Descripción: notes on basics of wind tunnel for engineering students...
Wind Tunnel Techniques
3003
Wind tunnel techniques course depicts the types, working and characteristics of wind tunnels in the laboratory. The flow characteristics and flow visualization in the tunnel are recorded for further observations.
Objectives
The course should enable the students to: 1) Understand the non-dimensional number by Buckingham theorem 2) Differentiate the wind tunnels on the basis of circuit, air flow and working.. 3) Know the calibration of a wind tunnel. 4) Understand the pressure and force measurements in wind tunnel 5) Deduce the flow visualization techniques used in the wind tunnel testing
Outcome The students should be able to: 1) Solve the Buckingham theorem to find the SI unit of a parameter 2) Clearly understand the working of blow down, in draft tunnels and their specifications 3) Know about horizontal buoyancy, flow angularities are checked while calibration 4) Know about component axis balance and internal balances are read and understood for the measurements in wind tunnel 5) Get a clear idea about the smoke and tuft flow visualization procedures in WT testing
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Syllabus Unit I
Topic
No. of Lectures
Principles of Model Testing
6
Buckingham theorem, Non-dimensional numbers, Scale effects, Types of similarities
II
Wind tunnels
8
Classification, Special problems of testing in subsonic, transonic, supersonic and hypersonic speed regions, Layouts, Sizing and design considerations
III
Calibration of Wind Tunnels
11
Test section speed, Horizontal buoyancy, Flow angularities, Turbulence measurements, Associated instrumentation, Calibration of supersonic tunnels
IV
Wind Tunnel Measurements
12
Pressure and velocity measurements, Force measurements, Three and six component balances, Internal balances
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Syllabus Unit V
Topic
No. of Lectures
Flow visualization techniques
8
Surface and turft flow visualization techniques, Dye injection techniques, Optical methods of flow visualization
References 1 Barlow, Jewel B. Rae, William H. and Pope, Alan “Low speed wind tunnel testing”, III Ed., ISBN 13: 9780471557746, ISBN 10: 0471557749, 1999 (Wiley India Edition, 2010, About Rs. 1,000). 2 Pope, Alan and Goin, Kenneth L, “High-Speed Wind Tunnel Testing”, ISBN-13: 978-0882757278; ISBN-10: 088275727X, 1978.
Instructor: Test
Dr. N. Sitaram
E_mail:
[email protected]
First Periodical Second Model Seminar/ Attendance End-semester Test* Periodical Test* Exam Assignment/Quiz Examination
Weightage Duration
10% 2 periods
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10% 2 periods
20% 3 hours
10% -
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10% -
50% 3 hours
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Introduction CFD vs. EFD Computational Fluid Dynamics Requirements High end computer, software, printers, plotters etc.
Experimental Fluid Dynamics Experimental facility (Wind tunnel), Instrumentation etc.
Manpower
One or two for small CFD work, A small team for large CFD work, each member specializing in grid developing, solver, post processing etc.
Usually requires moderate number of people for small EFD for preparing and installation of model, for measurements etc. Requires large number of people with different skills for large experiments.
Initial Cost
Moderate
High
Time required
Moderate for computation, but large for post processing
Large for preparation, small for data acquisition and post processing
Information acquired
Hugh
Limited
CFD and EFD are complimentary. Both require careful working. Needs verification and validation, Extensive CFD can be carried out followed by limited comparison with EFD
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Principles of Model Testing Buckingham theorem “Dimension” is characteristic of the object, condition, or event and is described quantitatively in terms of defined “units”. A physical quantity is equal to the product of two elements: A quality or dimension A quantity expressed in terms of “units”
Dimensions Physical things are measurable in terms of three primitive qualities (Maxwell 1871) Mass (M) Length (L) Time (T) NOTE: Temperature, electrical charge, chemical quantity, and luminosity were added as “primitives” some years later.)
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Principles of Model Testing Buckingham theorem Examples Length (L) Velocity (L/T) Force (ML/T2) Units: Measurements systems: CGS, MKS, SI SI units are now the international standard (although many engineers continue to use Imperial or U.S.)
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Principles of Model Testing Buckingham theorem SI Primitive Units Dimension
Symbol
Unit
Symbol
Length
L
meter
Mass
M
kilogram kg
Time
T
second
s
Temperature
T
Kelvin
K
Elec. Current
I
Ampere
A
Luminous intensity
C
Candela cd
Amount of substance
N
Mole
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m
mol
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Principles of Model Testing Buckingham theorem SI Derived units Description Derived unit Symbol Dimension
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Force
Newton
N
kg m/s2
Energy
Joule
J
kg m2/s2
Pressure
Pascal
Pa
kg/(ms2)
Power
Watt
W
kg m2/s3
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Principles of Model Testing Buckingham theorem Fundamental rules: All terms in an equation must reduce to identical primitive dimensions Dimensions can be algebraically manipulated.
