
10
12.5
15
H[km]
Ma
∞
0 0.5 1 1.5 2
0
2.5
5
7.5
10
12.5
15
x 106
ReU[m−1]
Figure 5. Exemplary Ariane 5 ascend. Altitude and unit
Reynolds number over Mach number.
is mainly invoked by a temperature curvature change in
the atmosphere. Taking into account the differences of
the base diameter as characteristic length scale between
flight and experiment of about two orders in compari-
son with the operational range VMK (Fig. 3), it becomes
clear that the Reynolds number is impossible to match
for such a configuration with experiments in VMK. For
more details, see Ref. [16]. In the previous similarity
studies in Ref. [43, 45], the Reynolds number did not
occur as influential parameters. To answer the question
about its influence, literature on wake flows is consulted.
The base pressure is taken here as reference. On the one
hand, since it was on the focus of many investigations for
a long time, and on the other hand, since it seems to be a
governing parameter as shown before. Note that this is a
necessary, but not a sufficient criteria for flow similarity.
Murthy and Osborn [69] assembled a bibliography about
base flow phenomena and presents various correlations
approaches. The Reynolds number is explicitly men-
tioned in the Kurzweg Correlation and Love Correlation.
It is stated that for wake flows without injections/jets, the
Reynolds number has an negligible influence on the base
pressure as long as the incoming boundary cylinder on
the main body is turbulent. In Ref. [69], an influence
is attributed to the Mach number, angle of attack, bound-
ary layer thickness and boundary layer characteristics and
surface temperature. Experiments conducted by Zaprya-
gaev et al. [21] in the recent past show the same insen-
sitivity to Reynolds number variations. The other influ-
ences should be kept in mind for the final definition of
0.5 0.6 0.7 0.8 0.9 1 1.1 1.2
0
1
2
3
4
5
6
Ma∞
(pe/p∞)Exp/(pe/p∞)Ariane
ϵExp = 10
ϵExp = 20
ϵExp = 40
ϵExp = 58.5
Figure 6. Pressure ratios between experiment and flight
over Mach number for various expansion ratios.
the wind tunnel model for a potential adjustment of the
boundary layer with appropriate measures like boundary
suction, blowing, roughness elements or surface temper-
ature adjustments.
For the further discussion, the Reynolds number is
consequently neglected. ’Screws’ that can be changed to
adjust the similarity for a predefined Mach numbers are
the stagnation temperature of the wind tunnel, and for the
combustion chamber the pressure, the expansion ratio
of the nozzle and the expansion ratio. The sensitivity
of these parameters has been evaluated in Ref. [16] and
an excerpt is shown in Fig. 6 to Fig. 10. Based on the
discussed approach, it shows the individual gas property
ratio of the nozzle exit and the ambient flow of the
experiments in relation to Ariane 5-like conditions as
function of the simulation Mach number for different
expansion ratios. As baseline, the combustion chamber
of the wind tunnel model is set to 6.89 MPa (PennState),
a mixture ratio of 6and an expansion ratio of 58.5(red
line).
It shows that by varying the expansion ratio, the pres-
sure ratio (see Fig. 6) - the essential parameter for plume
formation (G-1/F-5) - can be kept similar over a wide
range without having a significant influence on the veloc-
ity ratio (Fig. 7). The ratios with respect to density ratio
0.5 0.6 0.7 0.8 0.9 1 1.1 1.2
0.85
0.9
0.95
1
1.05
1.1
Ma∞
(ve/v∞)Exp/(ve/v∞)Ariane
Figure 7. Velocity ratios between experiment and flight
over Mach number for various expansion ratios.
(Fig. 8), impulse density ratio (Fig. 9) and temperature
ratio (Fig. 10) can be kept in the same order. It seems
like an expansion ratio of ⇠20 can lead to comparable
results. As mentioned before, these parameters are used
to determine the operational conditions for the future fa-
cility. If similarity is required at a specific Mach number,
a detailed analysis with respect to the viscous behavior of
the plume is required.
The influence of the other ’screws’ is studied in Ref. [16].
It shows that a change of mixture ratio between O/F =
4to 8does not offer this wider range of comparable ra-
tios. Mainly the temperature ratio can be influence while
the other ratios stay in a very narrow regime. An ad-
justment of the pressure ratio is not possible this way
and flow separation in the nozzle has to be expected
(Ref. [14]). This can be avoided by increasing the com-
bustion chamber pressure to a most likely detrimental
level of 15 MPa, which is also advantageous for the den-
sity and impulse density ratio while keeping the temper-
ature and velocity ratio at a comparable level. Note that
mixture ratio ’screws’ is also changed without consider-
ing the effect on an existing set-up. The parameter vari-
ation shows that the most favorable parameter to adjust
is the expansion ratio since it can be changed without in-
creasing the loads on the combustion chamber.
