30
2 Wind Tunnels and Other Aerodynamic Test Facilities
2.2 From Wind Tunnel Test to Reality
Earlier ground tests were conducted while considering real operating conditions of
the aircraft so this led to wind tunnels being designed for testing full-scale model.
However, with the increase of the flight speeds, altitude and the size of the aircraft, it
became necessary to test wind tunnel models whose dimensions can be well inferior
to those of the real vehicle. On the other hand, in the automotive industry full-scale
land vehicles are widely tested, but in these cases the test speeds are lower compared
to those encountered in aviation.
Testing with a model requires satisfying dynamic similarity conditions in order
to be able to extrapolate the results obtained in the wind tunnel to the actual flight.
The most common similarity conditions to be respected in classical aerodynamics
are often quantified by dimensionless coefficients or parameters whose definition
results from the examination of the equations of motion. These are the following:
– For similar wall boundary conditions, the geometry of the real vehicle needs to be
reproduced or at least very representative. However, in practice, the smaller the
scale, the more difficult or impossible it is to represent all the geometrical details of
the real object (fasteners, panel connections, antennas, scoops, etc.). Wind tunnel
models are thus similar only from a point of view of the global shapes, which are
usually sufficient to determine the aerodynamic forces with the exception of the
drag which necessitates further details of the surface features.
– At high speeds (Mach number greater than approximately 0.5); the simulation of
compressibility effects makes it necessary to use a fluid with the same thermodynamic properties as during operating conditions. This constraint does not exist at
very low speeds where the effect of density can be neglected and thus eliminated
from the governing equations (except in the presence of heat transfer or reactive
flows). In this case representative tests can be carried out in water, in hydrodynamic
flumes or water tunnels (see Sect. 3.4), which facilitates flow visualisation.
– The reproduction of compressibility effects (in particular shock waves) imposes
a very high accuracy on the Mach numbers between operation and wind tunnel
test. In addition, the crossing of the sound barrier and the passage to supersonic
flight are marked by significant changes in aerodynamic forces and their point of
application. The compressibility effect is necessary as soon as the Mach number
of the flow exceeds a value 0.3.
– In order to reproduce the effects related to the fluid viscosity, it is necessary to
ensure that the Reynolds number already defined in Sect. 1.2 is matched:
Re =
ρV L
μ
which accounts for the density of the gas ρ, V, the velocity, L, a characteristic
dimension of the physical model and, μ, the dynamic of the gas. Most often, the
reference quantities are relative to the uniform upstream flow of velocity V ∞ , with
ρ = ρ ∞ and μ = μ ∞ . For an aircraft cruising at a Mach number of 0.8, the
2 Wind Tunnels and Other Aerodynamic Test Facilities
2.2 From Wind Tunnel Test to Reality
Earlier ground tests were conducted while considering real operating conditions of
the aircraft so this led to wind tunnels being designed for testing full-scale model.
However, with the increase of the flight speeds, altitude and the size of the aircraft, it
became necessary to test wind tunnel models whose dimensions can be well inferior
to those of the real vehicle. On the other hand, in the automotive industry full-scale
land vehicles are widely tested, but in these cases the test speeds are lower compared
to those encountered in aviation.
Testing with a model requires satisfying dynamic similarity conditions in order
to be able to extrapolate the results obtained in the wind tunnel to the actual flight.
The most common similarity conditions to be respected in classical aerodynamics
are often quantified by dimensionless coefficients or parameters whose definition
results from the examination of the equations of motion. These are the following:
– For similar wall boundary conditions, the geometry of the real vehicle needs to be
reproduced or at least very representative. However, in practice, the smaller the
scale, the more difficult or impossible it is to represent all the geometrical details of
the real object (fasteners, panel connections, antennas, scoops, etc.). Wind tunnel
models are thus similar only from a point of view of the global shapes, which are
usually sufficient to determine the aerodynamic forces with the exception of the
drag which necessitates further details of the surface features.
– At high speeds (Mach number greater than approximately 0.5); the simulation of
compressibility effects makes it necessary to use a fluid with the same thermodynamic properties as during operating conditions. This constraint does not exist at
very low speeds where the effect of density can be neglected and thus eliminated
from the governing equations (except in the presence of heat transfer or reactive
flows). In this case representative tests can be carried out in water, in hydrodynamic
flumes or water tunnels (see Sect. 3.4), which facilitates flow visualisation.
– The reproduction of compressibility effects (in particular shock waves) imposes
a very high accuracy on the Mach numbers between operation and wind tunnel
test. In addition, the crossing of the sound barrier and the passage to supersonic
flight are marked by significant changes in aerodynamic forces and their point of
application. The compressibility effect is necessary as soon as the Mach number
of the flow exceeds a value 0.3.
– In order to reproduce the effects related to the fluid viscosity, it is necessary to
ensure that the Reynolds number already defined in Sect. 1.2 is matched:
Re =
ρV L
μ
which accounts for the density of the gas ρ, V, the velocity, L, a characteristic
dimension of the physical model and, μ, the dynamic of the gas. Most often, the
reference quantities are relative to the uniform upstream flow of velocity V ∞ , with
ρ = ρ ∞ and μ = μ ∞ . For an aircraft cruising at a Mach number of 0.8, the