Uses: Check consistency of equations Deduce expression for physical phenomenon
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Principles of Model Testing Buckingham theorem Simple Example: Drag on a Sphere
Drag depends on FOUR parameters: sphere size (D); velocity (V); fluid density (); fluid viscosity ()
Difficult to know how to set up experiments to determine dependencies
Difficult to know how to present results (four graphs?)
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Principles of Model Testing Buckingham theorem Simple Example: Drag on a Sphere
F = ρV D Only one dependent and one independent variable Easy to set up experiments to determine dependency Easy to present results (one graph)
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Principles of Model Testing Buckingham theorem Simple Example: Drag on a Sphere
Experimentally derived relation between the dependent and independent variables
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Principles of Model Testing Buckingham theorem Step 1: List all the parameters involved Let n be the number of parameters Example: For drag on a sphere, F, V, D,
Hence
n=5
Step 2: Select a set of primary dimensions For example M (kg), L (m), T (s). Example: For drag on a sphere, choose MLT
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Principles of Model Testing Buckingham theorem Step 3 List the dimensions of all parameters Let r be the number of primary dimensions Example: For drag on a sphere r = 3
Parameter:
Unit:
F
ML/T
V 2
L/T
D 2
L
3
M/L
M/LT
Step 4 Select a set of m dimensional parameters that includes all the primary dimensions Example: For drag on a sphere (m = r = 3) select ϱ, V, D
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Principles of Model Testing Buckingham theorem Step 5 Set up dimensionless groups ps There will be (n – m)=2 equations Example: For drag on a sphere
ML M T L
M LT =
Π = Fϱ V D
L T
L
Exponents to be determined to satisfy dimensional homogeneity Equating exponents of M, L and T on left and right sides For M: 0=1+a
Hence a=-1
For T: 0=-2-b
Hence
For L: 0=1-3a+b+c
0=1+3-2+c
b=-2 Hence
c=-2
Hence the first dimensionless number can be written as
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=
F ρV D
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Principles of Model Testing Buckingham theorem Step 6 Check to see that each group obtained is dimensionless Example: For drag on a sphere
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Π =
F ρV D
Π =
=
ML L T M ρ ν
M L
T L
1 L
L LT L T M
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Principles of Model Testing Buckingham theorem Direct Rationalization of Dimensionless Groups Obtain the Π terms by simple reasoning: 1
Determine their number from Buckingham theorem. Make sure the appropriate number of groups is obtained.
2
Identify variables (g etc.) that contain force quantities and formulate the 2 2
corresponding force. For example, F=V L and F=VL. Then take ratios of these forces (F/F=VL/) to get dimensionless groups. 3
If the force groups do not comprise the number of dimensionless groups sought, then look for length type terms, velocity type terms, and/ or time type variables which can be divided to give dimensionless groups.
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Principles of Model Testing Buckingham theorem Direct Rationalization of Dimensionless Groups There may be more than one length in the problem, i. e. L and D which give the dimensionless group, L/D. Length type variables can also include area and volume, so That A/L2 and V/L3 are dimensionless groups. Velocity type terms include velocity, V, angular velocity,
w, and volume flow rate, Q,
which yield dimensionless groups, wr/V and Q/VL . 2
Time type terms consist of period, , of the motion and convective time scale, L/V, which give the dimensionless group, V/L
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Principles of Model Testing Types of forces encountered in fluid phenomenon Inertia Force, Fi: = mass X acceleration in the flowing fluid. Viscous Force, Fv: = shear stress due to viscosity X surface area of flow. Gravity Force, Fg: = mass X acceleration due to gravity. Pressure Force, Fp: = pressure intensity X C.S. area of flowing fluid.
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Principles of Model Testing Significant Dimensionless Groups in Fluid Mechanics These are numbers which are obtained by dividing the inertia force by viscous force or gravity force or pressure force or surface tension force or elastic force. As this is ratio of once force to other, it will be a dimensionless number. These are also called nondimensional parameters. The following are most important dimensionless numbers. Reynolds Number Froude Number Euler Number Mach Number These numbers are named after Fluid Mechanics Scientists who had established their significance.
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Principles of Model Testing Significant Dimensionless Groups in Fluid Mechanics Reynolds Number, Re: It is the ratio of inertia force to the viscous force of flowing fluid.