4.2. Test Environment
The proposed concept shown in Fig. 11 is a combination
of the wind tunnel model used by Emunds and H. Riedel
[51] for base flow investigations and the PennState burner
[29] or slightly modified, but similar burner [62]. Bos
et al. [70] presented a comparable concept with respect
to the support in their feasibility study for a hot plume
test facility. Upstream of the nozzle exit, the wind tunnel
nozzle (1) is equipped with support arms (2), which have
three tasks: They simply keep the wind tunnel model in
place, the upper and lower support arm supply the com-
0.5 0.6 0.7 0.8 0.9 1 1.1 1.2
0
0.5
1
1.5
2
2.5
3
3.5
Ma∞
(ρe/ρ∞)Exp/(ρe/ρ∞)Ariane
ϵExp = 10
ϵExp = 20
ϵExp = 40
ϵExp = 58.5
Figure 8. Density ratios between experiment and flight
over Mach number for various expansion ratios.
0.5 0.6 0.7 0.8 0.9 1 1.1 1.2
0
0.5
1
1.5
2
2.5
3
Ma∞
(ρeve/ρ∞v∞)Exp/(ρeve/ρ∞v∞)Ariane
ϵExp = 10
ϵExp = 20
ϵExp = 40
ϵExp = 58.5
Figure 9. Impulse density ratios between experiment and
flight over Mach number for various expansion ratios.
0.5 0.6 0.7 0.8 0.9 1 1.1 1.2
1
1.1
1.2
1.3
1.4
1.5
1.6
1.7
Ma∞
(Te/T∞)Exp/(Te/T∞)Ariane
Figure 10. Temperature ratios between experiment and
flight over Mach number for various expansion ratios.
bustion chamber with gaseous hydrogen (3) and gaseous
oxygen (4), respectively, and a one or several hydrogen
supports can be selected as exit of the sensor harness.
The support arms converge in a central mounting (5) on
top of which is the combustion chamber (6). The injector
(7) and the nozzle (8) is exchangeable to realize various
injection conditions and nozzle exit conditions, respec-
tively. The wind tunnel nozzle is equipped with two lev-
els of straighteners (9) downstream of the support arms
to minimize perturbations. Fig. 2 shows the location of
the assembly with respect to the wind tunnel. This up-
stream supported concept is advantageous since stream-
lines of the upstream part of the wind tunnel model can
not be ’squeezed’, and the flow can not be accelerated to
Mach 1. Thus, blockage [71] is no issue and the diameter
can be freely selected. Further, the upstream support min-
imizes disturbance, which usually induced from a strut
support. On the other hand, Weiss and Deck [57] showed
that the shear layer of the wind tunnel model influences
the pressure fluctuations and has to be taken into account
by CFD and/or the diameter has to be sufficiently large
with respect to the model in order to allow the negligence
of these disturbance.
For upgrade step 1, the ratio between the wind tunnel
model and the model is /D = 340 mm/67 mm ⇡5.1.
For upgrade step 2 with the increased wind tunnel nozzle
(= 600 mm), it is at /D ⇡9.0. It is expected that
the influence of the shear layer perturbation decreases if
this ratio is about twice or four times as large as for the
investigations conducted by Ref. [57]. These calculations
were conducted with /D ⇡2.86.
The supply of the wind tunnel model is executed as
shown in Fig. 12. It shows the process flow diagram
of the GOX/GH2 supply from the storage outside of the
building to the combustion chamber in the test section in
accordance with the previously defined reference condi-
tions. It basically consists of a GOX line, a parallel GH2
line and a nitrogen line for purging different segments of
the supply system. The facility essentially runs in two
different modes: Ignition mode and operational mode. In
the ignition mode, the mass flow is passing through the
pressure reducing valves DR-X04/DR-H04, which limits
in combination with the setting for the control valves CV-
X07/CV-H07 the mass flow for a secure and ’low’ pres-
sure ignition. A flare system is used for ignition and for
burning excess-gas. Optionally, other system can be inte-
grated in the system as well. In the operational mode, the
mass flow passes through the shut-off valves AV-X06/AV-
H06. The mass flow sensor at FT-X15/FT-H15 sends the
mass flow rate to the flow indicator and controller FIC,
which adjusts the control valves CV-X07/CV-H07 to the
setpoint mass flow for the combustion chamber. Shut-
down off the combustion chamber is realized by purg-
ing the combustion chamber. In case of emergency, the
test section of the wind tunnel and the supply system is
vented with nitrogen. The extra loop in the oxygen piping
system prevents excessive and detrimental flow velocities
during start-up.
