Wind Tunnels at NASA
Langley Research Center
The first major U.S. Government wind tunnel became operational
in 1921 and was located at the Langley Memorial Aeronautical
of the National Advisory Committee for Aeronautics
(NACA), which became NASA Langley Research Center in 1958. This
wind tunnel was crude when compared to the tunnels used at NASA
Langley today. Like the aircraft tested in them, wind tunnels
evolve as researchers discover ways to more accurately duplicate
Wind tunnels help researchers understand the forces acting on an
object as it moves through the atmosphere. They are also used to
measure and minimize the noise produced by aircraft and to optimize
engine efficiency. Although primarily used for aircraft, other
objects such as spacecraft, automobiles, ships, trucks, and
wheelchairs have been tested in Langley wind tunnels.
Why Wind Tunnels Are Used
At Langley, models in wind tunnels are used in conjunction with
computers and flight simulators to learn about the flight
characteristics of new aircraft designs and modifications.
Components such as structural materials, wings, ailerons,
horizontal stabilizers, fuselages, power systems, engine cowlings,
and landing gear all affect the flight characteristics of aircraft.
Small changes to one component can result in the modification of
another component of the aircraft. All effects of the changes may
not be clear until the aircraft experiences flight conditions.
Tests with models in wind tunnels allow the study of aircraft
designs without risk to a pilot or the expense of building a new
full-size test aircraft for every design improvement. Measurements,
or data, from wind tunnel tests are also used to refine computer
programs that predict the forces that act on a new aircraft or
How Wind Tunnels Simulate Flight
The forces that act on an aircraft are the same whether the
aircraft is moving through the air or the air is moving past a
stationary aircraft. Typically for wind tunnel tests, aircraft
models are stationary and the air flows past. Basically, a wind
tunnel is a tube through which air or some other gas flows so that
the effects of an object moving through an air stream can be
determined. A wind tunnel may be open and draw air from the room
into the test section, or the wind tunnel may be closed with the
air continuously circulating through the test section. To obtain
meaningful data, the researcher must insure that the airflow in the
wind tunnel is very similar to that found in flight. In the tunnel,
the researcher can control airflow conditions, such as speed,
temperature, humidity, density, and viscosity. In continuous flow
wind tunnels, the airflow is most often produced by a large fan.
For very high speed blowdown tunnels, the air is
collected in pressure vessels and released into the tunnel.
Because wind tunnels are in buildings (and the actual aircraft
flight environment is not!), care must be taken to avoid
introducing airflow abnormalities from the tunnel itself. To
accurately simulate flight, the airflow in a wind tunnel must be
smooth. Wind tunnels must also be free of the effects of turbulent
or unsmooth airflow, which can be caused by forcing the air around
the tunnel circuit. Devices such as turning vanes, screens, and
slots in tunnel walls help to maintain smooth airflow.
Simulating airflow at flight conditions in a wind tunnel is
complex. Wind tunnels usually specialize in simulating a particular
aspect of flight. Subsonic speed wind tunnels study flight that is
slower than the speed of sound. Transonic speed tunnels study
flight that is slightly below, at, and slightly above the speed of
sound. Supersonic speed tunnels examine flight faster than the
speed of sound and hypersonic speed tunnels look at flight more
than five times the speed of sound. There are special tunnels for
propulsion research, aircraft icing research, aircraft spin control
research, and even full-scale model tests. The speed of airflow in
a wind tunnel is usually expressed as a Mach number. For example,
moving at twice the speed of sound is Mach 2, moving at half the
speed of sound is Mach 0.5.
Early work in fluid mechanics, or the
study of how fluids and gases behave and their effect on objects in
a flow, indicated that the airflow around a scale model would not
correspond exactly to the flow around a full-scale aircraft. To
ensure the correlation of model data to full-scale aircraft data,
researchers also determine the Reynolds number of flow in a wind
The ratio between the speed of a craft and the speed of sound in
the surrounding medium (the atmosphere) is called a Mach number.
The speed of air flowing through a wind tunnel is usually expressed
in terms of the speed of sound, which is approximately 761 miles
per hour at sea level. However, the speed of sound through the
atmosphere varies with temperature. Sound travels more slowly
through cooler air. Aircraft usually fly at higher Mach numbers in
the upper atmosphere where the air is colder.
Reynolds number is a nondimensional parameter representing the
ratio of the momentum forces to the viscous forces in fluid (gas or
liquid) flow. Reynolds number expresses the relationship of the
density of the fluid, velocity, the dimension of an object, and the
coefficient of viscosity of the fluid relationship. Osborne
Reynolds (1842-1912) demonstrated in experiments that the fluid
flow over a scale model would be the same for the full-scale object
if certain flow parameters, or the Reynolds number, were the same
in both cases.
For example, the Reynolds number of 1/4-scale models tested at
flight velocities at atmospheric pressure would be too low by a
factor of 4. Because the Reynolds number is also proportional to
air density, a solution to the problem could be to test 1/4-scale
models at a pressure of 4 atmospheres. The Reynolds number would
then be the same in the wind tunnel tests and actual full-scale
In addition to Mach number and Reynolds number, researchers
measure airflow around a model or specific parts of a model,
pressure exerted on the model, lift, drag, and engine thrust. There
are many techniques that researchers use to obtain these
measurements from wind tunnels and models.
Some models have small ports with pressure transducers that
measure pressures on the model at specific locations. These force
measurements can be recorded at a high rate of speed by a computer
with data acquisition software. Strain gauges can also be used to
measure pressure at specific locations. Pressure sensitive paint
(PSP) can acquire pressure measurements on the entire model or
entire sections of a model. The model may be mounted on a balance
to directly measure the aircraft lift or drag. A wake rake, or a
row of transducers, can also measure drag. Model attitude or the
angle of the aircraft to the airflow is measured using either a
sensitive angular encoder or high-precision accelerometer.
Various flow visualization techniques, such as ultraviolet oil,
paint, or fluorescent minitufts (short strings) are used to
investigate airflow at specific locations on the surface of model.
Video cameras can be used to record aspects such as airflow
visualization, wing deformation, and model angle of attack
measurements at normal video recording speeds, or at high speed to
provide more detailed motion analysis. Schlieren and shadowgraph
systems, infrared photogrametry, and laser velocimeters are used to
visualize and document airflow.
Thermocouples measure temperatures either on the model or inside
the tunnel at specific locations. Temperature sensitive paint (TSP)
can be used to determine temperature changes on broader areas.
Measurements can be made with a focused microphone array to
determine noise levels produced by specific components of an
0.3-Meter Transonic Cryogenic Tunnel
The Langley 0.3-Meter Transonic Cryogenic Tunnel is used for
testing airfoil (wing) sections and other models at Reynolds
numbers up to 100 x 106 per foot and Mach numbers from
0.1 to 0.9. The adaptive walls, floor, and ceiling in the test
section can be moved to eliminate or reduce effects of air around
the model being artificially constrained by the tunnel walls, thus
better representing flight in the atmosphere. The Mach number,
pressure, temperature, angle of attack, and adaptive wall shapes
are automatically controlled by the tunnel computer system. The
normal test medium is gaseous nitrogen. Air can be used as the test
medium at ambient temperature.
Because of the large operational temperature envelope, one end
of the tunnel is free floating and is allowed to contract and
expand along the length, width, and height. The fan housing section
is the fixed point for the tunnel and encloses the 12-blade
aluminum fan. The tunnel has interlocks and fail-safe systems that
shut down and vent the appropriate systems when electric,
lubrication, hydraulic, cooling water, pneumatic systems failures,
or gas leaks are detected.
8-Foot High Temperature
The Langley 8-Foot High Temperature Tunnel (8-Ft HTT) is a
combustion-heated hypersonic blowdown-to-atmosphere wind tunnel
that provides simulation of flight enthalpy (relationship of an
object, its environment, and energy) for Mach numbers of 4, 5, and
7 and Reynolds numbers from 0.3 x 106 to 5.09 x
106 per foot, depending on the Mach number. The test
section will accommodate air-breathing hypersonic propulsion
systems and structural and thermal protection system components. A
radiant heater system can be used to simulate ascent or re-entry
14- by 22-Foot Subsonic Tunnel
The Langley 14- by 22-Foot Subsonic Tunnel (14- by 22-Ft ST) is
an atmospheric, closed-return tunnel that can reach a velocity of
348 ft/s. The Reynolds number ranges from 0 to 2.2 x 106
per foot. Test section airflow is produced by a 40-ft diameter,
9-blade fan. Tunnel configurations include a fully closed test
section, a closed test section with slotted walls, and an open test
section closed only on the floor.
The tunnel provides an improved understanding of the
aerodynamics of vertical and short takeoff and landing (V/STOL)
aircraft. The 14- by 22-Ft ST is also ideally suited for low-speed
tests to determine high-lift stability and control, aerodynamic
performance, rotorcraft acoustics, turboprop performance, and basic
wake and airflow surveys.
16-Foot Transonic Tunnel
The Langley 16-Foot Transonic Tunnel (16-Ft TT) is an
atmospheric, closed-circuit tunnel with a Mach number range of 0.2
to 1.25 and a Reynolds number range from 1 x 106 to 4 x
106 per foot. The test section of the tunnel is
octagonal with a distance of 15.5 ft across the flats. The twin
34-ft diameter drive fans have counterrotating blades. The 16-Ft TT
has capabilities for conducting propulsion airframe integration
(PAI) and has supported most major military programs both in the
developmental stage and in on going propulsion integration
research. The F-14, F-15, F-18, and B-1, as well as the more recent
Navy Advanced Technology Fighter (NATF), the AX, and the Joint
Advanced Strike Technology (JAST) Program have been tested in this
tunnel. The tunnel has also supported NASA programs by doing
extensive tests for the Space Shuttle and X-33, X-34, X-37, X-38,
Hyper-X, and experimental programs such as the Unmanned Air Combat
20-Foot Vertical Spin Tunnel
The Langley 20-Foot Vertical Spin Tunnel (20-Ft VST) is a
closed-throat, annular return wind tunnel operating at atmospheric
conditions. The test section velocity can be varied from 0 to
approximately 85 ft/s. Test section airflow is produced by a
3-blade, fixed-pitch fan. The motor allows rapid changes in fan
speed, which result in maximum flow accelerations in the test
section of -25 ft/s2 to 15 ft/s2.
Dynamically scaled, free-flying aircraft models can be tested
for spinning, tumbling, and other out-of-control situations.
Spacecraft models can be tested for free-fall and dynamic stability
characteristics. The spin-recovery characteristics of aircraft are
studied by using remote actuation of the aerodynamic control
surfaces of the models. Emergency spin-recovery parachutes systems
for flight test aircraft can also be determined.
Jet Exit Test Facility
The Langley Jet Exit Test Facility (JETF) is a ground test
stand. Engine nozzle mass-flow rates and nozzle axial thrust are
measured. The nozzle test range provides nozzle pressure ratios
that simulate static conditions for up to a Mach number of 3. The
JETF has been used to test conventional and advanced aircraft
propulsion system components.
The Langley Low-Turbulence Pressure Tunnel (LTPT) is a single
return, closed-circuit tunnel that can operate from 1 to 10
atmospheres. The LTPT has been used for tests of high-lift
airfoils, basic research, and theory validation. LTPT's
capabilities of low disturbance, variable density tests and
high-lift, multielement airfoil tests at Mach numbers from 0.05 to
0.5 and Reynolds numbers from 0.4 x 106 to 15 x
106 per foot are unique in the world. This tunnel is
ideal for preliminary aerodynamic configuration screening because
of low operational cost and relatively inexpensive models.
National Transonic Facility
The National Transonic Facility (NTF) at Langley is a high
pressure, cryogenic, closed-circuit wind tunnel with a Mach number
range from 0.1 to 1.2 and a Reynolds number range of 4 x
106 to 145 x 106 per foot. The test section
has 12 slots and 14 reentry flaps in the ceiling and floor. To
ensure minimal energy consumption, the interior of the pressure
shell is thermally insulated. The drive system consists of a fan
with variable inlet guide vanes for responsive Mach number control.
In the variable temperature cryogenic mode, nitrogen is the test
gas. In this mode, the NTF provides full-scale flight Reynolds
numbers without an increase in model size. In the ambient
temperature air mode, air is the test gas and a heat exchanger is
used to maintain the tunnel temperature. The NTF provides tests in
support of stability and control, cruise performance, stall buffet
onset, and configuration aerodynamics validation.
Transonic Dynamics Tunnel
The Langley Transonic Dynamics Tunnel (TDT) is specifically
dedicated to investigating flutter problems of fixed-wing aircraft.
However, the tunnel has been used to investigate other aeroelastic
phenomena such as fixed-wing buffet and divergence and to conduct
rotary-wing tests that investigate the performance, loads, and
stability characteristics of both helicopter and tilt-rotor
configurations. Researchers also use the tunnel to determine the
effects of ground-wind loads on launch vehicles. The effects of
gusts on aircraft can also be studied in the TDT. The tunnel
provides steady and unsteady aerodynamic pressure data to support
computational aeroelasticity and computational fluid dynamics
computer code development and validation.
The TDT is a closed-circuit, continuous flow, variable pressure
wind tunnel. The tunnel is capable of using either air or R-134a as
the test medium. Testing in R-134a has important advantages over
testing in air, particularly for aeroelastic models. These
advantages include improved full-scale aircraft simulation, higher
Reynolds numbers, easier fabrication of scaled models, reduced
tunnel power requirements, and in the case of rotary-wing models,
reduced model power requirements. The tunnel can operate up to a
Mach number of 1.2 and is capable of maximum Reynolds numbers of
about 3 x 106 per foot in air and 10 x 106
per foot in R-134a.
Unitary Plan Wind Tunnel
The Langley Unitary Plan Wind Tunnel is a closed-circuit,
continuous flow, variable density supersonic tunnel with two test
sections. Typical tests include force and moment, surface pressure
measurements, and visualization of on- and off-surface airflow
patterns. Tests involving jet effects, dynamic stability, model
deformation, global surface and off-body flow measurements, and
heat transfer are also performed.
One test section has a design Mach number range from 1.5 to 2.9
and the other has a Mach number range from 2.3 to 4.6. The tunnel
can provide continuous variation in Mach number during operation.
The maximum Reynolds number per foot varies from 6 x 106
to 11 x 106.