Velocity Volume Mass. . Velocity Fi Time Time Re Fv Shear Stress. Area Shear Stress. Area Q.V AV .V AV .V VL VL du V .A .A .A dy L
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Principles of Model Testing Significant Dimensionless Groups in Fluid Mechanics Euler Number, Eu: It is the ratio of inertia force to the pressure force of flowing fluid.
Velocity Fi Time Fp Pr essure. Area Mass.
Eu
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Q.V P. A
AV .V P. A
V2 P/
Volume . Velocity Time Pr essure. Area
V P/
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Principles of Model Testing Significant Dimensionless Groups in Fluid Mechanics Froude Number, Fe: It is the ratio of inertia force to the gravity force of flowing fluid.
Velocity Fi Time Fg Mass. Gavitational Acceleraion Mass.
Fe
Volume . Velocity Time Mass. Gavitational Acceleraion
AV .V V2 V Volume.g AL.g gL gL Q.V
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Principles of Model Testing Significant Dimensionless Groups in Fluid Mechanics Mach Number, M: It is the ratio of inertia force to the elastic force of flowing fluid.
M
Fi Fe
Q.V K .A
Where : C
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Velocity Time Elastic Stress. Area Mass.
AV .V K .A
L2V 2 2
KL
Volume . Velocity Time Elastic Stress. Area
V V K / C
K /
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Principles of Model Testing Dimensionless numbers for force quantities Force intensities, pressure and shear stresses, are non-dimensionalized with the dynamic pressure
Pressure coefficient:
=
Local shear stress coefficient:
=
Resultant forces, such as lift and drag, are non-dimensionalized with the dynamic pressure force,
ρ V A, where A is some characteristic area. Lift coefficient, C =
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and
Drag coefficient, C =
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Principles of Model Testing Dimensionless numbers for force quantities Torque, T, is force F times a moment arm, L, which is non dimensionalized by multiplying the dynamic pressure force by L to give torque coefficient.
Torque coefficient:
C =
Power is the rate of doing work, FL/T or FV, which requires the dynamic pressure force be multiplied by V to give power coefficient.
Power coefficient:
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C =
=
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Principles of Model Testing Elaborate Example: Power required to rotate a propeller The power, P required to rotate a propeller depends on the fluid density, , the propeller diameter, D, the rotational speed, w, the velocity, V, of the fluid approaching the propeller, the speed of the sound, C, and the angle, a, of a propeller blade.
P = f (, D, w, V, C, a) Seven variables (n=7) give four dimensionless groups (m=n-r=4). Two force like terms yield one dimensionless group. The dynamic pressure, can be used to non-dimensionalized forces. As power is F V,
Π =
P 1 ρV A 2
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Principles of Model Testing Elaborate Example: Power required to rotate a propeller For a rotating propeller, L=D and V=wD to obtain
Π =C =
P 1 ρω D 2
There are three velocity variables:
w, V and C, which give two dimensionless numbers.
V V Π = Advace ratio = Π = Mach number = ωD C AE 2751
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Principles of Model Testing Elaborate Example: Power required to rotate a propeller Need one more dimensionless group, but have not yet considered the angle, a. An angle can be defined in terms of its tangent which is the ratio of two lengths and is dimensionless. That is a is a Π
term.
Π =α Then
C = Π =
or
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P 1 2 ρω D
=
V V , ,α ωD C
Π ,Π ,Π ,Π
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Principles of Model Testing Dimensional Analysis Definition: Dimensional analysis is a process of formulating fluid mechanics problems in terms of non-dimensional variables and parameters. Why is it used :
• Reduction in variables [ If F (A1, A2, … , An) = 0, then f (P1, P2, … Pr < n) = 0, where, F = functional form, Ai = dimensional variables, Pj = non-dimensional parameters, m = number of important dimensions, n = number of dimensional variables,
r=n–m). Thereby the number of experiments required to determine f vs. F is reduced.
• Helps in understanding physics • Useful in data analysis and modeling • Enables scaling of different physical dimensions and fluid properties
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Principles of Model Testing Dimensional Analysis: Example
Drag = f (V, L, , C, t, e, T, etc.) From dimensional analysis,
Vortex shedding behind cylinder
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Principles of Model Testing Similarity and Model Testing Definition : Flow conditions for a model test are completely similar if all relevant dimensionless parameters have the same corresponding values for model and prototype.
• i model = i prototype for i = 1 to n • Enables extrapolation from model to full scale • However, complete similarity usually not possible. Therefore, often it is necessary to use Re, or Fr, or Ma scaling, i.e., select most important parameter and accommodate others as best possible.
