Wind tunnels

ENGINEERING TOOLS, TECHNIQUES AND TABLES

WIND TUNNELS: AERODYNAMICS,
MODELS AND EXPERIMENTS

No part of this digital document may be reproduced, stored in a retrieval system or transmitted in any form or
by any means. The publisher has taken reasonable care in the preparation of this digital document, but makes no
expressed or implied warranty of any kind and assumes no responsibility for any errors or omissions. No
liability is assumed for incidental or consequential damages in connection with or arising out of information
contained herein. This digital document is sold with the clear understanding that the publisher is not engaged in
rendering legal, medical or any other professional services.


Additional books in this series can be found on Nova’s website under the Series tab.

Additional E-books in this series can be found on Nova’s website under the E-books tab.


WIND TUNNELS: AERODYNAMICS,
MODELS AND EXPERIMENTS

JUSTIN D. PEREIRA
EDITOR

Nova Science Publishers, Inc.
New York

Copyright © 2011 by Nova Science Publishers, Inc.

All rights reserved. No part of this book may be reproduced, stored in a retrieval system or
transmitted in any form or by any means: electronic, electrostatic, magnetic, tape, mechanical
photocopying, recording or otherwise without the written permission of the Publisher.

For permission to use material from this book please contact us:
Telephone 631-231-7269; Fax 631-231-8175
Web Site: http://www.novapublishers.com

NOTICE TO THE READER
The Publisher has taken reasonable care in the preparation of this book, but makes no expressed or
implied warranty of any kind and assumes no responsibility for any errors or omissions. No
liability is assumed for incidental or consequential damages in connection with or arising out of
information contained in this book. The Publisher shall not be liable for any special,
consequential, or exemplary damages resulting, in whole or in part, from the readers’ use of, or
reliance upon, this material. Any parts of this book based on government reports are so indicated
and copyright is claimed for those parts to the extent applicable to compilations of such works.

Independent verification should be sought for any data, advice or recommendations contained in
this book. In addition, no responsibility is assumed by the publisher for any injury and/or damage
to persons or property arising from any methods, products, instructions, ideas or otherwise
contained in this publication.

This publication is designed to provide accurate and authoritative information with regard to the
subject matter covered herein. It is sold with the clear understanding that the Publisher is not
engaged in rendering legal or any other professional services. If legal or any other expert
assistance is required, the services of a competent person should be sought. FROM A
DECLARATION OF PARTICIPANTS JOINTLY ADOPTED BY A COMMITTEE OF THE
AMERICAN BAR ASSOCIATION AND A COMMITTEE OF PUBLISHERS.

Additional color graphics may be available in the e-book version of this book.

LIBRARY OF CONGRESS CATALOGING-IN-PUBLICATION DATA
Wind tunnels : aerodynamics, models, and experiments / editors, Justin D.
Pereira.
p. cm.
Includes index.
ISBN 978-1-61942-329-9 (eBook)
1. Wind tunnels. I. Pereira, Justin D.
TL567.W5W58 2011
629.134'52--dc22
2010047058

Published by Nova Science Publishers, Inc. † New York

CONTENTS

Preface vii
Chapter 1 Design, Execution and Numerical Rebuilding of Shock Wave
Boundary Layer Interaction Experiment in a Plasma Wind Tunnel 1
M. Di Clemente, E. Trifoni, A. Martucci, S. Di Benedetto
and M. Marini
Chapter 2 The Mainz Vertical Wind Tunnel Facility– A Review of 25 Years
of Laboratory Experiments on Cloud Physics and Chemistry 69
Karoline Diehl, Subir K. Mitra, Miklós Szakáll,
Nadine von Blohn, Stephan Borrmann and Hans R. Pruppacher
Chapter 3 Modeling and Experimental Study of Variation of Droplet Cloud
Characteristics in a Low-Speed Horizontal Icing Wind Tunnel 93
László E. Kollár and Masoud Farzaneh
Chapter 4 An Air-Conditioned Wind Tunnel Environment for the Study of
Mass and Heat Flux Due to Condensation of Humid Air 129
Akhilesh Tiwari, Pascal Lafon, Alain Kondjoyan and
Jean-Pierre Fontaine
Chapter 5 In-Situ Evaluation for Drag Coefficients of Tree Crowns 147
Akio Koizumi
Chapter 6 The Pre-X Lifting Body Computational Fluid Dynamics and Wind
Tunnel Test Campaign 167
Paolo Baiocco, Sylvain Guedron, Jean Oswald, Marc Dormieux,
Emmanuel Cosson, Jean-Pierre Tribot and Alain Bugeau
Chapter 7 Low-Speed Wind Tunnel: Design and Build 189
S. Brusca, R. Lanzafame and M. Messina
Index 221

PREFACE

This new book presents current research in the study of wind tunnels, including the
design, execution and numerical rebuilding of a plasma wind tunnel with the aim to analyze
shock wave boundary layer interaction phenomena; the Mainz vertical wind tunnel facility
experimenting on cloud physics and chemistry; an air-conditioned wind tunnel environment
for the study of mass and heat flux; using wind tunnel studies to evaluate the drag coefficient
of the tree crown and Pre-X aerodynamic/aerothermal characterization through computational
fluid dynamics and wind tunnels.
Chapter 1 - The present chapter reports the design, execution and numerical rebuilding of
a plasma wind tunnel experimental campaign with the aim to analyse shock wave boundary
layer interaction phenomena in high enthalpy conditions.
This particular flow pattern could arise in proximity of a deflected control surface, thus
generally causing a separation of the boundary layer and a loss of efficiency of the control
surface itself; moreover, high mechanical and thermal loads are generally induced at the flow
reattachment over the flap. Therefore, the analysis of this problem is crucial for the design
and development of the class of hypersonic re-entry vehicles, considering that, even though it
has been widely analyzed in the past, both from an experimental and theoretical point of
view, by describing its physical features, only few studies have been carried to analyse the
phenomenon in high enthalpy real gas and reacting flow conditions.
The activity has been developed by analysing the flow phenomenon of interest in
different conditions: i) hypersonic re-entry conditions considering the ESA EXPERT capsule
as a workbench, and ii) ground-based facility conditions considering the CIRA Plasma Wind
Tunnel “Scirocco”. The aim has been the correlation of the results predicted, by means of a
CFD code, and then measured through specific experiments suitably designed, in these two
different environments.
To this effect, a flight experiment has been designed to be flown on the EXPERT capsule
along the re-entry trajectory in order to collect flight data (pressure, temperature and heat
flux) on the shock wave boundary layer interaction phenomenon to be used for CFD
validation and, additionally, as a reference point for the extrapolation-from-flight
methodology developed accordingly. Requirements for the experimental campaign to be
performed in the “Scirocco” facility have been derived considering the most critical and
interesting points along the EXPERT trajectory. A suitable model, representative of the
EXPERT geometry in the zone of interest, i.e. the flap region, has been conceived by
defining the main design parameters (nose radius, length, width, flap deflection angle) and an