The reference conditions in Tab. 2 were used as basis to
1
2
7
3
4
5
6
8
9
Figure 11. Cross-sectional view of the wind tunnel model
concept.
specify the piping of the feeding system, which will obvi-
ously be assembled with standardized components. The
next standardized step to include RC 2 condition is at
160 MPa. This is the upper pressure limit for the feed-
ing system. The selected pipes are rated for maximum
volumetric flow rate of 0.0358 m3s1and 0.0029 m3s1
for hydrogen and oxygen, respectively. Consequently, the
maximum permitted mass flow is ⇠480 gs1for hydro-
gen and ⇠620 gs1for oxygen. In the invitation to ten-
der, it is specified that the facility is required to adjust the
mass flow to the nominal value with an accuracy of ±5%
within 3sfrom ignition mode and reaches an accuracy of
±1% within 6s. The hot plume testing facility offers a
large enough margin to RC 2 and a high degree of flexi-
bility is given to realize advanced experiments. Addition-
ally, the facility compares well to the facilities presented
in Tab. 1 and can be classified as intermediate-range test
bench.
In a typical test sequence, the combustion chamber is ig-
nited with a flare system (Igntion mode). After a stable
combustion is detected, the facility is switched to the op-
erational mode where the chamber pressure is increased
to an intermediate set point to establish choked flow con-
ditions in the model nozzle. The goal is to minimize
coupling effects induced by the start of the wind tunnel,
which is the next step. When the test conditions are set,
the chamber pressure is increased to the nominal value.
After the experiment, both systems are shut down simul-
taneously.
PT
X22
AV-H09
Combustion Chamber
Nitrogen Supply
Hydrogen Supply
Oxygen Supply
Emergency Flushing
N2
Storage
H2
Storage
300 bar
3x12x50l
O2
Storage
300 bar
1x12x50l
FIC
ST-N02
SV-N04
AV-N06
SV-N13
AV-N15
PI
H12
HV-H01 ST-H02
CK-H11
PT
H13
PT
X13
PI
X12
HV-X01 ST-X02
RO-X21
PT
X16
FT
X15
SV-X08
TT
X17
CV-X07
PI
N20
RO-N16
AV-H03
AV-X03
AV-X09
AV-X20
CK-X11
20.7...115 bar
36.7...397.4 g/s
-60…+34°C
25...340 bar
36.7...397.4 g/s
-10…+35°C
25...330 bar
39.9...66.2 g/s
-10…+35°C
PN400 PN160
PN250 PN10
AV-X06 AV-X10
PT
X14
TT
X19
PT
H16
FT
H15
SV-H08
TT
H17
CV-H07
AV-H06
PT
H14
PN400 PN160
DR-H04
DR-N12
Anteroom VMKForecourt Building 38 VMK
CK-N14
CK-N05
PN10 PN160
AV-H10
PT
X18
TT
H19
PT
H18
RO-N07
CK-N19AV-N18
20,7...115 bar
39.9...66.2 g/s
-10…+47°C
CK-N10AV-N11
RO-N08
RO-N17
AV-N09
SV-N24
AV-N25
To Vent
DR-X04
DR-N03
DR-N23
PT
N22
PT
N21
PT
N23
AV-N01
PI
N25
Figure 12. Process flow diagram for the GOX/GH2 supply.
Figure 13. Comparison of Mach number contours of fa-
cility (top) and free flight (bottom) case, detail of rocket
plume.
Figure 14. Comparison of Mach number contours of fa-
cility (top) and free flight (bottom) case.
4.3. Computational Fluid Dynamics
The wind tunnel concept presented in the previous sec-
tion is opposed to free flight conditions here. The Mach
number contours of the resulting flow fields are compared
in Fig. 13 and Fig. 14. Fig. 13 demonstrates how the
fundamentally different flow fields nonetheless lead to
practically indistinguishable flow conditions at the rocket
base. It can be seen in Fig. 14 that the base recirculation,
the nozzle separation, and the position of the Mach disc
are practically identical in both cases.
In order to allow a qualitative comparison, profiles are ex-
tracted in radial direction at three positions: at the wind
tunnel nozzle exit and at both 0.5·Rand Rbehind the
rocket nozzle, where R is the rocket nozzle radius. It can
be seen in Fig. 15 that the wind tunnel nozzle has been
well designed. The boundary layer profile stopping to
early is a data extraction artifact, the velocity reduces to
zero at a no-slip boundary. The boundary layer profiles
differ slightly. This might be due to the free flight case
having a homogeneous build-up in a constant pressure
environment, whereas the flow is accelerated through the
wind tunnel nozzle. The difference in velocity at the noz-
zle exit is negligible, the shear layer develops slightly dif-
ferently, possibly due to the difference in boundary layer
profile. Overall agreement between both cases is in terms
Figure 15. Radial profiles of axial velocity of free flight
(Free) and facility (Facility) case.
of density is satisfactory as well. This can be seen in
Fig. 16.
4.4. Measurement Techniques
For a coherent examination of the study, applicable mea-
surement techniques are discussed in a last step. Willert
et al. [72] studied relevant optical measurement tech-
niques, ergo non-intrusive image based methods. Un-
der consideration were flow field diagnostics like Parti-
cle Image Velocimetry (PIV), Doppler Global Velocime-
try (DGV) and optical diagnostic methods for reacting
flows like chemiluminescence, Laser Induced Floures-
cence (LIF), Filtered Rayleigh Scattering (FRS) and Co-
herent Anti-Stokes Ramann Scattering (CARS). As a re-
sult, a combined single line OH-LIF and PIV measure-
ment technique according to the work of Lange et al.
[73] was recommended, which offers correlated infor-
mation about temperature and velocity field with a high,
two-dimensional spatial resolution on a single shot ba-
sis. Additionally, FRS [74] is considered to be advan-
tageous as border crossing application for the tempera-
ture measurement in the recirculation region where the
OH concentration is unknown and temperatures in the
range of the ambient free stream are expected. OH-
LIF is restricted to an OH concentration in the order
of 0.1to 2⇥1016 molecules cm3[73]. Constraints
might arise for measurements close to the region with the
bright hydrogen-oxygen combustion and due to the lim-
ited knowledge about this technique in the high temper-
ature range [75]. The diagnostics can easily be compli-
mented with measurement techniques, which are read to
go at the department like high-speed Schlieren imaging,
IR-thermography, spectroscopy in the IR, UV and visi-
ble range, steady and unsteady pressure measurements et
cetera.
Figure 16. Radial profiles of density of free flight (Free)
and facility (Facility) case.
5. CONCLUSION
The study at hand presents how to extend the capabilities
of the existing wind tunnel Vertical Test Section Cologne
to simulate representative experiments for space trans-
portation systems in the transonic regime. An overview
was given to the issues and challenges like buffeting, drag
and heat loads for instance, and a conceptual solution was
suggested by integrating the PennState combustion cham-
ber in an upstream-supported wind tunnel model. The
wind tunnel model is supported upstream to minimize
strut disturbances and to transonic effects due to block-
age. For this exact configuration, it was shown the high-
est degree of similarity over a wide Mach number range
is most likely established by an adjustment of the expan-
sion ratio to about 20. A more profound study based on
the similarity rules of Ref.[43] for the viscous behavior
or with an axis-symmetric shear layer model that takes
compressibility, different gas compositions and tempera-
ture effects into account is necessary if investigations for
specific flow conditions are required. Proof of concept
was given by means of CFD simulations. The influence
of the wind tunnel nozzle on the mean base flow is neg-
ligible. Further investigations are recommended to eval-
uate the influence of the wind tunnel nozzle shear layer
on the unsteady base flow phenomena for a larger as-
pect ratio as investigated by Ref. [57]. The requirements
of the feeding system were defined, which are set to a
maximum mass flow rate of ⇠480 gs1for hydrogen
and ⇠620 gs1for oxygen at a maximum pressure of
16 MPa. By means of a process flow diagram, the real-
ization of such a facility was discussed. The study con-
cludes with an introductory description of a typical test
sequence.
ACKNOWLEDGMENTS
This project is financially supported by the German Re-
search Foundation (Deutsche Forschungsgemeinschaft –
DFG) within the Transregional Collaborative Research
Centre 40 (Sonderforschungsbereich Transregio 40). The
help and advice of the technical staff of the Supersonic
and Hypersonic Technology Department in Cologne is
gratefully acknowledged.
REFERENCES
[1] Arianespace. Press conference to Flight 157 - Ari-
ane 5 ECA, Paris, January 7, 2003.
[2] ESA. Arianespace Flight 157 - In-
quiry Board Submits Findings. URL
http://www.esa.int/Our_Activities/
Launchers/Arianespace_Flight_157_
-_Inquiry_Board_submits_findings2.
[3] K. Triesch and E.-O. Krohn. Die Vertikale
Meßstrecke der DFVLR in K¨
oln-Porz (Stand
1986), DFVLR-Mitt. 86-22. Wissenschaftliches
Berichtswesen der DFVLR, ISSN 0176-7739, Post-
fach 906058, 5000 K¨
oln 90, 1986.
[4] Vertical Test Section Cologne (VMK), Su-
personic and Hypersonic Technology Depart-
ment. URL http://www.dlr.de/as/en/
desktopdefault.aspx/tabid-194/407_
read-5445/.
[5] P. Leroux. Ariane 5 - Data relating to Flight 202.
Launch kit, May 2011.
[6] Ariane 5 User’s Manual. Number No. 3. Arianes-
pace, 2000. URL http://books.google.
de/books?id=CbA0mwEACAAJ.
[7] S. David and S. Radulovic. Prediction of Buffet
Loads on the Ariane 5 Afterbody. In In 6th Sympo-
sium on Launcher Technologies, Munich, Germany,
8-11 November 2005, 2005.
[8] H. Wong, J. Meijer, and R. Schwane. Experimental
and Theoretical Investigation of Base-Flow Buffet-
ing on Ariane5 Launch Vehicles. Journal of propul-
sion and power, 23(1):116–122, 2007.
[9] J.J. Meijer and C.M. van Beek. Analysis of Ari-
ane 5 Base Flow Measurements in the NLR-PHST
and FFA T1500 Wind Tunnels. National Aerospace
Laboratory, NLR-CR-99449, 1999.
[10] H. Wong and R. Schwane. Final Report on Numeri-
cal Support to Aerodynamics Buffeting Problems in
Ariane5. European Space Technology and Research
Centre, Doc. YPA/2236/HW, 1997.
[11] J. Muylaert, W. Kordulla, D. Giordano, L. Mar-
raffa, and R. Schwane. Aerothermodynamic Analy-
sis of Space-Vehicle Phenomena. ESA bulletin 105,
Aerothermodynamic Analysis, 2001.
[12] H.V. Fuchs, E. Mercker, and U. Michel. Large-
Scale Coherent Structures in the Wake of Axisym-
metric Bodies. Journal of Fluid Mechanics, 93(01):
185–207, 1979.
[13] H. Wong. A Brief Overview on some Resonance
Phenomena in Nozzle Flows, AIAA 2004-4013. In
40th AIAA/ASME/SAE/ASEE Joint Propulsion Con-
ference and Exhibit, 11 - 14 July 2004, Fort Laud-
erdale, Florida, 2004.
[14] M. Frey. Behandlung von Str¨
omungsproblemen
in Raketend¨
usen bei ¨
Uberexpansion. PhD
thesis, Universit¨
at Stuttgart, Holzgarten-
str. 16, 70174 Stuttgart, 2001. URL
http://elib.uni-stuttgart.de/opus/
volltexte/2001/800.
[15] S. Deck and P. Thorigny. Unsteadiness of an Ax-
isymmetric Separating-Reattaching Flow: Numeri-
cal Investigation. Physics of Fluids, 19(6):065103,
2007.
[16] D. Saile. TRP on Hot Plume Testing Facilities for
ELV Propulsion Characterization, WP2100: Scope
of Operation of Wind Tunnel. Ref.: Proposal-No.
3005921, 2013.
[17] D. Depr´
es. Analyse Physique et Mod´
elisation des
Instationnarit´
es dans les ´
Ecoulements d’Arri`
ere-
Corps Transoniques. PhD thesis, Universit´
e de la
M´
editerran´
ee Aix-Marseille II, 2003.
[18] P.-E. Weiss, S. Deck, J.-C.Robinet, and P. Sagaut.
On the Dynamics of Axisymmetric Turbulent Sepa-
rating/Reattaching Flows. Physics of Fluids, 21(7):
075103, 2009.
[19] G.P. Sutton and O. Biblarz. Rocket Propulsion
Rlements, 7th ed. John Wiley & Sons, 2010.
[20] D. Depr´
es, P. Reijasse, and J.P. Dussauge. Analy-
sis of Unsteadiness in Afterbody Transonic Flows.
AIAA journal, 42(12):2541–2550, 2004.
[21] V. Zapryagaev, A.V. Lokotko, S.V. Nikiforov, A.A.
Pavlov, A.V. Tchernyshev, W.J. Bannink, H. Ottens,
and J. Muylaert. Experimental Investigation of Base
Pressure with Hot Supersonic Jet for External Su-
personic Flow. In Fourth Symposium on Aerother-
modynamics for Space Vehicles, volume 487, page
579, 2002.
[22] B.H. Goethert. Base Heating Problems of Missiles
and Space Vehicles. ARS Paper, pages 1666–61,
1961.
[23] F.S. Simmons. Rocket Exhaust Plume Phenomenol-
ogy. Aerospace Press El Segundo, CA, 2000.
[24] S. M. Dash. Rocket motor plume flowfields: Phe-
nomenology and Simulation. In In AGARD, Rocket
Motor Plume Technology, ARARD-LS-188, ISBN
92-835-0713-4, volume 1, 1993.
[25] J.S. Draper. Plume Separated Region as a Flame-
holder. AIAA Journal, 14(3):358–362, 1976.
[26] K. Kinefuchi, K. Okita, I. Funaki, and T. Abe. Pre-
diction of In-flight Radio Frequency Attenuation by
a Rocket Plume by Applying CFD/FDTD Coupling.
49th AIAA/ASME/SAE/ASEE Joint Propulsion Con-
ference, July 14 - 17, 2013, San Jose, CA, AIAA
2013-3790, 2013.
[27] NASA. NASA Thermal Infrared Camperas Cap-
ture SpaceX Falcon 9 First Stage Re-entry, Sept.
21, 2014, 2014. URL https://www.youtube.
com/watch?v=riU3DZmU-jE.
[28] Unitary Plan Wind Tunnel NASA Langley. En-
try, Descent and Landing Project, Supersonic
Retropropulsion Test, July 2010, 2010. URL
https://www.youtube.com/watch?v=
i-coJg_vgxI.
[29] W.M. Marshall, S. Pal, R.D. Woodward, and R.J.
Santoro. Benchmark Wall Heat Flux Data for a
GO2/GH2 Single Element Combustor. AIAA Paper,
3572:2005, 2005.
[30] P. Sibtosh, W. Marshall, R. Woodward, and R.J.
Santoro. Test Case RCM-1: Penn State Preburner
Combustor. In 3rd International Workshop Rocket
Combustion Modeling, March 13-15, 2006, Vernon,
France, pages 1–9, 2006.
[31] Editor: S. Schlechtriem. Institute of Space Propul-
sion Lampoldshausen, Status Report 2011, 2011.
[32] C. Manfletti, J. Sender, and M. Oschwald. Theoret-
ical and Experimental Discourse on Laser Ignition
in Liquid Rocket Engines. 2009.
[33] O.J. Haidn and M. Habiballah. Research on
High Pressure Cryogenic Combustion. Aerospace
Science and Technology, 7(6):473 – 491, 2003.
ISSN 1270-9638. doi: http://dx.doi.org/10.
1016/S1270-9638(03)00052-X. URL http:
//www.sciencedirect.com/science/
article/pii/S127096380300052X.
[34] S. Magniant and P. Vinet and R. Blasi and J.-P.
Dutheil and Taiichi Motomura. Airbus Defence
and Space LOX/Methane Propulsion Demonstrator,
IAC-14.C4.5.1. In 65th International Astronautical
Congress, Toronto, Canada, 2014.
[35] G. Le Forestier, P. Danous, G. Dantu, A. Preuss,
F. Grauer, and R. Strunz. First Firing Test Cam-
paign of European Staged-Combustion Demonstra-
tion. In 3rd EUCASS conference, Versailles, 6th-9th
July, 2009.
[36] I.A. Klepikov, B.I. Katorgin, and V.K. Chvanov.
The New Generation of Rocket Engines, Operat-
ing by Ecologically Safe Propellant “Liquid Oxy-
gen and Liquefied Natural Gas (Methane)”. Acta
Astronautica, 41(4):209–217, 1997.
[37] C. Bonhomme, M. Theron, E. Louaas, A. Beaurain,
and E.P. Seleznev. French/Russian Activities on
LOx/LCH4 Area. In 57th Astronautical Congress
(International). Valence, Spain, 2006.
[38] J. Breteau, Th. Kachler, B. Vieille, R. Tontini,
R. Strunz, and G. Dantu. FLPP European Next
Generation Launcher Propulsion High Thrust En-
gine Demonstrator. In 3rd EUCASS Conference,
Versailles, 6th-9th July, 2009.
[39] Y. Letourneur, F. Leleu, D. Pinard, J. Krueger, and
M. Balduccini. Status of next generation expend-
able launchers concepts within the flpp program.
Acta Astronautica, 66(9):1404–1411, 2010.
[40] P. Caisso, A. Souchier, C. Rothmund, P. Alliot,
C. Bonhomme, W. Zinner, R. Parsley, T. Neill,
S. Forde, R. Starke, et al. A Liquid Propulsion
Panorama. Acta Astronautica, 65(11):1723–1737,
2009.
[41] K.-H. Kim and D.-S. Ju. Development of CHASE-
10 Liquid Rocket Engine Having 10tf Thrust Using
LOX & LNG (Methane). AIAA Paper, 4907:2006,
2006.
[42] H. Tamura, F. Ono, A. Kumakawa, and N. Yat-
suyanagi. LOX/Methane Staged Combustion
Rocket Combustor Investigation. Rep./AIAA, 1987.
[43] B.H. Goethert and L.T. Barnes. Some Studies of the
Flow Pattern at the Base of Missiles with Rocket
Exhaust Jets. Technical report, DTIC Document,
1960.
[44] B.H. Goethert. Base Flow Characteristics of Mis-
siles with Cluster-Rocket Exhausts. Aerospace En-
gineering, 20(3):28, 1961.
[45] M. Frey. TRP on Buffeting Reduction, WP3200:
Feasibility Study for the Plume Simulation. Doc.
No. TP24-TN122/2009, 1(1):19, 2009.
[46] S. Stephan and R. Radespiel and R. M ¨
uller-Eigner.
Jet Simulation Facility using the Ludwieg Tube
Principle. In 5th European Conference for Aeronau-
tics and Space Sciences (EUCASS), Munich, Ger-
many, 2013, 2013.
[47] B. Stahl, H. Esch, and A. G ¨
ulhan. Experimental In-
vestigation of Side Jet Interaction with a Supersonic
Cross Flow. Aerospace Science and Technology, 12
(4):269 – 275, 2008. ISSN 1270-9638. doi: http://
dx.doi.org/10.1016/j.ast.2007.01.009. URL http:
//www.sciencedirect.com/science/
article/pii/S1270963807001010.
[48] Bernhard Stahl, Hans Emunds, and Ali G¨
ulhan.
Experimental Investigation of Hot and Cold Side
Jet Interaction with a Supersonic Cross-Flow.
Aerospace Science and Technology, 13(8):488
– 496, 2009. ISSN 1270-9638. doi: http://dx.
doi.org/10.1016/j.ast.2009.08.002. URL http:
//www.sciencedirect.com/science/
article/pii/S1270963809000443.
[49] A. Preci and A. G ¨
ulhan. Free Flight Investiga-
tion of Atmospheric Entry Capsules in Low Sub-
sonic Flow. In In Proceedings of AIAA Atmospheric
Flight Mechanics Conference, 5-9 January 2015,
Kissimmee, Florida, 2015.
[50] A. Preci, A. G¨
ulhan, E. Clopeau, P. Tran, L. Fer-
racina, and L. Marraffa. Experimental Investigation
of Dynamic Characteristics for MarcoPolo-R ERC
in Low Subsonic Flow Using the Free Flight Tech-
nique. In to be presented at 8th European Sympo-
sium on Aerothermodynamics for Space Vehicles, 2-
6 March 2015, Lisbon, Portugal, 2015.
[51] H. Emunds and edited by A. Ferri H. Riedel. Con-
tribution of the Institute f¨
ur Angewandte Gasdy-
namik of the DFVLR, Porz-Wahn, In: Improved
Nozzle Testing Techniques in Transonic Flow: Re-
port of a Study organized by the AGARD Propul-
sion and Energetics Panel with the Collaboration of
the Fluid Dynamics Panel and including Papers and
Discussions from the Special Session on this Topic
held at the 35th Fluid Dynamics Panel Meeting in
Rome, Italy, 4-10 Sept. 1974. Number 208. North
Atlantic Treaty Organization, Advisory Group for
Aerospace Research and Development, 1975.
[52] P.K. Tucker, S. Menon, C.L. Merkle, J.C. Oefelein,
and V. Yang. Validation of High-Fidelity CFD Sim-
ulations for Rocket Injector Design. AIAA Paper,
5226:2008, 2008.
[53] J. L. Herrin and J. C. Dutton. Supersonic Base Flow
Experiments in the Near Wake of a Cylindrical Af-
terbody. AIAA journal, 32(1):77–83, 1994.
[54] C. J. Bourdon and J. C. Dutton. Planar Visual-
izations of Large-Scale Turbulent Structures in Ax-
isymmetric Supersonic Separated Flows. Physics of
Fluids (1994-present), 11(1):201–213, 1999.
[55] H. Esch. Die 0,6x0,6-m-Trisonische Meßstrecke
der DFVLR in K¨
oln-Porz (Stand 1986), DFVLR-
Mitt. 86-21. Wissenschaftliches Berichtswesen der
DFVLR, ISSN 0176-7739, Postfach 906058, 5000
K¨
oln 90, 1986.