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Principles of Model Testing Dimensional Analysis and Similarity Geometric Similarity: The model must be the same shape as the prototype. Each dimension must be scaled by the same factor.
Kinematic Similarity: Velocity as any point in the model must be proportional Dynamic Similarity: All forces in the model flow scale by a constant factor to corresponding forces in the prototype flow.
Complete Similarity: is achieved only if all above three conditions are met.
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Principles of Model Testing Dimensional Analysis and Similarity Complete similarity is ensured if all independent groups are the same between model and prototype. What is ? We let uppercase Greek letter denote a nondimensional parameter, e.g., Reynolds number Re, Froude number Fr, Drag coefficient, CD, etc. •Consider automobile experiment •Drag force is F = f (V, , L) •Through dimensional analysis, we can reduce the problem to
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Principles of Model Testing Flow Similarity and Model Studies Example: Drag on a Sphere
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Principles of Model Testing Flow Similarity and Model Studies Example: Drag on a Sphere For dynamic similarity …
….then ….
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Principles of Model Testing Scaling with Multiple Dependent Parameters Example: Centrifugal Pump Pump head:
h=g1(Q, wD, )
Pump power:
P=g2(Q, wD, )
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Principles of Model Testing Similitude-Type of Similarities Geometric Similarity: is the similarity of shape.
Lp Lm
Bp Bm
Dp Dm
Lr
Where: Lp, Bp and Dp are Length, Breadth, and Diameter of Prototype and Lm, Bm, Dm are Length, Breadth, and Diameter of Model.
LR= Scale ratio
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Principles of Model Testing Similitude-Type of Similarities Kinematic Similarity: is the similarity of motion.
V p1 Vm1
Vp 2 Vm 2
Vr ;
a p1 am1
ap2 am 2
ar
Where: vp1& vp2 and ap1 & ap2 are velocity and accelerations at point 1 & 2 in prototype and vm1 & vm2 and am1 & am2 are velocity and accelerations at point 1 & 2 in model. Vr and ar are the velocity ratio and acceleration ratio
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Principles of Model Testing Similitude-Type of Similarities Dynamic Similarity: is the similarity of forces.
Fi p Fv p Fg p Fi m Fv m Fg m
Fr
Where: (Fi)p, (Fv)p and (Fg)p are Inertia, Viscous and Gravitational Forces in Prototype and (Fi)m, (Fv)m And (Fg)m are Inertia, Viscous and Gravitational Forces in Model. Fr Is The Force Ratio
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Principles of Model Testing Flow Similarity and Model Studies Scaling with Multiple Dependent Parameters Example: Centrifugal Pump
=
Head coefficient:
,
Power coefficient:
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Principles of Model Testing Flow Similarity and Model Studies Scaling with Multiple Dependent Parameters Example: Centrifugal Pump (Negligible viscous effects) If
then
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then neglecting Reynolds number (viscous effects)
gh gh = ω D ω D
and
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Why Wind Tunnel Testing? Consider an aircraft flying in still or almost still atmosphere.
U
It is possible to get the aircraft performance from flight tests, but complicated and resource (time, money, human resources etc.) consuming. Also only limited information can be obtained. Alternatively a still aircraft model can be tested in a wind tunnel with air flowing at a speed of U. This reduces the resources required and more information can be obtained. A wind tunnel can be used to test models of aircraft and other vehicles and components of aircraft (ex. wings). Alternative testing resources: Whirling arms (no longer used), Water tunnels Can CFD replace EFD (wind tunnels)?
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CFD vs. EFD Computational Fluid Dynamics
Experimental Fluid Dynamics
Requirements
High end computer, software, printers, plotters etc.
Experimental facility (Wind tunnel), Instrumentation etc.
Manpower
One or two for small CFD work, A small team for large CFD work, each member specializing in grid developing, solver, post processing etc.
Usually requires moderate number of people for small EFD for preparing and installation of model, for measurements etc. Requires large number of people with different skills for large experiments.