viii Justin D. Pereira

experimental campaign has been delineated, the aim being to reproduce on this model the
same mechanical and thermal loads experienced ahead and over the EXPERT full- scale flap
during the re-entry trajectory. Suitable facility operating conditions have been determined
through the developed extrapolation-from-flight methodology; the design and the analysis of
shock wave boundary layer interaction phenomenon has been done by focusing the attention
mainly to the catalytic effects over the interaction induced by the different behaviour in terms
of recombination coefficient of the materials involved in the problem under investigation.
Once defined the design loads, the model has been realized and tested in the Plasma
Wind Tunnel Scirocco under the selected conditions. The numerical rebuilding, showing a
reasonable good level of reproduction, has been also carried out, even though the validation
of the entire extrapolation-from-flight and to-flight developed methodology could be
completed only after the EXPERT flight currently planned in mid 2011.
Chapter 2 - The Mainz vertical wind tunnel is so far a worldwide unique facility to
investigate cloud and precipitation elements under conditions close to the real atmosphere.
Hydrometeors such as water drops, ice crystals, snow flakes, and graupels are freely
suspended at their terminal velocities in a vertical air stream under controlled conditions
regarding temperature (between -30°C and +30°C), humidity (up to the level of water
saturation), and laminarity (with a residual turbulence level below 0.5%) of the air stream.
Cloud processes in warm, cold, and mixed phase clouds have been investigated in the fields
of cloud physics and chemistry, aerosol–cloud interactions, and the influence of turbulence.
The experiments include the behaviour of cloud and rain drops, ice and snow crystals, snow
flakes, graupel grains and hail stones and the simulation of basic cloud processes such as
collisional growth, scavenging, heterogeneous drop freezing, riming, and drop-to-particle
conversion. Atmospheric processes have been investigated under both laminar and turbulent
conditions in order to understand and quantify the influence of turbulence.
The results are essential for applications in cloud chemistry models to estimate the
atmospheric pathway of trace gases, in cloud and precipitation models to improve the
description of the formation of precipitation (growth and melting rates), and in now- and
forecasting of precipitation to improve the evaluation of radar and satellite data.
Chapter 3 - Variation of the characteristics of aerosol clouds created in icing wind tunnels
is studied theoretically and experimentally. The characteristics of interest are the droplet size
distribution, liquid water content, temperature, velocity, and air humidity, which are among
the most important factors affecting atmospheric icing. Several processes influence the
trajectory, velocity, size and temperature of the droplets, such as collision, evaporation and
cooling, gravitational settling, and turbulent dispersion. The authors have developed a two-
dimensional theoretical model that takes these processes into account, and predicts how they
influence the changes in the characteristics of the droplet cloud during its movement in the
tunnel. The most recent development pays special attention to two of the possible collision
outcomes, i.e. coalescence after minor deformation and bounce, together with the transition
between them. Indeed, these outcomes are frequent when the relative velocity of the droplets
is small, as is the case for a cloud formed after the injection of water droplets in the direction
of air flow. An experimental study is also carried out with different thermodynamic
parameters at different positions in the test section of the tunnel, which makes it possible to
observe the evolution of cloud characteristics under different ambient conditions. The droplet
size distribution and liquid water content of the aerosol clouds were measured using an
integrated system for icing studies, which comprises two probes for droplet size

Preface ix

measurements and a hotwire liquid water content sensor. Droplet trajectories were observed
using particle image velocimetry. The experimental results are also used to validate the model
by comparing them to model predictions. Satisfactory agreement between the experimental
and calculated results establishes the applicability of the model to determine the evolution of
droplet size distribution and liquid water content in an aerosol cloud in the streamwise
direction, together with their vertical variation.
Chapter 4 - The development of an artificial ecosystem inside a closed environment is
one of the future challenging problems, which is mandatory for the long duration manned
space missions like lunar base or mission to Mars. Plants will be essential companion life
forms for such space missions, where human habitats must mimic the cycles of life on earth
to generate and recycle food, oxygen and water. Thus the optimized growth of higher plants
inside the closed environment is required to obtain efficient biological life support systems.
The stability and success of such systems lie on the control of the hydrodynamics and on an
accurate characterisation of the coupled heat and mass transfer that develop at interfaces
(solids, plants,..) within the space habitat. However, very few data can be found on the precise
characterization / prediction of the mass transfer at interfaces, and more particularly in space.
In most studies the mass flux is deduced from the measured / calculated heat flux by a heat
and mass transfer analogy.
Hence, the authors have developed a ground based experimental set-up to measure the air
flow velocities and concomitant mass transfer on specific geometries under controlled air
flow conditions (flow regime, hygrometry, temperature). The final goal is to derive a
theoretical model that could help for the prediction of the hydrodynamics and coupled
heat/mass transfer on earth, and eventually in reduced gravity. The authors have used a
closed-circuit wind tunnel for our experiments, which can produce very laminar to turbulent
flows with controlled temperature and hygrometric parameters inside the test cell. The initial
experiments have been performed in dry air with an average velocity between 0.5-2.5 m.s-1.
The velocity profiles near a clean aluminium flat plate in horizontal or vertical positions have
been studied for low Reynolds number flows by hot wire anemometry. The measurements
with the horizontal plate showed a boundary layer thickness in agreement with the Blasius’
solutions. Condensation of humid air was induced on an isothermal flat plate, which was
cooled by thermoelectricity. The mass transfer on the plate was controlled and recorded with
a precise balance. The obtained results are analyzed, and compared to the available data on
condensation.
Chapter 5 - In order to make a hazard prediction of trees against wind damage, such as
stem breakage or uprooting, it is essential to quantitatively estimate the wind force acting on a
tree. The drag coefficient of the tree crown, which is necessary to estimate wind force, has
been evaluated using wind tunnel studies. Most of the specimens used for wind tunnel studies
were dwarf trees, because of the restrictions due to wind tunnel size. However, with regard to
the wind-force response, the similarity rule is not applicable to the relationship between dwarf
trees and actual-sized trees. In fact, the drag coefficients of small trees were found to be
considerably greater than those of actual-sized trees. To estimate the drag coefficients of
actual-sized trees accurately and easily, a field test method was developed. Using this method,
wind speed and stem deflection were monitored simultaneously. The wind force acting on the
tree crown was calculated from the stem deflection; the stem stiffness was evaluated by
conducting tree-bending tests. The field tests were conducted on black poplars and a Norway
maple; the results showed that the drag coefficients decreased with an increase in wind speed.

x Justin D. Pereira

This decrease can be explained mainly by the decrease in the projected area of the crown,
because of the swaying movement of the leaves and branches. Although the variation in the
drag coefficients was large at low wind speeds because of the swaying behavior of the stem
subjected to a variable wind force, the variation at wind speeds above 10 m/s was small. The
average drag coefficient for black poplars at a wind speed of 30 m/s was estimated by the
curve fitting of a power function to the wind velocity-drag coefficient relationship, and this
value was found to be not greater than that of actual-sized conifers previously studied in wind
tunnel experiments. These results suggest that the wind permeability of poplar crowns is
greater than that of conifer crowns due to the difference in leaf flexibility. Although the drag
coefficients in the defoliation season were smaller than those measured in the leaved season at
low wind speeds, the difference in drag coefficients became less pronounced at high wind
speeds.
Chapter 6 - Pre-X was the CNES proposal for demonstrating the maturity of European
technology for gliding re-entry spacecraft. The program finished in year 2007 with the end of
the phase B and a successful PDR. Then it was stopped with the aim of joining the ESA
project IXV.
The main goal of this experience is to demonstrate the implementation of reusable
thermal protections, perform aero thermo dynamics experiments and efficiency of a suitable
guidance navigation and control system. The attitude control is realised by elevons and
reaction thrusters overall the hypersonic flight, with a functional and experimental objective.
This paper presents the Pre-X aerodynamic / aerothermal characterisation through
computational fluid dynamics and wind tunnels tests performed during the phases A and B of
the programme. The tests permitted to cover the Mach range from 0.8 to 14 and to investigate
the main effects of aerodynamic and aerothermal phenomena. In the preceding phases the
aerodynamic shape and centring had been defined.
The logic and main results of this activity are presented in this paper.
Chapter 7 - In this chapter the authors deal with a procedure for the design and build of a
low speed wind tunnel for airfoil aerodynamic analyses and micro wind turbine studies.
The designed closed-circuit wind tunnel has a test chamber with a square cross section
(500 mm x 500 mm) with a design average flow velocity of about 30 m/s along its axis.
The designed wind tunnel has a square test chamber, two diffusers (one adjacent to the
test section and one adjacent to the fan to slow the flow), four corners (with turning vanes) to
guide the flow around the 90° corners, an axial fan to guarantee the mass flow rate and
balance any pressure loss throughout the circuit, a settling chamber with a honeycomb (to
eliminate any transverse flow), a series of ever-finer mesh screens (to reduce turbulence) and
a nozzle to accelerate flow and provide constant velocity over the whole test chamber. The
pressure losses of single components were evaluated as well as the global pressure loss (the
sum of pressure losses of all the single components). Once the pressure losses were evaluated,
the axial fan was chosen to guarantee the design’s volumetric flow, balance pressure losses
and above all maximise its performance. The definitive dimensions of the wind tunnel are
10.49 m x 3.65 m.
Once the design targets were defined, the test chamber dimensions, maximum wind speed
and Reynolds numbers were calculated.
At the end of the design process, the wind tunnel energy consumption was estimated and
on-design and off-design performance was evaluated to obtain the wind tunnel circuit
characteristics for a defined velocity range (0 – 50 m/s).

Preface xi

The best circuit and axial fan matches were performed in both the open and closed test
section configurations. Using the matching procedure between the fan and wind tunnel’s
mechanical characteristics (global pressure loss as a function of wind velocity), the fan
operating parameters were set up for optimum energy conservation.

In: Wind Tunnels: Aerodynamics, Models and Experiments ISBN 978-1-61209-204-1
Editor: Justin D. Pereira © 2011 Nova Science Publishers, Inc.

Chapter 1

DESIGN, EXECUTION AND NUMERICAL REBUILDING
OF SHOCK WAVE BOUNDARY LAYER INTERACTION
EXPERIMENT IN A PLASMA WIND TUNNEL

M. Di Clemente1, E. Trifoni1, A. Martucci1, S. Di Benedetto1
and M. Marini1
1
Italian Aerospace Research Centre
Via Maiorise – 81043 Capua (CE), Italy

ABSTRACT
The present chapter reports the design, execution and numerical rebuilding of a
plasma wind tunnel experimental campaign with the aim to analyse shock wave boundary
layer interaction phenomena in high enthalpy conditions.
This particular flow pattern could arise in proximity of a deflected control surface,
thus generally causing a separation of the boundary layer and a loss of efficiency of the
control surface itself; moreover, high mechanical and thermal loads are generally induced
at the flow reattachment over the flap. Therefore, the analysis of this problem is crucial
for the design and development of the class of hypersonic re-entry vehicles, considering
that, even though it has been widely analyzed in the past, both from an experimental and
theoretical point of view, by describing its physical features, only few studies have been
carried to analyse the phenomenon in high enthalpy real gas and reacting flow conditions.
The activity has been developed by analysing the flow phenomenon of interest in
different conditions: i) hypersonic re-entry conditions considering the ESA EXPERT
capsule as a workbench, and ii) ground-based facility conditions considering the CIRA
Plasma Wind Tunnel “Scirocco”. The aim has been the correlation of the results
predicted, by means of a CFD code, and then measured through specific experiments
suitably designed, in these two different environments.
To this effect, a flight experiment has been designed to be flown on the EXPERT
capsule along the re-entry trajectory in order to collect flight data (pressure, temperature

1
Phone : +39 0823 623577, Fax : +39 0823 623700
Email: m.diclemente@cira.it, e.trifoni@cira.it, a.martucci@cira.it, s.dibenedetto@cira.it, m.marini@cira.it

2 M. Di Clemente, E. Trifoni, A. Martucci et al.

and heat flux) on the shock wave boundary layer interaction phenomenon to be used for
CFD validation and, additionally, as a reference point for the extrapolation-from-flight
methodology developed accordingly. Requirements for the experimental campaign to be
performed in the “Scirocco” facility have been derived considering the most critical and
interesting points along the EXPERT trajectory. A suitable model, representative of the
EXPERT geometry in the zone of interest, i.e. the flap region, has been conceived by
defining the main design parameters (nose radius, length, width, flap deflection angle)
and an experimental campaign has been delineated, the aim being to reproduce on this
model the same mechanical and thermal loads experienced ahead and over the EXPERT
full- scale flap during the re-entry trajectory. Suitable facility operating conditions have
been determined through the developed extrapolation-from-flight methodology; the
design and the analysis of shock wave boundary layer interaction phenomenon has been
done by focusing the attention mainly to the catalytic effects over the interaction induced
by the different behaviour in terms of recombination coefficient of the materials involved
in the problem under investigation.
Once defined the design loads, the model has been realized and tested in the Plasma
Wind Tunnel Scirocco under the selected conditions. The numerical rebuilding, showing
a reasonable good level of reproduction, has been also carried out, even though the
validation of the entire extrapolation-from-flight and to-flight developed methodology
could be completed only after the EXPERT flight currently planned in mid 2011.

1. INTRODUCTION
The high cost of access to space is the main limitation to scientific research and space
commercialization, and for this reason all the countries in Europe are thinking how design
advanced spacecrafts in order to achieve low launch costs in the near future (Ref. [4],[12]).
Spacecrafts like the US Space Shuttle Orbiter represent the first generation of reusable launch
systems but several system studies have been conducted during the 80’s to investigate
possible future concepts for the next generation of RLVs. In the frame of the ESA-FESTIP
Program in late 90’s, system concept studies were carried out and an extensive investigation
of a wide range of RLV concepts (more than 10 configurations) was performed (ref. [37]). In
the following decade several programmes, at European and national level, were launched to
promote the development of some of the identified enabling technologies required for the
future generation of reusable space transportation systems that shall be safer and less
expensive with respect to the US Space Shuttle. Enabling technologies for such vehicles and
derived systems must be inherently reliable, functionally redundant, wherever practical and
designed to minimize or eliminate catastrophic failure modes. Reliability could be improved
through performance margin that translates to robust design, and this presupposes the
maturation of some specific macro-technologies:

• Re-entry heating. the aerospace vehicles have to handle the typical large thermal
loads encountered during re-entry to Earth from LEO, due to the necessity of
reducing the vehicle speed before landing;
• Hypersonic flight navigation. the future space vehicles will have to fly for large part
of their mission to speed much greater that the speed of the sound, and will have to
maneuver safely in such conditions;

Design, Execution and Numerical Rebuilding of Shock Wave… 3

• Reusability. the most important characteristic from the operational point of view is
the tendency to be as much like current airplanes thus translating into the reusability
concept.

Starting from the necessity of a proper level of maturity of these high level technologies,
some guidelines and critical points to be developed at lower level, in order to match the
particular requirements for the RLV design, were identified by past space systems and
technological programs. Among the others, can be identified:

• Configuration Design
• Extrapolation to Flight
• Transition Prediction
• Control Surface Aerothermodynamics

It is clear that many other technological areas are being involved and ask for other
significant developments (propulsion, flight mechanics, stability and control, guide and
navigation, configuration optimization, etc.) but, in any case, to develop the future space
transportation system a considerable work should be devoted to the aeroshape definition in
order to improve performance, flyability and controllability, propulsion integration, heat load
reduction, stage separation, coupling between forebody aeroelasticity and propulsion system,
coupling between viscous drag and heat loads. The aerodynamic efficiency (E=CL/CD) should
be increased since they will experiment large part of flight at moderate altitude at high Mach
number, strongly asking for more efficient aerodynamic design. Also the transition process
from laminar to turbulent boundary layer should be predicted with greater accuracy since it
plays an important role in the design of aerospace planes thermal protection system, and the
currently available theoretical know-how (i.e. the stability theory) could not yet guarantee for
a safe and reliable transition prediction (Ref. [18], [19], [20]).
Among the others, the study of aerodynamic efficiency of control surfaces plays a role of
primary importance (Ref. [17]). In fact, the necessity of manoeuvrability and high cross-range
during ascent or re-entry phase requires the capacity to increase control surfaces aerodynamic
efficiency whose analysis is strictly connected to the study of shock wave boundary layer
interaction (SWBLI) occurring around them. The increase of knowledge must regard,
especially in the SWBLI phenomenon in high enthalpy conditions, the prediction, with a good
level of approximation, of its behaviour in flight conditions.
In a classical approach, the design of space vehicles (e.g. the Space Shuttle) is based
heavily upon experimental data although, due to the inherent limitations of similarity laws,
ground based facilities cannot simulate completely the physics of flows experienced by such
vehicles during re-entry. To overcome these limitations different strategies could be adopted:
in US data obtained from in-flight experiments, particularly with the X-series vehicles, have
been used to complement the test data obtained from ground-based facilities; on the other
hand, since the times of Hermes Program, Europe chose to complement the knowledge
available from the cold wind tunnels, which are not able to model the high-temperature and
real gas effects typical for higher speeds and altitudes, by means of high enthalpy or hot-flow
facilities. ESA, therefore, supported the updated of existing cold flow wind tunnels, and also
the construction of facilities with new capabilities, as for example the PWT Scirocco of


CIRA, to investigate the heat loads and gas surface interactions on materials and large size
structures (ref. [10], [38]).
In any case, the prediction of hypersonic flows, both for the complexity of the required
physical modelling and for the impossibility to duplicate in wind tunnels real flight conditions
due to the high energy required, is still one of the main problem related to the development of
the new class of space vehicles. Moreover, high efficiency space vehicles require complex
investigations because of the large contribution of the viscous effects to the aerodynamic
forces and heating, while the effects of the gas modelling are important since the small blunt
nose, necessary to increase the aerodynamic efficiency, does not shield the rest of the vehicle,
thus implying the presence of large chemical effects on most of the vehicle surface.
The main data sources for the aerothermodynamic design of a space vehicle are
computational fluid dynamics, wind tunnel tests and flight experiments, generally on
simplified geometries:

• wind tunnel tests are important because they allow carrying on “controlled”
simulations and therefore to better understanding the flow-physics phenomena;
although ground-based facilities provide fundamental information for flight, no one
facility can provide all of the aerothermodynamic information required for the design
of a vehicle. As today it is well recognised, duplication of all flight characteristic
parameters (Mach, Reynolds, Damkhöler, state of the gas) in a ground facility is not
possible, particularly flight Reynolds number and high enthalpy effects are critical
and difficult to be reproduced at ground;
• flight experiments data represent the “truth” to be predicted, i.e. they show the real
performance of the vehicle in representative conditions and, therefore, they are
unique for vehicle qualification although they are quite costly, require considerable
time and have uncertain repeatability and accuracy. Many phenomena can not be
directly measured and de-coupling of effects is not always an easy task;
• numerical simulations still play a fundamental role in the study of aerot-
hermodynamics; moreover, the highest confidence in any ground-based or flight data
set occurs when the results obtained with CFD are in agreement with them, the so-
called extrapolation-to-flight technique. Even if today CFD is contributing
significantly to the aerothermodynamic design of advanced vehicles, it still suffers
from lack of physical modelling, robustness and accuracy of the mathematical
algorithms, grid generation flexibility and hardware limitations; thus good wind-
tunnel and flight data are still necessary for validation and/or calibration of CFD
codes used to predict surface and flow field variables for the full-scale vehicle at re-
entry flight conditions.

The best approach for improving confidence in aerothermodynamic design tools, from a
computational and ground-based experimental point of view, is to validate those tools and
design approaches with respect to flight experiments. As matter of fact, although in the last
years Europe has dedicated significant effort to improve the quality and reliability of
aerothermodynamic predictions, due to their key importance in the design and development
of any hypersonic space vehicle, and a considerable effort has been devoted to the realization
of ground based plasma facilities and development of advanced numerical tools with the state


of the art physical model, in-flight experimentation is still needed to validate the
computational codes and to establish meaningful and reliable ground-to-flight extrapolation
methodology.
Above Mach 10, where in particular high-temperature effects become dominant, CFD
represents the only prediction tool, and therefore the appropriate validation of numerical
codes is a great concern. Generally it is achieved, by comparing data measured in high-
temperature facilities with those obtained by numerical prediction and in many cases a
numerical approach is used to define the experimental test cases and for the interpretation of
the measured data. CFD codes are subsequently being used for flight simulations above Mach
10 even if this ‘extrapolation method’ assumes that the physical models enabling good results
for the simulation of the experimental test cases, provide good results also for free flight.
Therefore, free-flight data are required to remove any doubt about the validity and accuracy
of the CFD predictions, and to confirm the extrapolation methodology as well. The main
argument for the in-flight experimentation is therefore the need for realistic and combined
loads levels which are representative for the operational environment of a RLV. Such tests
must be performed complementarily to on-ground testing for validating critical enabling
technologies of the reference RLV concept.
The analysis of this phenomenology is complicated by the fact that, in hypersonic regime,
scaling laws have not yet been found. Plasma Wind Tunnels, which allow the same energy
levels of the real flight, are in fact characterized by the test chamber flow rather dissociated
conditions, and this has a large influence on the flow-field around the test article, while cold
hypersonic wind tunnels, where the simulation is focused on the duplication of Mach and
Reynolds numbers, permit only to reproduce the classical aerodynamic forces and the related
coefficients even though with strong limitations. The influence of real gas effects and viscous
interaction effects on control flap efficiency and heating is one of the main
aerothermodynamic issues for the next generation RLV design, together with the qualification
testing of the thermal protection system in ground-based facility and the consequent
extrapolation to flight for experimental results.
In order to assess these issues, a numerical approach has been followed to define a wind-
tunnel experimental campaign on a representative model to reproduce the in-flight expected
values of mechanical and thermal loads acting on a typical control device, in interpreting the
measured data and finally for the extrapolation to the flight conditions of the experimental
results, as the local conditions in the wind tunnel facility only partially duplicate those in
flight. Moreover, a flight experiment whose results could be used as point of reference for
such phenomena has been also designed. As matter of fact, aerothermodynamic design issues,
as the analysis of flap efficiency for control and navigation, has been addressed using
advanced numerical codes, ground-based facilities and flight testing.
Following the previous considerations, it arises the need to develop an extrapolation-
from-flight and to-flight methodology able to combine and mutually validate the flight and
ground data on the problem of interest. Even though the prediction of mechanical and thermal
loads acting on the control surfaces of hypersonic vehicles is crucial for the design of their
aerodynamic shapes and thermal protection system, at the moment the lack of hypersonic
flight data that can serve as a point of reference for the validation process, makes it
impossible, especially for some of the most challenging hypersonic problems.


Figure 1. Extrapolation from flight and to flight procedure.

The issue of an extrapolation-to-flight methodology for high enthalpy flow must be in the
light of a progressive building up of confidence in the design of a space vehicle. The
development of such methodology, whose rationale is shown in Figure 1, has been carried out
referring to the ESA EXPERT capsule (ref. [40]), which has the indubitable advantage to be a
simple geometry, conceived as an experimental test-bed for in-flight experimentation and
designed to avoid degradation and flow contamination. In order to develop the methodology
and to extrapolate plasma facility results to real flight conditions, it is necessary, first of all, to
characterize by means of CFD simulations the flight conditions in the flap region and design a
flight experiment in order to instrument the vehicle and to collect flight data during the re-
entry mission, to be used for the post-flight analysis to validate the entire procedure. On the
other hand it is necessary to design, perform and numerically rebuild a number of
experimental tests in a plasma wind tunnel facility that can be representative of the flight
conditions with respect to the SWBLI phenomenon over the flap. Finally, plasma wind tunnel
results must be correlated, by means of the relevant parameters of the interaction as viscous
interaction parameter and rarefaction parameter, with those predicted (and then measured)
during the flight, the goal being to understand the test conditions necessary to reproduce
(simultaneously or separately) the mechanical (pressure) and thermal (heat flux, temperature)
loads acting on the control surface device.

1.1. EXPERT Capsule

The development of the extrapolation to flight methodology has been carried out
referring to the ESA EXPERT capsule whose in-flight test program focuses on a generic
capsule-like configuration designed in such a way to enhance the most interesting
aerothermodynamic phenomena of a typical re-entry vehicle performing a sub-orbital ballistic
hypersonic flight. The main objective of the project is to collect in-flight data on the most
critical aerothermodynamic phenomena via dedicated classical and advanced flight test
measurement assemblies (i.e. EXPERT Scientific Payloads), and this in order to improve the
knowledge about the differences between ground experiments and real flight conditions; each
particular phenomenon related with the high energy re-entry mission (gas-surface interaction,
induced and natural laminar-to-turbulence transition, real gas effects on shock wave boundary

Design, Execu
D ution and Num
merical Rebuilding of Shoc Wave…
ck 7

la
ayer interactio shock lay chemistry has been separately an
on, yer y) nalyzed, and a specific
xperiment has been design for each of them in t frame of the Technica Research
ex s ned the al
Prrogram related to the capsu developme The scient
d ule ent. tific data will then be used to validate
st
tate-of-the-art numerical to ools for aeroothermodynam application and groun
mic ns nd-to-flight
ex
xtrapolation pr rocedures (Re [39]).
ef.
Each Paylo will be q
oad qualified acco
ording to the relevant Ass
e sembly, Integgration and
Verification Plan; in para
V P allel a numb ber of exper rimental activ vities in the field of
e
ae
erothermodynamics are carr out to acq
ried quire all necessary pre-flight information on specific
ph
henomena allo owing for an optimized pos st-flight phase Among the others, specia attention
e. al
ha been given to the Shock Wave / Boun
as n k ndary Layer In nteraction phe enomenon wh hose effects
on the open flap are being in
n ps nvestigated w two differe Scientific Payloads, i.e. Payload 6,
with ent
th
hrough instrummentation of f flaps and cavi ities with mai inly classical sensors, and Payload 7,
th
hrough the cha aracterization of the bounda layer appr
ary roaching the f flap, whose deevelopment
ha been carried out in the fra of the pre
as d ame esent research activity.
The referen geometry of the EXPER capsule is a body of rev
nce RT s volution with an ellipse-
cl
lothoid-cone two-dimension longitudina profile cut b 4 planes an equipped w 4 fixed
nal al by nd with
op flaps. Th elliptical n
pen he nose has a ra adius of 600 mm at the s stagnation point and an
ec
ccentricity of 2.5. The fixed flaps have a deflection of 20 deg, a wid of 400 mm and an x-
d f dth m
ax projected length of 300 m (see Figur 2).
xis l mm re

Fi
igure 2. EXPER capsule.
RT

2.
2 MATHEM
MATICAL MODEL

The analys of shock wave bounda layer inte
sis ary eraction, for t developm
the ment of the
ex
xtrapolation-fr
rom-flight me
ethodology, ha been carried out consider
as d ring the CIRA numerical
A
co
ode H3NS which allows for the ae
w s erothermodyna amic analysis over comp
s plex three-
di
imensional ge eometries and suitable to s
d simulate commpressible flow at high ent
w thalpy (ref.
[2
28]).
One of the main charact personic flows is that, due to the high temperatures
teristics of hyp
ex
xperienced be ehind the bow shock, the gas cannot be considere as a perf
w e ed fect gas as
co
ommonly assu umed for low speed flows; air molecule at temperat
w ; es tures higher th 800 K
han
st to vibrate and for temp
tart e peratures arou 1500 K t dissociatio of oxygen molecules
und the on
be
egin whereas for higher tem mperature also nitrogen mo olecules disso
ociate. The mo odelling of
hi
igh temperatu phenome
ure ena is quite difficult bec cause of the difference a among the


characteristic time of fluid-dynamics and that of chemical reactions and vibration. This
situation generally leads to the thermochemical non-equilibrium. In fact, there are many
problems in high-speed gas dynamics where the gas doesn’t reach the equilibrium state; a
typical example is the flow across a shock wave, where the pressure and temperature are
rapidly increased within the shock front. By the time equilibrium properties have been
approached, the fluid element has moved a certain distance downstream of the shock front.
The modelling of these phenomena cannot be limited to a calculation of equilibrium
conditions at a certain temperature and pressure, but a number of equations must be added to
the classical Navier-Stokes formulation, one for each vibrating or dissociated species.
In the code considered for numerical computations, governing equations have been
discretized using a finite volume technique with a centred formulation over structured multi-
block meshes. This approach is particularly suitable to the integral form of the equations; in
fact, in a first order approximation, it is simply obtained by integrating the equations for each
cell and considering the variables constant inside each volume. The integral formulation
ensures that mass, momentum and energy are conserved at the discrete level. Suitable models
have to be taken into account to define the thermodynamics, the transport coefficients and
turbulent variables as reported in detail in Appendix 1. The inviscid fluxes at cell interfaces
are computed using a Finite Difference Splitting (FDS) Riemann solver, which is especially
suitable for high speed problems (Ref.[3]). This method solves for every mesh interval the
one-dimensional Riemann problem for discontinuous neighboring states (the states at both
sides of the cell face). The second order approximation for FDS is obtained by means of a
higher order ENO (Essentially Non Oscillatory) reconstruction of interface values. The
viscous fluxes are calculated by central differencing, i.e. computing the gradients of flow
variables at cell interfaces by means of Gauss theorem. Time integration is performed by
using an Euler forward scheme with a semi-implicit pre conditioner based on the derivative of
the source chemical and vibrational terms.

3. PWT SCIROCCO EXPERIMENT PRELIMINARY DESIGN
A number of experiments to be performed in the CIRA Plasma Wind Tunnel “Scirocco”,
representative of the capsule flight conditions with respect to the shock wave boundary layer
interaction phenomenon occurring around the 20 deg flap, has been designed: PWT driving
conditions, model configuration and attitude and model instrumentation have been defined,
by means of a massive CFD activity performed by using the CIRA code H3NS. These
experiments have been designed in order to allow for the duplication of characteristic
parameters (viscous interaction parameter, rarefaction parameter, reference pressure and heat
flux) of the interaction to reproduce on a full-scale flap model both pressure and heat flux
levels estimated in critical re-entry flight conditions. The final goal has been to develop an
extrapolation-to-flight methodology for such flows since the full duplication of flow
characteristic numbers (Mach, Reynolds, Damkhöler) and state of the gas is not feasible in
ground facilities.
A parametric analysis of the facility operating conditions and model characteristic
dimensions (nose radius, length, flap dimensions, etc.) has been carried out in order to define
the operating conditions and experimental set up that permit a simultaneous reproduction of


mechanical and thermal loads acting on flap in flight conditions over the selected model. The
rest of the activity has been devoted mainly to the choice of the different protection materials,
its equipment (sensors distribution) and the experimental tests at the selected flow conditions.
The final model configuration reproduces the full-scale EXPERT 20 deg flap, mounted
on a holder composed by a flat plate with a rounded leading edge (made of copper and
actively cooled) and lateral edges. The flap has been realized in C-SiC, which is the same
material foreseen for the realization of the EXPERT capsule flap, whereas for the flat plate it
has been used Haynes 25 whose main thermo-mechanical characteristics are quite similar to
PM1000 which is the material foreseen on the capsule. The effects of catalysis jump, due to
the coupling of different materials, have been analysed by considering the available
recombination coefficients for the materials of interest, or modelling the different parts as
fully or not catalytic.

3.1. PWT “Scirocco” Facility Description

The CIRA Plasma Wind Tunnel “Scirocco” is devoted to aerothermodynamic tests on
components of aerospace vehicles; its primary mission is to simulate (in full scale) the
thermo-fluid-dynamic conditions suffered by the Thermal Protection System (TPS) of space
vehicles re-entering the Earth atmosphere.
“Scirocco” is a very large size facility, whose hypersonic jet has a diameter size up to 2 m
when impacts the test article and reaches Mach number values up to 11. The jet is then
collected by a long diffuser (50 m) and cooled by an heat exchanger. Seventy MW electrical
power is used to heat the compressed air that expands along a convergent-divergent conical
nozzle. Four different nozzle exit diameters are available: 0.9, 1.15, 1.35 and 1.95 m,
respectively named C, D, E and F. The overall performance of “Scirocco” in terms of
reservoir conditions is the following: total pressure (P0) varies from 1 to 17 bar and total
enthalpy (H0) varies from 2.5 to 45 MJ/kg. Facility theoretical performance map in terms of
reservoir conditions produced by the arc heater is shown in Figure 3. Lower enthalpy values
are obtained by using a plenum chamber between the arc heater column exit and the nozzle
inlet convergent part, which allows transverse injection of high pressure ambient air to reduce
the flow total enthalpy.

Figure 3. Arc heater theoretical performance map.


The energetic heart of the facility is the segmented constricted arc heater, a column with a
maximum length of 5.5 m and a bore diameter of 0.11 m. At the extremities of this column
there are the cathode and the anode between which the electrical arc is generated. A power
supply feeds the electrical DC power to the electrodes for the discharge. A compressed air
supply distributes dry compressed air to the various segments of the arc heater column, being
able to supply a mass flow rate ranging from 0.1 to 3.5 kg/s, heated up to 10000 K.
The last important subsystem of “Scirocco” is the vacuum system, which generates the
vacuum conditions in test chamber required by each test. The system consists of ejectors that
make use of high pressure water steam as motor fluid (30 bar and 250 °C). The achievement
of the operating conditions (P0, H0) in test chamber is assured by the presence, before the
insertion of the model, of a 100mm-diameter hemi-spherical calibration probe made of
copper, cooled, that measures radial profiles of stagnation pressure (Ps) and stagnation heat
flux (Qs) at a section 0.375 m downstream of the conical nozzle exit section, by means of
high precision pressure transducers and Gardon-Gage heat flux sensors, respectively. Facility
regulations (mass flow, current) are tuned in order to measure on the calibration probe a
certain couple of values (Ps, Qs) which correspond to the desired set point in terms of the
couple (P0, H0).

3.2. Facility Performance Evaluation

The definition of a representative experiment in the CIRA Plasma Wind Tunnel
“Scirocco” has been done by considering the most interesting points of the EXPERT
reference trajectory, i.e. point P1, that is the point characterized by the highest stagnation
point heat flux, and point P2, characterized by high heat flux and a relatively low pressure,
potentially critical for passive/active oxidation transition of the C-SiC, which is the material
of the nose and the flaps of the capsule.
A preliminary analysis of PWT Scirocco capabilities for the duplication of SWBLI flows
has been carried out, the aim being to understand what it is possible to reproduce in this
plasma facility in terms of the characteristic parameters of the interaction as pressure and heat
flux (peak values and reference values, upstream of the separation), viscous interaction
parameter, (χ L ) ( )
≈ M ∞ / Re ∞L , rarefaction parameter, VL ≈ M ∞ / Re ∞L , separation
3

length experienced during the flight that has been preliminary predicted through CFD
simulations. Even if in this phase preliminary flight values have been used, it was important
only to develop an extrapolation from flight methodology that allowed to set the facility
operating parameters necessary to duplicate assigned flight values, that have been then
updated through three-dimensional non equilibrium computations.
Starting from the nominal operating envelope of PWT facility, considering the conical
nozzle D (length 3.1 m from the throat section, exit diameter 1.15 m) and assuming fully
laminar flows, a certain number of numerical simulations has been performed, basing on the
currently explored region of the envelope, where qualification and validation tests have been
already executed in order to have a clear idea of what can be simulated in terms of SWBLI
interesting parameters inside the facility; in particular, the effects of total pressure (at low,
medium and high total enthalpy) and total enthalpy (at low, medium and high total pressure)
have been investigated. Then, additional computations have been performed in order to


duplicate the estimated flight values of the interesting parameters of the interaction around the
body-flap. The complete CFD test matrix, for this preliminary phase, is reported in Table 1 in
terms of total pressure (P0) and total enthalpy (H0), while in Figure 4 these points have been
shown inside the PWT Scirocco theoretical envelope.

Table 1. CFD Matrix for preliminary computations

P0 (bar) H0 (MJ/kg)
PWT-1 2.45 11.90
PWT-2 2.45 15.00
PWT-3 2.40 18.80
PWT-4 4.70 10.40
PWT-5 5.20 15.00
PWT-6 5.20 18.80
PWT-72 7.90 11.00
PWT-8 7.90 15.00
PWT-9 7.90 18.80
PWT-10 13.00 11.00
PWT-111 12.00 15.00
PWT-121 10.00 15.00
PWT-13 7.90 17.90
PWT-14 10.00 11.00

45

40

35 PWT Operating Envelope
Preliminary Computations
30 Additional Computations
H0, MJ/kg

25

20

15

10

5

0
0 2 4 6 8 10 12 14 16 18
P0, bar

Figure 4. PWT envelope and operating conditions.

For each of the considered points, the PWT nozzle flow has been simulated in the
hypothesis of fully laminar thermo-chemical non equilibrium flow, then the centreline
conditions at X = Xnozzle exit + 0.15m (= 3.25 m) have been taken as free stream conditions to
perform the simulation of flow around the model (located preliminarily 0.15 m downstream
of the nozzle exit section).

2 For these conditions an angle of attack of the model equal to 10 deg has been also considered


This CFD-based procedure was also successfully applied to previous experiments design,
where it was clearly shown that the same test chamber flow is predicted by means of the
uncoupled (single simulation of nozzle flow with geometrical extension to the model
stagnation point section) as done in this work and coupled (complete simulation of flow
through PWT facility, with test chamber details) simulations, thus assessing the accuracy of
the overall experimental test design.
The preliminary geometry, which can be considered representative of the EXPERT
geometry around the body-flap region (scale 1:2), was a 0.60 m long blunted flat plate with a
0.15 m long flap forming a 20 deg angle with the plane and a cylindrical nose whose radius
was equal to 0.25 m as schematically reported in Figure 5.

Figure 5. Preliminary shape of the test article.

All the computations have been performed in non-equilibrium fully laminar conditions,
assuming a fully catalytic wall with a fixed temperature equal to 300 K or the radiative
equilibrium condition, considering the symmetry plane of the model with a two dimensional
approach. For numerical reasons a horizontal plate has been considered at the flap trailing
edge; the effect of this plate, that could fix the reattachment at the end of the flap, has been
analyzed for the conditions PWT-7 and PWT-3 of Table 1, that are the conditions,
respectively, characterized by the highest and the lowest Reynolds number.
This effect seems to be negligible (see Figure 6 and Figure 7) being the reattachment
mechanism not “driven” by the expansion at the flap trailing edge, but it occurs on the flap
“far enough” from its trailing edge. From the computed results of the present analysis, and
from considerations about the complexity of baseflow (useless) prediction, it can be
concluded that the geometry model with the flat plate extension behind the flap can be always
employed for present simulations.

mach
7.5 0.4
7
1.4 6.5
6 0.3
5.5
Y (m)

1.2 5
4.5
0.2
4
3.5
1 3
0.1
2.5
2 0.5 0.6 0.7 0.8
X (m)
0.9 1

0.8 1.5
Y (m)

1
0.5
0
0.6 -0.5

0.4

0.2

0
0 0.2 0.4 0.6 0.8 1 1.2 1.4
X (m)

Figure 6. Mach contours for the condition PWT-7 with base flow.


1200 200000 1200 200000

Pressure 180000 180000
Pressure - Base
1000 Heat Flux 1000
Heat Flux - Base
160000 Pressure 160000
Pressure - Base
Heat Flux
140000 Heat Flux - Base
140000

Heat Flux (W/m2)

Heat Flux (W/m2)
800 800

Pressure (Pa)

Pressure (Pa)
120000 120000

600 100000 600 100000

80000 80000
400 400
60000 60000

40000 40000
200 200
20000 20000

0 0 0 0
0.4 0.6 0.8 0.4 0.6 0.8
X (m) X (m)

Figure 7. Pressure and Heat Flux distribution for PWT-7 (left) and PWT-3 (right).

The pressure distribution along the wall for all the computations is reported in Figure 8
(left); after the rapid expansion starting from the stagnation point, a quasi-constant pressure
region is observed along the flat plate up to the zone of shock wave boundary layer
interaction induced by the presence of the flap; at the separation location there is a pressure
jump due to the separation shock, then pressure reaches a plateau in the recirculation region
and a peak after the reattachment on the flap followed by a sharp expansion at the end of the
flap. The overall distribution is typical of such SWBLI interaction. Pressure levels on the
model are mainly influenced by the value of the total pressure being negligible the effect of
the total enthalpy; therefore, it is clear that higher values of the pressure in the interaction
zone can be achieved with higher values of the total pressure (or additionally giving an
incidence to the model in the test chamber).
In the same figure (right) it is reported the wall heat flux distribution for the same
computations; in this case both the total pressure and total enthalpy influence wall heat
transfer, that increases as these two variables increase (following roughly the dependency
upon the product p 0 H 0 ).

8000 1E+06

H0=11.90 P0=2.45 800000 H0=11.90 P0=2.45
6000 H0=15.00 P0=2.45 H0=15.00 P0=2.45
H0=18.80 P0=2.40 H0=18.80 P0=2.40
Heat Flux (W/m2)

H0=10.40 P0=4.70 H0=10.40 P0=4.70
Pressure (Pa)

H0=15.00 P0=5.20 H0=15.00 P0=5.20
H0=18.80 P0=5.20 600000 H0=18.80 P0=5.20
H0=11.00 P0=7.90 H0=11.00 P0=7.90
4000 H0=15.00 P0=7.90 H0=15.00 P0=7.90
H0=18.80 P0=7.90 H0=18.80 P0=7.90

400000

2000
200000

0 0
0 0.2 0.4 0.6 0.8 0 0.2 0.4 0.6 0.8
X (m) X (m)

Figure 8. Wall pressure (left) and heat flux (right) distributions in PWT operating conditions.

In the recirculation region there is a decrease of the heat flux, typical of fully laminar
interactions, followed by an increase on the flap and a peak just immediately after the
reattachment point, where boundary layer thickness reaches the minimum value.


For some of the computations, the effect of the wall temperature has been estimated by
considering the radiative equilibrium condition. The results of these computations are
reported in Figure 9 for pressure (left) and heat flux (right) wall distribution. As general trend,
with the radiative equilibrium temperature at the wall, the separation bubble is larger (except
for the condition PWT-3 where the effect is negligible) and the peak loads over the flap (both
thermal and mechanical) are lower than those predicted with a fixed wall temperature equal to
300K.
This is due to the higher temperatures in the boundary layer in the case of the radiative
equilibrium condition, and then to the lower values of density causing an increase of the
boundary layer thickness; the upstream propagation of pressure disturbances is enhanced in
the case of radiative equilibrium and, consequently, an early separation is predicted.
The effects on mechanical loads is a reduction of ∼4% with the condition of equilibrium
radiative wall whereas is ∼3% for thermal loads as reported also in Table 2. It can be
concluded that in these conditions surface temperature has only a small effect on thermal and
mechanical loads acting on the flap.

1600 400000

H0=11.90 P0=2.45 T=300K H0=11.90 P0=2.45 T=300K
1400 H0=11.90 P0=2.45 T=Tradeq H0=11.90 P0=2.45 T=Tradeq
H0=18.80 P0=2.40 T=300K H0=18.80 P0=2.40 T=300K
H0=18.80 P0=2.40 T=Tradeq H0=18.80 P0=2.40 T=Tradeq
1200 H0=15.00 P0=5.20 T=300K 300000 H0=15.00 P0=5.20 T=300K
H0=11.00 P0=7.90 T=300K
Heat Flux (W/m2)

H0=11.00 P0=7.90 T=300K
Pressure (Pa)

1000 H0=18.80 P0=7.90 T=300K H0=18.80 P0=7.90 T=300K
800 200000

600

400 100000

200

0 0
0.2 0.4 0.6 0.8 0.2 0.4 0.6 0.8
X (m) X (m)

Figure 9. Twall effects on wall pressure (left) and heat flux (right) distributions.

Table 2. Tw effects: comparison of the peak values on the flap

Tw=300K Tw=Trad.eq.
Ppk qpk Ppk qpk
(Pa) (kW/m2) (Pa) (kW/m2)
PWT – 1 331.7 78.8 306.5 72.9
PWT – 3 328.8 131.6 328.7 129.7
PWT – 5 680.4 171.2 639.8 167.0
PWT – 7 1093.7 146.9 1054.3 142.8
PWT – 9 949.0 293.2 907.2 286.5


3.3. Definition of PWT Model

The wide amount of CFD results obtained in different PWT conditions has permitted the
development of the extrapolation-from-flight procedure: it allows to determine the
experimental test conditions (P0, H0 and model attitude) able to duplicate the representative
mechanical and thermal loads ahead and over the flap. However, in order to give the final
requirements for the detailed model design and then for the execution of the tests, it has been
necessary to consider also different aspects of the problem, not only the aerothermodynamic
ones.
A detailed numerical analysis has been carried out to analyse the effects of geometric
variation of the model on the flow variables, in particular, the effects of the nose radius, the
flap dimension and the model’s finite span have been considered. Sensitivity analysis has
been carried out considering the PWT operating condition characterized by a reservoir
enthalpy H0=15MJ/kg and a reservoir pressure P0=10 bar, being this condition the one
determined for the duplication of the point P1 flight conditions over the model as it will be
described hereinafter.

3.3.1. Nose Radius
Computations with the radiative equilibrium wall assumption have shown that
temperature in the nose region could reach 2000 K. If there will be the possibility to have an
active cooling system (at least in the nose region) the size of the nose could be decreased in
order to not exceed the model weight limit for the “Scirocco” Model Support System (MSS).
A sensitivity analysis to the nose radius has been then carried out for one of the selected
operating conditions inside the PWT operating envelope, by considering three different
models with the same length of the plate ahead the flap and three different nose radii, equal to
0.25 m (the first hypothesis), 0.1 m and 0.05 m; it has been found that the influence of nose
radius is small in terms of mechanical loads (see Figure 10, left) even if a slight decreases of
about 5% is predicted in the reference and peak values whereas, for what concerns the
thermal loads (see Figure 10, right ), a slight increase of the values in front of the flap and a
small decrease of the peak values is predicted as also reported in Table 3.

8000 H0=15 MJ/kg P0=10 bar AoA=12 deg H0=15 MJ/kg P0=10 bar AoA=12 deg

1.5E+06

6000
Rnose = 25 cm
Heat Flux (W/m2)

Rnose = 25 cm Rnose = 10 cm
Pressure (Pa)

Rnose = 10 cm Rnose = 5 cm
Rnose = 5 cm 1E+06

4000

500000

2000

0 0.2 0.4 0.6 0.8 0 0.2 0.4 0.6 0.8
X (m) X (m)

Figure 10. Effects of nose radius on wall pressure and heat flux.


Table 3. Nose radius effects: mechanical and thermal loads ahead and over the flap

Rnose Pref Qref Ppk qpk
(m) (Pa) (kW/m2) (Pa) (kW/m2)
0.25 1024.91 120.85 2226.34 266.34
0.10 1002.02 152.36 2095.86 252.26
0.05 890.92 167.18 2004.07 239.16

The reduction of the nose radius causes a decrease of the separated region mainly due to
the movement towards the flap hinge line of the separation point whereas the reattachment
point is located more or less in the same position for all the analyzed configurations (see
Figure 11).

H0=15 MJ/kg P0=10 bar AoA=12 deg

0.03
Skin Friction Coefficient

Rnose = 25 cm
0.02 Rnose = 10 cm
Rnose = 5 cm

0.01

0

0 0.2 0.4 0.6 0.8
X (m)

Figure 11. Nose radius effects: skin friction distribution.