[56] Trisonic Test Section Cologne (TMK), Su-
personic and Hypersonic Technology Depart-
ment. URL http://www.dlr.de/as/en/
desktopdefault.aspx/tabid-194/407_
read-5447/.
[57] P.E. Weiss and S. Deck. Numerical Investiga-
tion of the Robustness of an Axisymmetric Sepa-
rating/Reattaching Flow to an External Perturbation
Using ZDES. Flow, turbulence and combustion, 91
(3):697–715, 2013.
[58] A. Ponomarenko. RPA-Tool for Rocket Propulsion
Analysis. In Space Propulsion Conference, 2014.
[59] S.J. Isakowitz, J.P. Hopkins, and J.B. Hopkins. In-
ternational Reference Guide to Space Launch Sys-
tems. AIAA (American Institute of Aeronautics &
Astronautics), 1999.
[60] M. Sodomann. TOSCA Program Abilities,
Doc.No.: FF-SR 3000-0061-01, 1996.
[61] US NOAA and US Air Force. US Standard Atmo-
sphere, 1976. Technical report, NOAA-S/T, 1976.
[62] D. Saile and D. Kirchheck and A. G ¨
ulhan and C.
Serhan and V. Hannemann. Design of a GH2/GOX
Combustion Chamber for the Hot Plume Interaction
Experiments at DLR. In 8th European Symposium
on Aerothermodynamics for Space Vehicles, 2015.
[63] K. Hasegawa, K. Kusaka, A. Kumakawa, M. Sato,
M. Tadano, and H. Takahashi. Laser ignition char-
acteristics of GOX/GH2 and GOX/GCH4 propel-
lants. In Joint Propulsion Conference, 2003.
[64] V. Schmidt, D. Klimenko, O. Haidn, M. Oschwald,
A. Nicole, G. Ordonneau, and M. Habiballa. The-
oretical Investigation and Modeling of the Ignition
Transient of a Coaxial H2/O2-Injector. 2003.
[65] T. Gerhold, M. Galle, O. Friedrich, and J. Evans.
Calculation of Complex Three Dimensional Config-
urations Employing the DLR-TAU-code. In In Pro-
ceedings of the 35th Aerospace Sciences Meeting
and Exhibit, number AIAA-1997-167, Reno, USA,
1997.
[66] A. Mack and V. Hannemann. Validation of the Un-
structured DLR-TAU-Code for Hypersonic Flows.
In In Proceedings of the 32nd AIAA Fluid Dynamics
Conference and Exhibit, AIAA-2002-3111, 2002.
[67] S. Gordon and B.J. McBride. Computer Program
for Calculation of Complex Chemical Equilibrium
Compositions and Applications I. Analysis. NASA
RP-1311-P1, 1994.
[68] B.J. McBride and S. Gordon. Computer Program
for Calculation of Complex Chemical Equilibrium
Compositions and Applications II. User’s Man-
ual and Program Description. NASA RP-1311-P2,
1999.
[69] S.N.B. Murthy and J.R. Osborn. Base Flow Phe-
nomena with and without Injection - Experimental
Results, Theories, and Bibliography. Aerodynamics
of base combustion.(A 76-37230 18-02) New York,
American Institute of Aeronautics and Astronautics,
Inc., pages 7–210, 1976.
[70] O. Bos, O. de Bruijn, N.P. van Hinsberg, T.J.J. Lom-
baerts, P. Lucas, D.M. Luchtenburg, J.A.S. Wit-
teveen, and K.G. van der Zee. Feasibililty Study
of a Hot Plume Test Facility. 2002.
[71] B.H. Goethert. Transonic Wind Tunnel Testing.
Technical report, DTIC Document, 1961.
[72] C. Willert, M. Fischer, J. Heinze, and M. Vo-
ges. TRP on Hot Plume Testing Facilities for ELV
Propulsion Characterization, WP2300: Measure-
ment Techniques. Ref.: Proposal-No. 3005921,
2013.
[73] L. Lange, J. Heinze, M. Schroll, C. Willert, and
T. Behrendt. Combination of Planar Laser Opti-
cal Measurement Techniques for the Investigation
of Pre-mixed Lean Combustion. 2012.
[74] U. Doll, M. Fischer, G. Stockhausen, and C.E.
Willert. Fundamentals of Filtered Rayleigh Scat-
tering. 2012.
[75] S.P. Kearney, S.J. Beresh, R.W. Schefer, and T.W.
Grasser. Filtered Rayleigh Scattering Diagnos-
tic for Multi-Parameter Thermal-Fluids Measure-
ments: LDRD Final Report. Sandia Report,
SAND2004-0158, 2004.