Initial Cost Time required
Moderate
High
Moderate for computation, but large for post processing
Large for preparation, small for data acquisition and post processing
Information acquired
Hugh
Limited
CFD and EFD are complimentary. Both require careful working. Needs verification and validation, Extensive CFD can be carried out followed by limited comparison with EFD
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National Aerospace Laboratory Trisonic Wind Tunnel Facility
Test section: 1.2mx1.2m
Operation: Intermittent blowdown
Mach Number Range: 0.2 to 4.0
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Test duration: 40 secs (Typical)
Reynolds number: 8 to 60x106 per meter
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Lockheed Martin High Speed Wind Tunnel Circuit Layout Mach Number Range: Transonic range:
0.3 to 1.8
Supersonic range: 1.6 to 4.8
Reynolds Number Range 4 to 34x106/foot
Equivalent Airspeed 150 to 1,200 knot (280 to 2200 kmph) (144 to 1132 mps)
Run Time: 15 to 110 secs AE 2751
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NASA Ames Wind Tunnel World largest wind tunnel Test sections: 40’x80’& 80’x120’ Speeds:
300 knots 100 knots 556 kmph 185 kmph 154 mps 51 mps
80’x120’ test section can test full size Boeing 737 aircraft at velocities up to unit Reynolds numbers of 1.1x106/ft.
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F/A-18 Fighter Aircraft Testing in NASA Ames Wind Tunnel
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Classification of Wind Tunnels I
Based on Size: Small: Test section ˜ 0.3 mx0.3 m, mainly for college instructional purposes Medium: Test section ˜ 1 mx1 m, used for university research purposes Large: Test section ˜ 2 mx2 m, used for testing industries and research laboratories Very Large: Test section ˜ 4 mx4 m, used for testing large models in industries and research laboratories (for ex: NASA Ames Research Laboratory)
II Based on Wind Velocity: Low velocity: Velocity less than 100 m/s, incompressible flow used in instructional and university research purposes Subsonic velocity: Compressible flow (M5) used for industry and research laboratory research purposes. Rarefied gas with very low density.
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Classification of Wind Tunnels III Based on Test Section Configuration: Open Circuit: Used mainly for small and intermediate tunnels. Test section may be without solid boundaries (open jet) or with solid boundaries (closed jet) Closed Circuit: Used mainly for large and very large tunnels, may have closed or open (Fluid Mechanics Laboratory, Department of Applied Mechanics, IIT Madras). May have single return, double return or annular return.
IV Based on Operation: Continuous: Most of the tunnels operate continuously. Short Duration: These tunnels usually operate for very small time (Intermittent: A few minutes, Blow down: a few seconds) These types of tunnels are used in high speed turbomachinery testing. Saves enormous amount of energy for testing.
V Special Tunnels: High Reynolds Number Tunnels Low Turbulence Tunnels: The tunnel test section usually have a turbulence level of 1%. Certain applications require very low turbulence levels ˜ 0.1% (MTL wind-tunnel at the Department of Mechanics, KTH, Sweden) Variable Density Tunnels: Independent variation of Mach and Reynolds numbers. Cryogenic tunnels.
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Applications of Wind Tunnels I
Aeronautical applications Most of wind tunnels are used for aeronautical applications for instructional purpose, for research carried out at universities, industries and research laboratories. The research carried out may be for fundamental understanding of fluid flow phenomena or for developmental testing of components for aircrafts. Even moderate sized models of actual aircraft are tested in large wind tunnels. Aeronautical wind tunnels can be further classified as high Reynolds number wind tunnels, V/STOL wind tunnels, Free-flight wind tunnels, Spin tunnels or vertical wind tunnels, Stability tunnels, Propeller tunnels, Propulsion tunnels, Icing tunnels, Low turbulence tunnels, Two-dimensional tunnels.
II Smoke tunnels III Automobile wind tunnels IV Aeroacoustic wind tunnels (Anechoic tunnels)
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Applications of Wind Tunnels V Water tunnels Used for flow visualization studies, underwater vehicle development and for cavitation studies. Usually small size and works at low velocities to obtain realistic Reynolds numbers.
III General purpose wind tunnels Used for study of people drag (bike racers, skiers etc.), birds and insects, wind power devices solar collectors, radar antennas and satellite television receivers, sails and above-water parts of ships, bridges, etc.
IV Environmental or meteorological wind tunnels These tunnels are designed to simulate Earth’s natural boundary layer, which typically has a thickness of 300 to 600 m. They are used for study of static loads and associated experiments on buildings, dynamic loads on buildings, unsteady aerodynamics in wind engineering, agricultural/wind breaks, agricultural/agronomy, agricultural/soil erosion breaks, snow drifting, evaporation and related issues, pollution dispersion, etc. The reader is referred to Barlow J. B., Rae Jr. W. H. and Pope A., “Low Speed Wind Tunnel Testing”, IIIrd Edition, John Wiley & Sons, Inc.,1999 for some more details of the above wind tunnels. Indian edition of this text book is also available at about Rs. 1,000/-.
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Applications of Wind Tunnels
Snowdrift study for lodge, near building in photo, at a ski resort. