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A short history of the European Transonic Wind Tunnel ETW
John Green a,n, Jurgen Quest b
aAircraft Research Association, Manton Lane, Bedford MK41 7PF, United Kingdomb ETW GmbH, Ernst-Mach-Strasse, 51147 Koln, Germany
a r t i c l e i n f o
Keywords:
Wind tunnel history
Cryogenic wind tunnels
Wind tunnel design
Wind tunnel test techniques and model
instrumentation
Transonic aerodynamics
European aeronautical collaboration
a b s t r a c t
This paper is written as a contribution to the celebration of 50 years of Progress in Aerospace Sciences
and of the centenary of the birth of its founder, Dietrich Kuchemann. It reviews the evolution of the
European Transonic Wind Tunnel, ETW, from early conceptual studies to its entry into service and its
current capabilities and achievements. It traces the development, from the earliest days, of experi-
mental aerodynamics and of the basic aerodynamic understanding that gave rise to the main periods of
wind tunnel building before and after World War II. By about 1960, this activity appeared to have come
to a natural halt. The paper gives an account of the role of Kuchemann in arguing the need in 1968 for a
further step in wind tunnel capability, to provide transonic testing at high Reynolds numbers. It
describes his leading role in gaining acceptance of the concept, formulating the specification and
promoting studies of alternative, radical design options for the co-operative European project that
became ETW. The progress of ETW through design, construction, commissioning and into full operation
is recorded. The paper discusses the many technical innovations that have been introduced in order to
meet customer requirements in the challenging field of aerodynamic testing in a cryogenic environ-
ment and, finally, looks to the future and the further technical challenges that it holds.
& 2011 Elsevier Ltd. All rights reserved.
Contents
1. Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 320
2. Experimental aerodynamics in the beginning 17421917. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 321
2.1. Early insights, 17421904 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 321
2.2. The evolution of the wind tunnel up to 1917. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 323
3. The coming of age of the wind tunnel 19171945 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 325
3.1. The pursuit of full scale Reynolds numbers in the 1920s and 1930s. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 325
3.2. The significance of compressibility and the first high-speed tunnels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 326
4. The great period of wind-tunnel building 19451959 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 327
5. Emergence of the need for higher Reynolds number 19591968 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 329
6. Definition of the requirement and the solution for Europe 19681978 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 331
6.1. The role of AGARD and Kuchemann . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 331
6.2. The work of LaWs and MiniLaWs 19711974 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 333
6.3. AEROTEST and AC/243 (PG.7) 19721973. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3346.4. The LaWs specification and the four original design concepts. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 335
6.4.1. The transonic Ludwieg Tube tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 336
6.4.2. The Evans Clean Tunnel (ECT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 337
6.4.3. The Injector-Driven Tunnel (IDT). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 338
6.4.4. The Hydraulic-Driven Tunnel (HDT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 338
6.5. Engineering studies of the four design concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 339
6.6. The coming of cryogenics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 341
6.7. The underlying physics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 343
6.8. Evolution of the specification, from LEHRT to ETW 19751978 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 345
Contents lists available at ScienceDirect
journal homepage: ww w.elsevier.com/locate/paerosci
Progress in Aerospace Sciences
0376-0421/$- see front matter& 2011 Elsevier Ltd. All rights reserved.
doi:10.1016/j.paerosci.2011.06.002
n Corresponding author. Tel./fax: 44 1525 290631.
E-mail address: [email protected] (J. Green).
Progress in Aerospace Sciences 47 (2011) 319368
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7. Designing the ETW 19781988 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 346
7.1. Phase 2.1 preliminary design 19781985 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 346
7.2. Phase 2.2 final design and the Rogers task force 19851988. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 351
8. Establishing the GmbH 1988. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 352
9. The construction phase 19881993 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 352
10. Hardware characteristics of the facility . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 353
10.1. The settling chamber and its downstream contraction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 353
10.2. The drive system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 354
10.3. The nitrogen system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 354
10.4. Model handling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35410.5. The test section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 355
11. Getting Wind on 19932000 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 356
11.1. Tuning and calibrating. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 356
11.2. Client testing in the 1990s. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 357
11.3. Developing techniques for gathering fully corrected data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 357
12. Operating in the 21th century 20002010 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 359
12.1. Contributing to European research . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 359
13. Developing and enhancing test techniques. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 360
13.1. Further development and enhancement of test techniques . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 363
13.2. Laminar wings are back. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 365
14. Summary and outlook . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366
References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 367
Further reading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 368
1. Introduction
The website of ETW GmbH boldly asserts, The European Transo-
nic Wind Tunnel, ETW, in Cologne, Germany, is the most modern
wind tunnel in the world; a unique test facility for the development
of new transport aircraft. Fig. 1is an aerial view of the facility. This
paper gives an account of its evolution, achievements to date, current
capabilities and the part that Dietrich Kuchemann played in its
creation.
There are two high Reynolds number transonic wind tunnels
in the world, the National Transonic Facility (NTF) at NASA
Langley and ETW in Cologne. The tunnels are similar in size and
operating principle both are cryogenic, using gaseous nitrogen
at near liquefaction temperatures as the working fluid and bothfar surpass all other wind tunnels in their ability to test aircraft
models at Reynolds numbers equal to, or near to, those of flight.
The NTF first ran in 1983. Its reported cost was $85 million. ETW
first ran 10 years later, in 1993. Its construction cost was 562
million Deutschmark at 1987 prices, roughly twice the cost of the
NTF when adjusted for inflation, and in some key respects it is a
more advanced facility than the NTF. But we recall that Isaac
Newton said, If I have seen further, it is only by standing on the
shoulders of giants. ETW and Isaac Newton may seem an unusual
juxtaposition, but there is a parallel here, in that the ETW has
beyond doubt benefited immensely from the pioneering work of
NASA that led to the NTF. Without the NTF, there would be no
ETW as we know it today.
It is also possible that, without Dietrich Kuchemann, there
would be no ETW at all. It was he who led the intellectual debate
within NATO that resulted eventually in four nations, France,
Germany, The Netherlands and the United Kingdom, deciding in
1973 to co-operate in a project to build a high Reynolds number
transonic tunnel for Europe. The drive towards such a tunnel was
triggered by events in the 1960s that were reported at a Specialists
Meeting[1]of the AGARD Fluid Dynamics Panel (FDP) in September
1968. The subject of the meeting was Transonic Aerodynamics andKuchemann, who was a member of the Programme Committee, was
asked to prepare a Technical Evaluation Report on the meeting.
It was a happy chance for Europe that the task fell to Kuchemann.
He was, at that time, the Head of Aerodynamics Department at
the Royal Aircraft Establishment in Farnborough, internationally
respected and with deep insight into the application of the results
of aerodynamic research to aircraft design, particularly the design of
aircraft operating at transonic conditions. He was also a believer in
getting things done rather than merely philosophising and his
energy and commitment to making progress played a vital part in
shaping the concept of a co-operative European transonic tunnel
and in convincing the four nations of the need for it.
Much of the work that was done under his leadership was on
alternative, novel concepts for a tunnel with air at ambienttemperature as the working fluid. By 1974, however, the concept
of a cryogenic transonic tunnel, in which higher Reynolds num-
bers are achieved by testing in gaseous nitrogen at very low
temperatures, had been shown at NASA Langley to be an attrac-
tive possibility. In October 1975, at the Specialists Meeting of the
AGARD FDP on Wind Tunnel Design and Testing Techniques [2],
the first paper was from NASA Langley, presenting the results
obtained in a pilot cryogenic transonic tunnel and setting out
the plans for the NTF. Kuchemann, in his closing address as the
outgoing Chairman of the Panel, noted the potential of the
cryogenic tunnel, welcomed the progress that had been made in
the USA and, speaking of Europe, ended his address, So all I need
to do now is to quote one of the speakers who said: Now, let us do
it! Just over 4 months later, on 23 February 1976, he diedFig. 1. Aerial view of the European Transonic Wind Tunnel.
J. Green, J. Quest / Progress in Aerospace Sciences 47 (2011) 319 368320
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unexpectedly, after a short illness. He was not to see his vision
come to fruition but, through his inspirational leadership, he had
set the course that the four nations followed to its logical
conclusion. There were obstacles and delays on the road ahead
but the end result meets the requirements that were set out
under his guidance in the early 1970s and stands as a testament
to his insight, perseverance and ability to inspire his colleagues.
The authors of this paper span more than 40 years of involvement
in ETW. In 1970 John Green, a member of Aerodynamics Departmentat the Royal Aircraft Establishment, was invited (instructed) by his
Department Head (Kuchemann) to write a paper on viscous flows
over wings for the AGARD FDP meeting[3]in May 1971, for which
Kuchemann was Chairman of the Programme Committee. In the
following years he contributed other papers to the studies led by
Kuchemann. In the period 19781981, the early years of Phase 2.1,
the Preliminary Design phase of the ETW project, he was the UK
member of the Steering Committee,1 chairing the committee in
1980. In the early 1990s, having left Government service, he was a
member of ETW Advisory Committee 1.2 He has written the first part
of the paper, covering the years from 1742, when projectile drag at
transonic speeds was first measured, to 1988 when the four nations
resolved to carry the ETW project through to its conclusion. Jurgen
Quest joined ETW in 1988, after the Technical Group had moved
from Amsterdam to Cologne, at the time of the establishment of ETW
GmbH and the start of Phase 3, the construction and operation phase.
He is currently ETW Chief Aerodynamicist and he has written the
second part of the paper, from 1988 to the present day.
2. Experimental aerodynamics in the beginning 17421917
2.1. Early insights, 17421904
In the late 19th century and for the first third of the 20th century,
the University of Gottingen was an academic centre of world
renown.3 It was host to an outstanding collection of mathematicians
and scientists whose research in many fields, not least aerody-
namics, paved the way for many of the advances of the 20thcentury. Dietrich Kuchemann was born and educated in Gottingen,
took his doctorate at Gottingen University under Ludwig Prandtl in
1936 and, for the next 10 years, continued his aerodynamic research
at the A.V.A. (Aerodynamische Versusch Anstalt) Gottingen. There is
no doubt that he was imbued with the spirit that prevailed there in
his student years. In May 1975 he spoke eloquently of the Gottingen
spirit der Gottinger Geist in his opening address to a symposium
of the AGARD Fluid Dynamics Panel on Flow Separation held in the
town[4]. He characterised the spirit as a determination to know, to
understand, coupled with the firm intention that knowledge gained
should be applied usefully, should be of benefit to human society.
That spirit was clearly in evidence throughout his determined efforts
to drive forward the European studies that led finally to the ETW.
We begin this paper with a brief review of the evolution of ourunderstanding of the aerodynamic phenomena that gave rise to the
need for Europe to build the ETW. Central to that understanding are
three fundamental conceptual advances, all linked in some way to
Gottingen but, to begin, we go back to the 18th century, and to
experiments rather than theory, for the starting point in our narrative.
In 1742 Benjamin Robins devised a ballistic pendulum, Fig. 2,which he employed to make what were, for that time, some
remarkably accurate measurements of the drag of a ball fired
from a musket [5]. His purpose was to demonstrate that the
resistance of the air had an important influence on the trajectory
of a cannon ball and that, as a consequence, all ballistic calcula-
tions at that time, which took the resistance to be negligible, were
ill founded. In this he succeeded, but he also discovered that, over
the range of velocities covered by his experiments, the drag of the
ball did not vary as the square of the velocity as predicted by
the accepted authority at that time, Sir Isaac Newton [6]. Over the
speed range that his experiments covered, (Mach 0.71.5 in
modern terminology) he found that the drag increased more
rapidly than the square of the velocity (Fig. 3). He had measured
transonic drag rise[7].Ernst Mach, born almost 100 years after Robins published his
paper, also studied ballistics experimentally. His most notable
contribution was the use of Schlieren photography to observe
gunshots and to display the pattern of shock waves created by a
bullet travelling at high speed. He observed that the inclination
of the shock wave was a function of the ratio of the speed of
the bullet to the speed of sound. This observation was initially of
little interest to aeronautical scientists until, as flight speeds
increased, one of the leaders in the field, Jakob Ackeret, who
had worked under Prandtl in Gottingen from 1921 to 1927,
published in 1927 an article on gasdynamics [8] in which he
proposed the term Mach number, MV=a, for this ratio of
velocities. Flight Mach number is the first parameter that ETW
is required to replicate.
Fig. 2. Ballistic pendulum of Benjamin Robins, 1742.
Source: Ackroyd, UKs contribution to development of aeronautics, Part 1, Aero J.,
January 2000.
1 With the foundation of ETW in 1988, the Steering Committee was expanded
slightly and became the Supervisory Board, the governing body of ETW.2 AC 1 was established by the Supervisory Board to provide advice on matters
related to the expected development of aerospace science and engineering,
especially in Europe.3 The coming to power of the Nazi party in Germany in 1933 was followed
almost immediately by the great purge of Jewish scientists, which resulted in
many of the most distinguished academics at Gottingen leaving the country.
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Shortly after Ernst Machs experiments in Prague with gunshots,
Osborne Reynolds, in Manchester in 1883, made some rather different
but no less significant experiments with water flowing through a
tube. The tube was of glass and a thin stream of coloured water
flowing down the middle of the tube was used to visualise the
behaviour of the flow. The man, his apparatus and the flow patterns
that he observed are shown inFig. 4. Patterns ac, seen by spark
illumination at three different, increasing flow rates, were termed by
Reynolds direct for a and sinuous for c today we call this laminar
and turbulent flow. By performing experiments with tubes of three
different diameters, and by varying the temperature of the water
and the flow rate through the tube, Reynolds established that the
character of the flow depended on a non-dimensional quantity
RnrVl=mwherer is the fluid density,Vits velocity,m its absoluteviscosity andl a length scale (in his experiments Reynolds chose the
tube diameter). At low values of this quantity the flow was laminar,at high values turbulent. From a consideration of the equations of
motion, Reynolds reasoned that the quantity represented the ratio of
inertial to viscous forces acting on a small volume of fluid and that
at some characteristic value of this quantity the flow would begin to
form eddies.
The term Reynolds number for the quantity Rn4 was first
proposed in 1908 [9] by the physicist Arnold Sommerfeld, a
graduate of Gottingen. But, even before it had been given a name,
its fundamental importance to aerodynamics had been recognised
and both Lord Rayleigh[10], in his 1884 Presidential Address to
the British Association in Montreal, and Lanchester [11]in 1907,
in his seminal book Aerodynamics, had identified equality of
this quantity as a requirement for fluid flows to be dynamically
similar.
In the early years of flight, although the significance of Reynolds
number was recognised, it was understood that it could not be
replicated in the ground test facilities of the time, whirling arms and
small wind tunnels. The true full-scale aerodynamics could be
realised only in flight by the full-scale machine. Fortunately, the
aerodynamic properties of early aircraft were not strongly depen-
dent on Reynolds number and failure to replicate flight values in the
ground test facilities of the time did not seriously undermine the
usefulness of these facilities.
In the period immediately after World War II, when many new,
large wind tunnels were built, both Mach number and Reynolds
number were recognised as important parameters of the new
generation of high speed aircraft. Although it was now possible to
replicate flight Mach numbers in the wind tunnel, the maximum
achievable Reynolds numbers were lower than flight by an order ofmagnitude. Hence the post-war practice evolved of testing and
reporting results at specific Mach numbers and of developing
methods of adjusting the data for the difference in Reynolds number
between tunnel and flight. This approach appeared to be satisfactory
for the first two decades that followed World War II.
Eventually, however, the approach was undermined by advances
in wing design that increased the importance of the behaviour of the
wing boundary layer. It was in 1904 that Prandtl, then a professor of
mechanics at the technical school in Hannover, presented his theory
of the boundary layer at the Third International Mathematical
Congress at Heidelberg[12]. Its impact was great and was a factor
no doubt in his appointment as director of the Institute for Technical
Physics at the University of Gottingen later that year. In fact,
Prandtls theory enabled the gulf between the theoretical results of19th century hydrodynamicists and the practical results of the
aeronautical experimenters finally to be bridged. It is arguably the
single most important concept in the evolution of aerodynamics,
explaining the key role of the boundary layer in determining
aerodynamic behaviour and also enabling the full power of inviscid
flow theory to be brought to bear on predicting the flow about
aircraft. In the years after 1904, Prandtl and his colleagues in
Gottingen made many further, fundamentally important contribu-
tions to our understanding of the flow about aerofoils and wings and
the behaviour of the boundary layer.
Fig. 3. Comparison between sphere drag measured by Robins using ballistic pendu-
lum and present day result.
Source:Ref.[7],Fig. 9.
Fig. 4. Osborne Reynolds with experimental apparatus and observed flow
patterns, 1883.
Source:internet.
4 The usual symbol for Reynolds number is R. Here we useRn so that we may
useR for the universal gas constant.
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Theory and experimental capabilities advanced together in the
early decades of the 20th century with the support of wind tunnel
testing becoming increasingly important in the development of any
new aircraft. As aircraft design evolved and flight speeds increased,
so wind tunnel facilities grew in size, complexity and cost. By 1904,
the three key concepts, Mach number, Reynolds number and the
boundary layer had emerged. However, more than 60 years were to
pass before the need for a wind tunnel to simulate not only Mach
number but also to approximate flight Reynolds number closely, inorder to simulate the behaviour of the boundary layer in flight as
accurately as possible, became generally recognised.
2.2. The evolution of the wind tunnel up to 1917
In 1742, Benjamin Robins determined the drag of musket balls
by measuring the velocities of balls fired over different ranges
with a fixed charge of powder and calculating the deceleration
from the reduction in impact velocity with increase in range. Four
years later, in 1746, he reported experiments with a whirling arm
apparatus (Fig. 5) in which a weight rotated a drum that carried
the test object on a long arm. Drag was determined by the weight
while velocity of the test object was measured by timing a
number of revolutions of the arm. This gave him more accuratedrag data, for a range of shapes, but only in low speed flow.
The whirling arm concept was taken up by others, notably
Sir George Cayley, who in 1804 used a whirling arm to measure
the lift force on a square plate at angles of incidence between
31 and 181 (Fig. 6). Using these data he designed, built and
successfully flew a model glider (Fig. 7) believed to have been
the first successful heavier than air vehicle in history. In the 19th
century several other researchers used the whirling arm, notably
Otto Lilienthal, who between 1866 and 1889 built several whir-
ling arms of different sizes and measured the lift and drag
characteristics of a variety of aerofoils[13]. He also made similar
measurements of the forces on stationary aerofoils in the wind
over open ground. Because the whirling arm created a swirling
motion in the air around it, there were doubts about the validity
of the data it produced and Lilienthal concluded that his mea-
surements in the natural wind were the more reliable. He used
these in the design of the gliders in which he made more than
2500 flights between 1891 and his final, fatal flight in 1896. In
1895 he published tables derived from his natural-wind measure-
ments and these, republished in the USA in 1897, were used by
the Wright brothers to design their gliders of 1900 and 1901.
Meanwhile, in Britain, Francis Wenham, following unsatisfactory
experiments with a whirling arm, in 1871 persuaded the Aeronau-
tical Society of Great Britain to raise the funds to build a wind
tunnel, the worlds first. It consisted of a duct 12 ft long and
18 in18 in in cross section with a fan upstream of the model
driven by a steam engine. It had poor flow quality but nevertheless,
from tests on a variety of wing shapes, two significant results
emerged. First, that at small angles of incidence the lift force varies
in proportion to the sine of the angle of incidence, rather than to the
square of the sine. Secondly, that wings of high aspect ratio had
higher lift to drag ratios than those of low aspect ratio.5 In the early
1880s, also in Britain, Horatio Phillips built a wind tunnel of similar
proportions but driven by a steam ejector. This produced a steadierflow and led to Phillips developing and patenting a series of
cambered aerofoils, considered the first truly modern aerofoils.
Others followed Wenham and Phillips in building and experiment-
ing in wind tunnels, but with little further impact until the decisive
step forward taken by the Wright brothers in the autumn of 1901.
The Wrights had designed their first glider using Lilienthals
tables of normal and axial force. When they took it to Kitty Hawk,
North Carolina in September 1900, they had some limited success
but found that its lift was rather lower than had been expected.
Results the following year, with a new glider with increased wing
area, also fell well below expectations. The Wrights concluded that
Lilienthals tables were not reliable6 and in the autumn of 1901 built
themselves a wind tunnel similar to Wenhams, with a 16 in16 in
test section and a two-bladed fan driven by a gasoline engine(Fig. 8). They measured the lift and drag of some 200 model wings
Fig. 5. Benjamin Robins whirling arm, 1746.
Source:NASA Centennial of Flight.www.centennialofflight.gov.This is a re-drawn
version of the Robins original. The latter is in Ackroyd on the same page asFig. 2.
Fig. 7. Sir George Cayleys glider, 1804.
Source: Scanned from download from rsnr.royalsocietypublishing.org, paper by
Ackroyd on Cayley (2002) p. 175.
Fig. 6. Sir George Cayleys whirling arm, 1804.
Source:rsnr.royalsocietypublishing.org, paper by Ackroyd on Cayley (2002) p. 173.
5 The sine squared law, Newtons theory [6], had led to the widely accepted
conclusion that heavier than air flight was not practicable. Others, notably Cayley,
had found a linear variation of lift with incidence for low aspect ratio surfaces but
it was Wenhams discovery of high lift to drag ratios for high aspect ratio surfaces
that gave members of the Aeronautical Society reason to believe that heavier than
air flight would one day be achieved.6 The error in Lilienthals tables, which were based on his measurements of
the forces on an aerofoil in a natural wind, arose from his use of a plate
anemometer calibrated on a value of plate drag coefficient quoted by Smeaton
in 1759 from whirling arm results obtained by his friend, a certain Mr. Rouse of
Harborough. The Wright Brothers determined from their wind tunnel tests that
the Smeaton coefficient was incorrect and should have been 0.0033 rather than
the 0.005 that had been widely used for the previous century and a half.
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with different aerofoil sections and planform, using a simple balance
of their own design which gave accurate and repeatable results.
Their 1902 glider (Fig. 9) designed on the basis of their wind tunnel
results, had nearly twice the span of the 1900 glider. At Kitty Hawk,
in 5 weeks in September and October during which they made
between 700 and 1000 glides, the brothers developed the flight
controls for this machine so that it was fully controllable in three
dimensions. It was also more efficient aerodynamically than any-
thing that had gone before, with a lift-to-drag ratio of 8. They had
established a solid basis for the larger machine with which, inDecember 1903, they made the first controlled powered flights. It is
an achievement that would not have been possible without their
wind tunnel test programme in 1901.
Early in the 20th century, Gustav Eiffel began aerodynamic
investigation by measuring the aerodynamic forces on objects
dropped from the second platform of the Eiffel Tower, 377 ft above
ground level. He followed this in 1909 by building a wind tunnel of
novel design in the shadow of the Eiffel Tower. Its test section was
1.5 m diameter and its fan was driven by an electric motor drawing
on the towers power supply. In 1912 he built a similar but larger
tunnel at Auteuil (Fig. 10) and patented the design. Like the earlier,
smaller tunnels of Wenham, Phillips and the Wrights, it was an open-
return tunnel, housed in a hangar, but having an open jet test section
with air drawn into the jet nozzle though a bellmouth by a fan at the
outlet from the diffuser downstream of the test section. Eiffels
introduction of the bellmouth and diffuser meant that the pressure
in the test section was lower than the pressure in the hangar and the
test section therefore had to be inside a hermetically sealed enclo-
sure, the experimental room. Eiffels experiments led to a number of
significant advances; he pioneered the testing of models of complete
aircraft and, in resolving a factor of two disagreement between his
results and those of Prandtl in Gottingen, in 1914 he demonstrated
for the first time the sharp drop in the drag of a sphere as Reynolds
number is increased above 300,000 approximately, when the
boundary layer on the sphere changes from laminar to turbulent
[13]. The Eiffel wind tunnel concept was considered a success and
further, larger versions were built in the following decades.
Some 600 km to the North-East in Gottingen, at Prandtls
suggestion, the German Society for Airship Study (Motorluftschiff
Studiengesellschaft) in 1907 funded the construction of a simple
wind tunnel at a cost of 20,000 marks. It had a closed returncircuit of rectangular planform with a closed test section 2 m2.
There was a honeycomb flow straightener downstream of the fan
in the return leg and cascades of turning vanes at each corner.
However, because almost the entire circuit had the same cross-
sectional area and flow velocity as the test section, flow quality in
the test section was not particularly good. The tunnel was
constructed in 1908 and in 1909 began practical work on the
aerodynamics of airships. It was envisaged as a temporary facility
and in 1911 Prandtl made the first case for building something
more substantial. Negotiations for the funds for this were essen-
tially complete in 1914 when World War I broke out, the plans
were put on hold and the first wind tunnel, now concentrating on
aircraft aerodynamics, continued as an important test facility for
most of the war. In 1915 the case to build a second tunnel wasaccepted by the war administration and 300,000 marks were
made available15 times the funding for the first tunnel. The
project was completed and began operations in Spring 1917.
The second Gottingen tunnel (Fig. 11) was a great advance on
what had gone before and embodied for the first time many features
that have become standard in most tunnels built since then. In fact,
in the years that followed, wind tunnels tended to be classed as
either the Eiffel or the Gottingen design. The key features of the
Gottingen design were explained by Prandtl in a lecture in 1920
[14]. He had combined the idea of a contraction ahead of the open-
jet test section and a diffuser downstream, a concept he acknowl-
edged as coming from Eiffel, with a closed return circuit of
substantially greater cross-sectional area, and hence lower flow
velocities, than the test section. The Eiffel contraction and diffuser
Fig. 9. Wilbur Wright in the 1902 glider.
Source: NASA Centennial of Flight: www.centennialofflight.gov.
Fig. 10. Eiffel wind tunnel at Auteuil, 1912.
Source:The wind tunnels of NASA, NASA SP-440 Chapter 2.
Fig. 8. Wright Brothers wind tunnel, 1901.
Source:www.wright-brothers.org.
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increased the efficiency of the circuit and reduced the fan power
requirement while the closed return circuit, ending in a settling
chamber with a flow smoothing honeycomb, followed by a 5:1
contraction, produced a more uniform and steady flow than in
previous tunnels. The closed return circuit removed the need for the
testing room to be hermetically sealed, as was the case in the Eiffel
tunnel, and hence made the test area more accessible. Models were
mounted on carts that ran on transverse tracks and there were bays
on either side of the test section to enable one model to be prepared
while another was in the test sectiona feature that Prandtl
considered important and that has since been replicated in a
number of major wind tunnels, including ETW. The fan was driven
by a Ward Leonard set, an AC motor driving a DC generator to
supply the DC motor driving the fan. This gave accurate speed
control over a range from 50 to 1100 rpm and set a pattern that
became generally adopted for subsonic wind tunnel drive systems.
The use of testing with the model erect and inverted to determine
with precision the inclination of the flow in the tunnel, relative to
the direction of gravity, was another innovation that remains best
practice in todays wind tunnels.
There were some innovations, such as building the tunnel out of
reinforced concrete with its circuit in a vertical plane, that have beencopied less frequently, some of the advances in measurement
techniques and automatic speed control have been superseded
and, increasingly, closed test sections have been preferred to open
jets. Even so, many of the key features of the Gottingen tunnel can
be found in almost all the worlds major wind tunnels built since
1920. And two other practical considerations noted by Prandtl in
1920 have featured in the building of many major wind tunnels
since. First, though the drive power of the Gottingen tunnel was only
about 0.25 MW, that power was significant relative to the capacity
of the Gottingen local power supply at the time and an automatic
regulator was needed to avoid making large load increases suddenly.
Power availability has since featured in decisions on the location of a
number of large wind tunnels. Secondly, the design of the tunnel
circuit strikes a balance between running and capital costs. Thetunnel circuit is shorter and less aerodynamically efficient than
optimum, thereby reducing the cost of the tunnel shell. This trade-
off between capital and running costs has to be made in the design
of every major wind tunnel and was an important consideration
during the assessment of alternative drive systems for ETW.
3. The coming of age of the wind tunnel 19171945
3.1. The pursuit of full scale Reynolds numbers in the
1920s and 1930s
At the end of World War I the Gottingen tunnel could be
considered the state of the art. If we adopt the convention used
in specifying the ETW Reynolds number, that a typical wing chord is
0.1 times the square root of the cross-sectional area of the test
section, the maximum Reynolds number of the Gottingen tunnel
based on this typical chord was approximately 0.7 million. For the
Sopwith Camel, a typical WWI fighter aircraft, the chord Reynolds
number was approximately 4.7 million. For a larger aircraft, such as
the Vickers Vimy bomber, it was approximately 9 million. There was
thus an order of magnitude difference between characteristic tunnel
and flight Reynolds numbers.
Despite the Wright brothers having made the first controlled
powered flight in 1903 and having taken Europe by storm with
Wilbur Wrights demonstration of the abilities of the Wright Flyer in
Paris in 1908, aeronautical progress in the USA lagged far behind
progress in Europe in the following decade. This was recognised in
the USA as early as 1912 but it was not until March 1915 that
Congress, at the recommendation of the regents of the Smithsonian
Institution, passed legislation to establish the National Advisory
Committee for Aeronautics (NACA). In 1917 NACA established a
laboratory site in Hampton, Virginia and named it Langley Field.
NACA Wind Tunnel No. 1, a low speed tunnel of the Eiffel type with
a test section 1.5 m in diameter, began operation at Langley Field in
June 1920. Its characteristic Reynolds number based on one tenththe square root of its test section area was 0.37 million; its life was
relatively short and unproductive.
A year later, in June 1921, the NACA Executive Committee
decided to build a much more substantial and important tunnel,
the Langley Variable Density Tunnel (VDT), in order to test at
Reynolds numbers much closer to full scale flight. With Reynolds
number defined as
RnrVlm
,
wherer is the density,Vthe velocity,l a characteristic length and mabsolute viscosity, Rn can be increased, in air at ambient tempera-
ture, by increasing pressure and thereby density, increasing velocity
or increasing the characteristic length. The concept of increasing itby increasing pressure was proposed by Max Munk, who had
obtained his doctorate in Gottingen under Prandtl and had moved
to the USA to a post in NACA Headquarters in Washington in 1920.
Shown inFig. 12, the tunnel had a circular cross section with a
closed test section of 5 ft (1.5 m) diameter followed by a diffuser
embedded within an annular return circuit, all contained within a
cylindrical pressure vessel. The maximum velocity in the test
section was 50 mph, as against 90 mph for the no. 1 wind tunnel,
but it could test at pressures up to 20 atm and thus had a
characteristic Reynolds number of 4.2 million, comparable to
the chord Reynolds number of fighter aircraft of the time. The
tunnel became operational in 1923 and was used to obtain high
Reynolds number data on a wide range of aircraft and airship
types, A particularly substantial contribution was its testing of
Fig. 11. Prandtls wind tunnel at Gottingen, 1916.
Source:Ref. [13]p. 300.
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aerofoil sections; the aerodynamic data on 78 sections, published
in 1933 in NACA Technical Report 460, was an important land-
mark which prepared the ground for the design of the many new
aircraft developed in the USA before and during World War II.
The UK was impressed by the results from the VDT and, a
decade after the USA, built a similar tunnel at the National
Physical Laboratory (NPL) at Teddington. Over that decade Jones
[15] had published his paper on The Streamline Aeroplane,
aircraft design had progressed from biplanes braced with struts
and wires to the streamlined monoplanes competing in the
Schneider Trophy, the world airspeed record had doubled and,
with it, flight Reynolds numbers. Accordingly the NPL Com-
pressed Air Tunnel (CAT), which had essentially the same layout
as the NACA VDT, had a test section diameter and maximum
speed both 20% greater than the VDT. These increases, together
with an increase in maximum pressure to 25 atm, gave the CAT a
characteristic Reynolds number of 7.5 million.
Although the VDT and CAT were valuable sources of high
Reynolds number data, their layout, with flow smoothing honey-
combs in the relatively high speed stream immediately upstream
of the test section, resulted in comparatively high free stream
turbulence levels. These, as Doetsch [16] showed in 1936 fromtunnel turbulence data and NACA deduced from comparisons
between wind tunnel and flight, adversely affected the wind
tunnel results. The turbulence was understood to promote pre-
mature transition, a particular problem in research to develop
laminar flow aerofoils, and thus low turbulence became a desir-
able feature for future tunnels.7 To explore this question, NACA
built a pilot low turbulence tunnel which came into operation in
1939, designated the NACA Ice Tunnel.
NACA used the Ice Tunnel as the basis for the design of the Low
Turbulence Pressure Tunnel (LTPT) which went into operation at
Langley in 1941. The tunnel had a contraction ratio of 17.6:1, with
a combination of gauze screens and a honeycomb in the settling
chamber to minimise test section turbulence. The test section was
7.5 ft high3 ft wide, intended specifically for two-dimensionalaerofoil testing. With a maximum pressure of 10 atm and a max-
imum speed which varied with tunnel pressure but was about
130 mph at maximum pressure; its characteristic Reynolds number
was about 5.8 million. However, for tests on a two dimensional
aerofoil that fully spanned the tunnel, a chord of 2.0 ft was normal
and tests were typically done at Reynolds numbers of 3.0, 6.0 and
9.0 million. The tunnel played a key role in the development of the
NACA 6-series of laminar flow, low drag aerofoils that were adopted
for later WWII aircraft such as the highly successful P-51 escort
fighter.
Sixteen years before the LTPT went into service, and only 2
years after the VDT began operations, NACA decided to take the
complementary route to full-scale Reynolds number testing of
increasing tunnel size. The Propeller Research Tunnel (PRT), which
went into operation in July 1927, had an open jet test section 20 ft in
diameter and a stream velocity of 110 mph. The tunnel was used
mainly for tests on full-scale propellers, mounted in the fuselages of
real aircraft and driven by real engines. The propellers were full size,
running at their operational rotational speed and hence at virtually
full-scale Reynolds number. Many important advances came from
the ability given by this tunnel to test real hardware under realistic
aerodynamic conditions, including the development of the NACA
cowl for air-cooled engines, and led to NACA making the case for a
tunnel in which complete full-scale aircraft could be tested. Design
work on the Full-Scale Tunnel (FST) began in 1929 and the tunnel
began operations in spring 1931. It had an open jet test section of
30 ft60 ft (9.1 m18.3 m) and was driven by two 4000 hp (total
6 MW) motors, giving it a speed range of 25118 mph and a
maximum characteristic Reynolds number of 4.7 million. It played
a key role in US aircraft development in the 1930s and 1940s andremained in service for until 1995.
In 1939 another high Reynolds number tunnel came into opera-
tion at Langley, again with a drive of 8000 hp. This was the 19 ft
pressure tunnel, the first attempt anywhere to combine large scale
with high pressure. With a maximum pressure of 2.5 atm and a
maximum speed of 300 mph, its characteristic Reynolds number
was 11.9 million which enabled models of fighter and twin-engine
bomber aircraft to be tested at or near full-scale Reynolds number.
The advances in the USA were followed in Europe, both the UK
(Farnborough, 1934) and Germany (Braunschweig, 1940) building
8 m diameter tunnels in which, as in the NACA Propeller Research
Tunnel of 1927, full-scale propellers could be tested installed on an
aircraft. In France (Chalais-Meudon, 1934) a large tunnel with an
elliptical test section 16 m8 m was built. It was of the Eiffel type,with the open air rather than a hangar as the return circuit. Its
characteristic Reynolds number was 3.4 million. These and other,
smaller facilities played a part in enabling the respective national
industries to develop aircraft that would perform satisfactorily at
flight Reynolds numbers. All could be classed, however, as low-
speed tunnels, limited to testing aircraft at flight speeds at which
the flow around the aircraft could be treated as incompressible.
3.2. The significance of compressibility and the first high-speed
tunnels
As flight Mach numbers increase, the significance of compres-
sibility the local variation in air density caused by the passage of
the aircraft increases. Our insight into the behaviour of
Fig. 12. NACA Variable Density Tunnel (VDT) 1923.
Source:Ref. [13]p. 302.
7 We now know that free stream turbulence can also affect the development
of the turbulent boundary layer, as was recognised in specifying the turbulence
requirement for ETW (paper 4 in[35]).
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compressible flows began in 1816 with Laplaces correction of
Newtons theory for the speed of sound which recognised the
importance of the ratio of the specific heats g. Later in the centuryEarnshaw and Riemann (in Gottingen) studied waves of finite
amplitude in a compressible fluid, Rankine and Hugoniot formu-
lated the equations for a plane shock wave, Ernst Mach photo-
graphed shock waves and de Laval patented the convergent-
divergent nozzle for generating supersonic flow.
In Gottingen in 1908 Theodor Meyer, a doctoral student ofPrandtl, submitted a thesis in which the key relationships for
supersonic flow were developed, including the formulation of the
expansion fan in supersonic flow around a sharp corner (the
PrandtlMeyer expansion) and the equations for an oblique shock
wave. In his lecture in 1920 on the Gottingen wind tunnel, Prandtl
referred to the propeller drive system under development on
which a propeller of 1 m diameter will probably be driven at a
speed as high as 5000 rpm for the purpose of studying the
influences of the compressibility of the air. At the time, however,
the typical maximum aircraft speed was around Mach 0.15 and
compressibility, whilst it might affect the flow around a propeller
blade, had no significant effect on the flow around an aircraft.
In the years following World War I, interest in the effect of
compressibility on flow around propellers spread. However, with
wind tunnel drive power increasing as the test section velocity to
the power three, the power required to drive a continuous flow
tunnel of any size up to the speeds of propeller tips was too great
to be contemplated. In the USA NACA, after funding various small-
scale aerofoil tests in high speed jets at other sites, began in 1927
to design its own, small, high-speed tunnel. This was an inter-
mittent tunnel with an 11 in diameter test section, supplied by
atmospheric air drawn through the test section by a downstream
ejector (the same principle also underlay one of the candidate
drive systems considered for ETW). The air to drive the ejector
was supplied from the pressure vessel that was the outer shell of
the VDT, its capacity being sufficient for a run of approximately
1 min. The Reynolds number on an aerofoil with a chord of 2 in
was approximately 0.8 million.
Results from the tunnel [17]revealed a sharp rise in aerofoildrag as Mach number increased towards unity. The success of the
11-in high speed tunnel prompted the design of a larger version,
the 24-in high speed tunnel, and inspired the UK to build a 12 in
diameter tunnel of the same design which also mimicked the
NACA tunnel by drawing the air for its ejector drive from the NPL
Compressed Air Tunnel at Teddington. Both tunnels came into
operation in 1934 and made valuable contributions to improving
propeller performance over the next decade. In Germany the first
significant high-speed tunnel, built in Gottingen in the late 1930s,
was also a short-duration intermittent tunnel supplied by atmo-
spheric air but with the refinement that the air was dried by being
drawn through a silica-gel filter. It was based on a concept first
set out by Prandtl in 1912 of drawing air through the tunnel into a
large vacuum vessel. It had an open-jet test section and wasequipped with interchangeable nozzles, 11 cm 11 cm for sub-
sonic testing (0.5oMo1.0), 11 cm13 cm for supersonic testing
(the tunnel had a range of Laval nozzles, 1.2oMo3.2). It was in
this tunnel in the autumn of 1939 that Ludwieg (who in 1955
invented the Ludwieg Tube drive system considered for ETW)
made the first measurements on a swept wing at high subsonic
and supersonic speeds. He thereby confirmed the validity of the
swept wing concept for supersonic aircraft that had been advanced
by Busemann of Gottingen at the Volta Conference in Rome in 1935.
The limitations of the intermittent high-speed tunnels at Langley,
small model size and limited testing time, led in 1933 to NACA
beginning the design of a large continuous running tunnel. This, the
Langley 8-ft high-speed tunnel, was completed in March 1936. It
was the first, and for 5 years the only, large high-speed wind tunnel
in the world. Built of reinforced concrete, driven by an 8000 hp
motor and with a maximum Mach number of 0.75 in a test section
8 ft in diameter, it had a maximum characteristic Reynolds number
of 3 million and played an important part in the development of US
combat aircraft in World War II. A particularly valuable contribution
was the solution of the problem of severe buffeting and loss of
control that affected the Lockheed P38 fighter in a steep dive, arising
from shock wave oscillation on the wing at high subsonic speeds.
The problem was cured by the development in the tunnel of a diveflap on the lower surface of the wing.
In the early 1940s other large high-speed tunnels came into
operation. New 16ft tunnels were built at NACA Langley and NACA
Ames, in the UK a 10 ft7 ft high-speed tunnel was built at
Farnborough and in Germany three high speed tunnels with test
section diameters in the range 2.7 to 3.0 m were built. By 1945 there
were also a number of medium-sized supersonic tunnels in
Germany, including a continuous running tunnel of 94 cm94cm
test section at Braunschweig which covered both the subsonic range
and supersonic Mach numbers between 1.1 and 1.8. The most
ambitious German project, launched in 1940, was the 8 m diameter
high-speed tunnel to be built in the Austrian Alps in the Otztal. This
was to be the largest high-speed tunnel in the world. It was not
completed when World War II ended and the components were
thereafter transferred to Modane, in the French Alps, where it became
the major facility at the newly created ONERA test centre. Its original
specification was for a maximum speed of 300 m/s and a character-
istic Reynolds number of 8.5 million. This called for a drive power of
76 MW, a very severe demand which was to be met by locating the
tunnel in a mountain valley and driving a pair of contra-rotating fans
by a pair of water turbines (Pelton wheels) supplied from a mountain
reservoir 530 m above. The rebuilding at the ONERA Modane-Avrieux
centre was completed and the tunnel went into service as S1MA in
1952. The drive was, as originally conceived, by Pelton wheels
supplied from a reservoir in the mountains above. This has been
perhaps the most extreme example of the location of a major tunnel
being determined by its power requirement.
4. The great period of wind-tunnel building 19451959
In the years immediately following the end of World War II
there was a surge forward in planning new wind tunnels, driven
partly by the realisation of the advances in both wind tunnel and
aircraft design made in Germany during the war. After the war
some of the German tunnels were dismantled and re-built in the
USA, France and Britain and many leading German aerodynami-
cists were recruited into the government laboratories, often to
continue research in the fields in which they had been working
previously. Dietrich Kuchemann and several of his colleagues
came to RAE Farnborough at that time.
Even before the end of the war, German and British jetpropelled aircraft were in operational service and the potential
for future supersonic aircraft was evident. In 1945 NACA, which
already had pilot supersonic tunnels at Langley and Ames, set in
hand the design of three large, continuous operation supersonic
wind tunnels, a 4 ft4 ft tunnel at Langley, a 6 ft 6 ft tunnel at
Ames and, at the Lewis Flight Propulsion Laboratory at Cleveland,
Ohio, a tunnel for jet engine testing with a test section measuring
8 ft 6 ft. These all had substantial power requirements, the
Langley tunnel being limited to a maximum operating pressure
of 0.25 atm because there was only 6000 hp available to drive it.
In contrast, the propulsion tunnel at Lewis had a drive power of
87,000 hp. In the UK, also before the end of the war, it had been
decided to build a major government wind-tunnel and flight-test
centre. This was to include a number of high-speed wind tunnels
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and the availability of substantial electricity supplies was an
important factor in the decision to locate the centre at Bedford.
The weakness in these plans was that in 1945 there was no
credible way of testing an aircraft model of realistic size at Mach
numbers greater than 0.9 in the wind tunnels of the time, because
of the flow choking in the vicinity of the model. The problem of
wall interference, the influence of the wind tunnel wall on the
flow field around the model, had been known for many years and
theoretical treatments of the boundary effects for both closed andopen jet test sections, and the corrections to apply to test results
for these effects, had been developed for low-speed flow. In the
1940s there was work on wall corrections for high-speed subsonic
flow but the methods were not applicable at Mach numbers close
to unity. However, it was well known that the interference effects
from solid walls and open jet boundaries were of opposite sign.
Starting from this, Wright at NACA Langley [18] developed a
theoretical model of a tunnel with longitudinal slots in which the
opposing interference effects of the solid and open sections of the
wall cancelled each other to produce, ideally, an interference-free
flow.8 Wrights work in 1946 led to the construction of a pilot
tunnel with a 12 inch slotted test section. This was a success,
showing much less wall interference and enabling Mach number
to be increased progressively through Mach 1 to low supersonic
speeds simply by increasing fan speed.
The result was a decision by NACA to install slotted walls in
both the 8 and 16 ft high-speed tunnels at Langley. The 8 ft
tunnel, which in February 1945 had had its drive power increased
from 8000 to 16,000 hp to give it an empty tunnel Mach number
of 1.0, was the first to be modified. It became operational as a
transonic tunnel in early 1950 and was followed shortly by the
re-powered 16 ft tunnel. The 8 ft tunnel was the facility in which the
transonic drag rise problem of the first generation of supersonic
fighters was identified and, through Whitcombs development of the
Area Rule,9 was solved. The Convair F102 supersonic interceptor was
the first aircraft to encounter this problem. Although powered by the
worlds most powerful jet engine of the time, tests in the 8 ft tunnel
indicated, and flight tests on the prototype confirmed, that it could
not go supersonic in level flight. Area ruling, narrowing the fuselage inplaces and adding bulges ahead and aft of the waist, overcame the
problem and saved the project.
Up to this point the large high-speed wind tunnels had relied on
air exchange between the tunnel and the outside air to remove the
motor power that is put in through the fan. The original 8 ft tunnel,
with 8000 hp being put into the fan, required to exchange about 1% of
its airflow with the outside atmosphere to maintain tunnel tempera-
ture at an acceptable level. With its doubled power, needed to
overcome the increased losses caused by the slotted test section
and to drive the flow to higher Mach numbers, the required exchange
rate with the outside air doubled. In the summer, when the humidity
of the outside air at Langley was invariably high, humidity within the
tunnel was similarly high and the temperature drop in the
acceleration to high speeds caused dense fog in the test section,
water droplets interfering with instrumentation and a deterioration
in tunnel flow quality. Langley quickly put in hand the construction of
a new tunnel with slotted walls, the 8-ft transonic pressure tunnel,
which removed the fan power via a water-cooled heat exchanger and,
being a sealed tunnel, avoided the problems of moist air. It could
operate at up to 2 atm pressure and had a high contraction ratio plus
screens in the settling chamber to provide a high flow quality in the
test section. It went into operation in 1953 and was the first facility toincorporate the essential features of a modern transonic tunnel.
Following the early experiments with slotted walls at Langley
there was work at Ames and the Cornell Aeronautical Laboratory to
explore the alternative of porous and perforated walls (walls with a
mesh of small holes) as a possible means of reducing the strength of
the reflection at the wall of the bow shock wave from the model. This
was significant only at Mach numbers around 1.0, where the reflected
shock could strike the rear of the model, but for military aircraft the
near-sonic Mach number range was a critical one. It was found that
perforated walls did indeed cause less interference than slotted walls
in the low supersonic regime effectively eliminating the reflected
shock and it was decided therefore to convert the 16 ft diameter
high-speed tunnel at Ames, which had been operating since 1941,
into a 14 ft14 ft transonic tunnel with perforated walls. The
increased test section drag, caused by the displacement of air into
the plenum chamber by the model and its subsequent return to the
tunnel stream with much reduced total pressure, together with an
increase in maximum test section Mach number, required the drive
power to be quadrupled, from 27,000 to 110,000 hp (82 MW). The
Ames 14 ft transonic tunnel began operation in 1955.
During this hectic period of tunnel development in the USA there
was also intense activity in Europe to create a new generation of
wind tunnels. In 1952, France had transferred the components of the
German 8 m high-speed tunnel from the Otztal to Modane and had
completed the construction of the tunnel as S1MA, thereby inaugu-
rating the ONERA Modane-Avrieux centre. In the UK, work was in
progress on a new flight and wind tunnel test centre near Bedford, a
very ambitious government project known at the time as the
National Aeronautical Establishment (NAE). By 1952 the first windtunnel was already in operation there, a 3 ft 3 ft supersonic tunnel
driven by plant originally used in the 94 cm94 cm supersonic
tunnel at Braunschweig mentioned above.10 A new 8 ft8 ft sub-
sonic/supersonic tunnel was also under construction to add to the
capability provided by the 10 ft7 ft tunnel at RAE Farnborough,
which was the only large high-speed tunnel in the UK at
the time. Also in the UK, in January 1952, the Aircraft Research
Association (ARA) was founded by 14 aircraft and engine compa-
nies with the specific aim of building a large high speed wind
tunnel for industrial project development work. In the Nether-
lands, after a hiatus of 28 months caused by a funding crisis, work
began again on a high-speed tunnel for the government aero-
nautical laboratory, the then NLL, in Amsterdam. Because the
power requirements for this tunnel were high relative to thecapacity of the local electricity supply, the 20 MW required to
drive its electric motor was supplied from a battery of five oil-
fired steam turbine power plants acquired as war surplus from US
navy destroyers.
At the time of this activity, the work at NACA to develop a
transonic test section was classified and work on the high speed
tunnels in Europe was proceeding without the benefit of the
US test section design knowledge. There is an account [21] by
8 The idea of a test section with a combination of solid wall and free air
boundaries has been attributed[19]to Prandtl (Gottingen) and Glauert (Farnbor-
ough) in the 1920s and to work by Wieselsberger of Gottingen in 1942 and Ferri in
Italy. Even so, there seems no doubt that Wright developed his theoretical model
independently and was the first to put it to the test of experiment. To add a
footnote to a footnote, it is worth recording that Glauert, an Englishman who was
a predecessor of Kuchemann as Head of Aerodynamics Department at RAE, was
killed on Saturday 4 August 1934 in an accident on the edge of the Farnborough
airfield when army engineers were using explosive charges to remove a tree
stump. The first author recalls Kuchemann recounting how, when the news
reached Gottingen, Prandtl called the laboratory staff together and said Gentle-
men, Glauert has been killed; we will do no work today and sent them home.9 Although many German aerodynamicists had emigrated to the US at the end
of WWII, it appears that Whitcomb discovered the area rule for himself, unaware
that it had been discovered by Frenzl in Germany in 1943[20]and was covered by
Junkers patent 932410 of 21 March 1944.
10 From 1971 to 1973 this tunnel was part of the first authors responsibilities.
Each year, the annual inspection of the compressors brought from Germany after
the war led to long deliberations as to whether the fatigue cracks in the casings
were getting worse. Finally, because of the cracks, the tunnel was taken out of
service in 1983.
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von Karman, then Chairman of AGARD, of how he helped to
ensure that the NLL tunnel was finally designed as a transonic
tunnel, with a slotted test section inside a plenum chamber,
rather than as a high-speed tunnel with solid walls. By the time
the design of the ARA Transonic Wind Tunnel was finalised,
Europe had been given access to the US advances in transonic
test section design and it was decided to adopt the Ames model of
perforated test section walls for the ARA tunnel. By 1954 the
original UK plans for a separate National Aeronautical Establishmenthad been drastically scaled down and the test centre had been
absorbed within the Royal Aircraft Establishment. The NAE became
RAE Bedford. The test section of its supersonic 3 ft3 ft tunnel was
adapted to accommodate slotted walls and in 1956 the former
10 ft7 ft high-speed tunnel at Farnborough returned to service
with slotted walls installed to convert it into the RAE 8 ft6 ft
transonic tunnel. In 1956 the ARA Transonic Wind Tunnel (TWT) ran
for the first time. It had a 9 ft8 ft test section with a flexible
supersonic nozzle ahead of the perforated test section and was
powered by a combination of a 25,000 hp motor driving the main
fan and a separate 14,000 hp motor driving an auxiliary compressor.
The latter was used at transonic and supersonic speeds to draw the
test section boundary layer air out through the perforated walls,
thereby reducing pressure losses in the diffuser and enabling the two-
stage fan to drive the tunnel up to Mach 1.4.11 In 1957 ONERA
brought the S2MA transonic-supersonic tunnel into operation in
Modane. This had separate test sections for transonic and supersonic
testing, the transonic section measuring 1.75 m1.77 m, and, as
with S1MA, it was powered by Pelton wheels. Also in 1957, the
8 ft8 ft tunnel at RAE Bedford was commissioned. This had a drive
power of 60 MW and a flexible supersonic nozzle with solid walls
through the test section which gave it a Mach number range from
low subsonic up to 2.812 but excluding the strictly transonic range. In
1959 the high speed tunnel at NLL (now NLR) came into operation
with a slotted 2 m1.6 m test section and a 20,000 hp drive. This
was the last new facility of its kind; the suite of the main post-war
transonic tunnels in Europe was now complete.
While Europe was building its large transonic tunnels in the
1950s, the USA was doing likewise. In 1949 the US Congress passedthe Unitary Wind Tunnel Plan Act and the Air Engineering Develop-
ment Center Act. The Unitary Plan embraced NACA, the USAF,
industry and universities and an early draught envisaged 33 large
transonic, supersonic and hypersonic wind tunnels costing almost $1
billion. As with the original UK plan for the National Aeronautical
Establishment at Bedford, budget realities resulted in a final plan of
more modest scale. Even so, the USAF Air Engineering Development
Center, now the Arnold Engineering Development Center (AEDC),
established in 1951 at Tullahoma Tennessee in order to be close to
the abundant hydroelectric power available from the Tennessee
Valley Authority, was a massive undertaking. It included two
16 ft16 ft wind tunnels to cover the Mach number range from
0.2 to 4.74. The transonic tunnel, with perforated walls, first ran in
1956 and since then has played a key role in all US military aircraftdevelopment. The other important transonic facility created under
this legislation was the NACA Ames transonic tunnel, with a slotted
test section 11 ft11 ft. This was part of the Ames Unitary Plan Wind
Tunnel Complex, which comprised the transonic tunnel and two
smaller supersonic tunnels, all linked to a set of motors and
compressors with an installed power of 180,000 hp. The Ames 11 ft
tunnel went into service in 1957. Other large high-speed tunnels that
were converted to transonic tunnels in the 1950s included the Boeing
8 ft12 ft tunnel in Seattle, the Cornell Aeronautical Laboratory (now
Calspan) 12 ft high speed tunnel at Buffalo, converted to a test section
8 ft8 ft in 1956, and the high speed tunnel at the Naval Surface
Warfare Center in Carderoc, Maryland, converted to a transonic test
section 10 ft7 ft in 1958.13
In addition to the large, fan-driven wind tunnels in Europe and
the USA there were also many smaller fan-driven or blow-downtransonic and supersonic tunnels built during that period, pri-
marily in industry but some also at universities. In his book on
transonic wind tunnels [22], published by AGARD in 1961,
Gothert lists 19 transonic tunnels in Europe and 30 in the USA.
5. Emergence of the need for higher Reynolds number
19591968
By 1960 the NATO nations had at their disposal an impressive
array of transonic and supersonic wind tunnels suitable for aircraft
development testing (there had also been a large programme of
wind tunnel building at TsAGI in Zhukovsky, near Moscow, but little
was known of that in the West at the time). Overall, the largest
tunnels in the USA had higher maximum Reynolds numbers thantheir European counterparts but the difference was not great. The
investment in these facilities had been substantial and there was a
feeling that, for transonic and supersonic testing, the job had been
done. On both sides of the Atlantic, the NATO nations now had the
wind tunnels they needed.
Gotherts book on transonic wind tunnels [22] sets out in
detail the level of understanding that had been reached in 15
years of intensive post-war development. In 1962 the Interna-
tional Union of Theoretical and Applied Mechanics (IUTAM) held a
Symposium Transsonicum in Aachen. In looking back on this in
1969, Kuchemann saw it as a meeting held at a time when many
of the researchers in transonic aerodynamics had already moved
to other fields, mainly space research. They had come back to the
meeting to present and sum up results which, in many cases, theyhad obtained long before. Since 1962, work in transonic aero-
dynamics had continued only at a relatively low level, the main
research being carried out by a few workers.
The priority given in the 1940s and the 1950s to transonic
aerodynamic research, and to the development of transonic wind
tunnels, had arisen mainly from the quest to develop supersonic
fighter aircraft. The transonic region had been important primarily as
one which the aircraft had to traverse controllably; once aerodynamic
knowledge and engine thrust had advanced to the point where that
hurdle could be cleared comfortably, which by 1960 they had, the
interest in transonics fell away. This was, however, only a temporary
fall in interest. In 1958 jet travel across the Atlantic began, the Comet
4 in September and the Boeing 707 in October. Both aircraft were
based on the late 1940s to the early 1950s aerodynamics but theirintroduction was followed by a rapid growth in air travel in the 1960s
and a demand from the airlines for larger and more efficient aircraft.
Also, in the spring of 1960, the US Air Force released Specific
Operational Requirement 182for a long-range freight aircraft, to which
Lockheed responded successfully with a large, turbofan-powered
swept-winged design, the Lockheed Model 300, subsequently desig-
nated C-141.
11 Many transonic tunnels use auxiliary suction to supplement the main fan
drive. Gothert[22]discusses the minimisation of total drive power by optimising
the balance between fan and auxiliary suction power.12 In 1970 the tunnel compressor was modified, reducing the top Mach
number to 2.5 in order to increase Reynolds number at high subsonic speeds to
approximately 8 million. The tunnel was taken out of service in 2002 and has since
been dismantled.
13 The structure of this tunnel and its drive fans a pair of cast steel contra-
rotating fans 19ft in diameter came from a 3 m high speed tunnel at Ottobrun
near Munich, an ambitious project that had not begun operation when the war
ended; the Carderoc tunnel went out of service in 1990 when one of the fans
suffered a catastrophic fatigue failure. The first author visited Carderoc shortly
after the failure and witnessed the devastation caused, even though the fan was
contained within a concrete shell.
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As Kuchemann [23]noted, by the time of the AGARD Specia-
lists Meeting in Paris in September 1968 on the subject of
Transonic Aerodynamics there had been a general revival of
interest in the subject, with the participants from Industry stating
that the importance of continued technical advances in this field
cannot be overemphasised. The interest now was in the next
generation of transport aircraft, both civil and military. These
were subsonic aircraft with moderately swept wings on the upper
surfaces of which, at the cruise condition, there was an embeddedregion of supersonic flow terminated by a shock wave. Three
months before the meeting the Lockheed C-5A Galaxy had begun
flight testing and 10 days after the meeting the first Boeing 747
was rolled out. This was almost 5 years after the first flight of the
Lockheed C-141 Starlifter on 17 December 1963, the 60th anni-
versary of the Wright Brothers first powered flight, and it was the
aerodynamics of the C-141 wing that raised the concerns that led
finally to ETW. There is an excellent account of the emergence of
ETW from this starting point in the book The European Transonic
Wind Tunnel ETW A European Resource for the World of
Aeronautics [24]by Jan van der Bliek, former Director of NLR and
the original member for the Netherlands on the ETW Steering
Committee. There is inevitably appreciable overlap between that
book and this paper and in some places we have unashamedly
borrowed van der Blieks words. In general, however, we discuss
the technical issues more fully and the policy and political issues
less fully than he does.
In his 1961 book [22], Gothert made no reference to Reynolds
number but by then the test centres and the industry had between
them developed their test methods and their procedures for extra-
polating from wind tunnel to flight. In 1958 Braslow and Knox [25]
had published a method for designing boundary-layer tripsnarrow
bands of distributed roughness to cause transition from a laminar to
a turbulent boundary layer. These bands were fixed to the wind
tunnel model, close to the leading edges of lifting surfaces, around
the aircraft nose, etc., so as to ensure a turbulent boundary layer
over effectively the entire surface of the model. This simulated flight,
to the extent that the full-scale aircraft would also have a turbulent
layer all over, and the correction to quantities such as drag to allowfor the difference between the Reynolds numbers in wind tunnel
and flight could be made relatively straightforwardly on the basis of
the then current understanding of the turbulent boundary layer.
Individual test centres had their own preferred method of tripping
the boundary layer and each company had its own methodology for
extrapolating from wind tunnel to flight. There was an acknowl-
edged, accepted level of uncertainty in this process but, overall,
testing in the major wind tunnels of the time was considered to be a
satisfactory basis for the aerodynamic design of a new aircraft.
In the 1960s advances in wing design changed the situation.
The problem became apparent at high subsonic Mach numbers
where there is a region of supersonic flow terminated by a shock
wave on the wing upper surface. As either Mach number or lift is
increased, a bubble of shock-induced boundary layer separationforms at the foot of the shock. With further increase in Mach
number or lift coefficient the shock strength and the extent of the
bubble increases until there is a sudden drop in trailing edge
pressure and loss of lift. For the aerofoil designs used in the early
1960s, the characteristic behaviour was for the bubble to grow
slowly with shock strength, with the chordwise position of the
shock remaining relatively unchanged until a critical condition
was reached in which the bubble expanded suddenly to cause the
drop in lift. This behaviour was not particularly scale sensitive, as
explained by Pearcey et al. [26], and observed differences between
wind tunnel and flight were not great.
Advances in wing design aimed at reducing wing weight led to
increases in wing thickness and aerodynamic loading. The newer
designs had a longer region of supersonic flow on the upper surface,
followed by a steeper pressure rise towards the trailing edge. It was
found that, at wind tunnel Reynolds numbers, the effect of increas-
ing Mach number or lift on the more recent aerofoil designs was to
generate both a bubble separation beneath the shock wave and a
separation at the trailing edge, the interaction between the two
determining the chordwise shock position. This feature resulted in
the shock position being sensitive to Reynolds number.
The first manifestation of this effect that was of practical impor-
tance was on the C-141. Designed on the basis of wind tunnel testing,its flight test results were sufficiently at variance with the wind tunnel
to put the viability of the project at risk for a while. Figs. 13 and 14,
from Loving [27], show wing pressure distributions on the wing upper
surface in wind tunnel and flight at Mach 0.75 and 0.85, respectively.
The wind tunnel tests followed the standard practice of the time, with
transition fixed near the leading edge. At a Mach number of 0.75,
where the flow of the upper surface was subcritical (i.e. Mo1.0
everywhere), the difference between wind tunnel and flight was
relatively small and consistent with previous experience. At a Mach
number of 0.85, however, the flow over the upper surface was
supercritical, reaching a peak Mach number of 1.32 in the flight case,
and the chordwise positions of the shock waves terminating the
supersonic region differed by about 20% of chord between tunnel and
flight. The result was a nose-down pitching moment in flight that was
appreciably higher than in the tunnel and, in consequence, the
tailplane downloads needed to trim the aircraft were appreciably
Fig. 13. Comparison between wing upper surface pressure distributions in wind
tunnel and flight results for C-141 aircraft at subcritical conditions [27].
Source:NASA TN D-3580,Fig. 1.
Fig. 14. Comparison between wing upper surface pressure distributions in wind
tunnel and flight results for C-141 aircraft at supercritical conditions [27].
Source:NASA TN D-3580,Fig. 2.
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higher than had been calculated on the basis of the tunnel tests. In the
event, a complete re-stressing of the aircraft was needed before it
could be decided that the project would meet its design requirements.
Lovings investigations[27]of the tunnel-to-flight discrepancy on
the C-141 were carried out in the NASA Langley 8 ft pressure tunnel.
They revealed that the tunnel results were strongly dependent on the
chordwise location of the transition strip. Loving found (Fig. 15) that
as the trip was moved rearwards, the upper surface pressure
distribution increasingly approached that in flight. With the tip
removed completely, to allow natural transition, the position of the
terminal shock was essentially the same in tunnel and flight. Black-
well[28]followed Lovings work with experiments in the Langley 8 ft
pressure tunnel on a large two-dimensional aerofoil that could be
tested at chord Reynolds numbers typical of tunnel tests on a
complete three-dimensional model (3.0 m