From this analysis it results that the model with the nose radius equal to 0.1m seems to be
the best solution for the model configuration, also considering the fact that a lower value of
the radius could make difficult the handling and positioning of model instrumentation
whereas the model with the biggest value of the nose radius could result in a too heavy model
difficult to sustain during the test execution with the MSS.

3.3.2. Flap Dimensions
Another variation that has been considered with respect to the preliminarily selected
model has been done by considering the full scale flap dimensions, thus exploring the
possibility to test in PWT “Scirocco” the actual EXPERT open flap before the flight, whose
overall dimensions are 0.30 m in length and 0.40 m in width.
The effect of this variation has been examined with respect to the model with the nose
radius of 0.10 m, considering the same total length of the previous one since the extension of
the flat plate has been decreased from 0.35 m to 0.20 m. The results are shown in Figure 12;
considering the full scale EXPERT flap the size of the separation bubble decreases and the
thermal and mechanical loads over the flap increase. The effects in terms of wall pressure are
evident (Ppk increases of ∼24%) whereas are modest in terms of heat flux (qpk increases of
∼5%).


H0=15 MJ/kg P0=10 bar AoA=12 deg
8000 1200

Flap 1:2
1000

6000 Flap 1:1

Heat Flux [kW/m ]
2
800

Pressure [Pa]
pressure Flap 1:1
pressure Flap 1:2 600
tot. wall flux Flap 1:1
4000 tot. wall flux Flap 1:2
Geometry
400

200
2000

0
0 0.2 0.4 0.6 0.8 1
X [m]

Figure 12. Effects of flap dimensions.

3.4. Final Configuration and Materials

The final configuration of the model, whose characteristic dimensions are Rnose = 0.10 m,
Lplate = 0.20 m (the flap hinge is located at X=0.30m starting from the nose), Lflap = 0.30 m
(projection on the X-axis), corresponding to the full scale 1:1 flap and flap deflection angle =
20 deg is shown in Figure 13.

MATERIALS
0.4 Nose : TBD
Plate : PM1000
Flap : C-SiC
0.3
Lplate = 0.20 m Lflap = 0.30 m (scale 1:1)

0.2
Y (m)

δflap = 20 deg
0.1

R nose = 0.10 m
0

-0.1
0 0.1 0.2 0.3 0.4 0.5 0.6
X (m)

Figure 13. Final model configuration.

The final geometrical configuration of the model to be tested in the plasma wind tunnel
“Scirocco” is a trade-off between the aerothermodynamic requirements necessary to
reproduce the flight characteristic parameters of the interaction in PWT conditions, and the
thermo-mechanical design issues that have taken under consideration also different aspects of
the problem.
The model reproduces the EXPERT capsule flap (scale 1:1) characterized by 20 deg
deflection angle; it is mounted on an holder with a flat plate ahead the flap with rounded
leading and lateral edges. In Figure 14 it is reported a schematic representation of the model.
To be consistent with the EXPERT capsule, the model will be built by using as much as
possible the same materials to manufacture its different parts: the leading edge is a GLIDCOP
AL-15 copper cylinder with an active cooling system; the upper part is covered by a flat plate
of PM1000 equipped with pressure taps, thermocouples and combined heat flux/pressure


sensors; the flap is covered by a 4mm thick plate of C-SiC with a deflection angle of 20 deg
with respect to the flat plate, and it is equipped with pressure taps and thermocouples. The
lateral rounded panels, the entire lower panel and the parts below the PM1000 flat plate will
be realized in PROMASIL 1100; the wedge underlying the C-SiC flap will be realized in
amorphous carbon.
For what concerns the dimensions of the model, the cylinder leading edge has a radius of
100mm and a length of 400mm, the flat plate is 400m wide and 200mm long, the flap is
400mm wide and 300mm long. All the lateral edges are rounded with a radius of 50mm in
order to avoid localized over heating, whereas the flap plate has a radius of curvature at the
lateral edges equal to 4mm (i.e. its thickness).
The model will be installed on the PWT Model Support System (MMS) by means of a
proper interface that consists of a commercial steel circular beam built with AISI 316L; the
interface is covered by proper thermal insulator of PROMASIL 1100 to avoid any critical
solicitation due to the plasma interaction with the model surface. Such a covering has a
cylindrical shape for the proper alignment of the upper and lower parts of the test article with
the MSS body surface. It is realised to avoid the presence of gaps between the surfaces and
the possibility of any peak heating occurrence.

Figure 14. Model for PWT tests.

4. EXTRAPOLATION FROM FLIGHT PROCEDURE
The definition of representative experiments in PWT has been done by considering the
most interesting points of the EXPERT reference trajectory: point P1 (M∞=13.40, h=37Km),
characterized by the highest stagnation point heat flux, and point P2 (M∞=18, h=50Km)
characterized by high heat flux and a relatively low pressure, potentially critical for
passive/active oxidation transition of the C-SiC.

4.1. Facility Operating Conditions

In par. 0, requirements for the execution of the PWT test campaign will be shown:
according to the extrapolation from flight procedure, those requirements must be duplicated


inside the facility on the representative model that has been defined and dimensioned. To this
purpose, it is necessary to define the facility operating conditions and model positioning and
attitude within the test chamber to achieve the goal.
The wide amount of CFD results obtained in different operating conditions has permitted
the development of the extrapolation-from-flight procedure, in such a way to determine the
experimental test conditions (P0, H0 and model angle of attack) that allow for the duplication
of the representative mechanical and thermal loads ahead and over the flap of the model.
For each of the computations carried out in PWT conditions, considering the effect of
facility operating conditions, of the angle of attack of the model, wall temperature, radius of
the nose and length of the flap, different variables of interest have been analyzed:

• Boundary layer edge variables at the separation location: M, Re, V and ρ
• Reference values at the separation location: χ, V, Pref and qref

The comparison between the characteristic SWBLI parameters estimated in flight
conditions and the results obtained in PWT for the selected test conditions, is shown, in terms
of reference pressure (Pref) and heat flux (qref), in Figure 15 and Figure 16, respectively. For
the range of trajectory around the maximum stagnation point heat flux (Point P1, H0∼13.2
MJ/kg), it seems not possible to duplicate the reference values of pressure and heat flux which
instead could be well enough duplicated for points at higher enthalpy, that it clearly means at
higher altitudes (Point P2, H0∼18 MJ/Kg, h∼50 Km).

2.50E+05

CFD Flight 2D
Qref Flight 2D
Qref Flight 3D
Point P1
2.00E+05 Point P2
Qref - PWT
Qref - PWT Additional Runs

1.50E+05
Qref (W/m2)

1.00E+05

5.00E+04

0.00E+00
0.00E+00 2.00E+06 4.00E+06 6.00E+06 8.00E+06 1.00E+07 1.20E+07 1.40E+07 1.60E+07 1.80E+07 2.00E+07
H0 (J/Kg)

Figure 15. Comparison between qw ahead the flap in flight and PWT conditions.

1.00E+05

CFD Flight 2D
Pref - FLIGHT 2D
Pref FLIGHT 3D
Point P1
1.00E+04 Point P2
Pref - PW T
Pref - PW T Additional Runs
Pref (Pa)

1.00E+03

1.00E+02

1.00E+01
8.00E+06 1.00E+07 1.20E+07 1.40E+07 1.60E+07 1.80E+07 2.00E+07
H0 (J/Kg)

Figure 16. Comparison between Pw ahead the flap in flight and PWT conditions.

Wind tunnels

Empfohlen

Empfohlen

Weitere ähnliche Inhalte

Was ist angesagt?

Was ist angesagt? (20)

Andere mochten auch

Andere mochten auch (19)

Ähnlich wie Wind tunnels

Ähnlich wie Wind tunnels (20)

Kürzlich hochgeladen

Kürzlich hochgeladen (20)

Wind tunnels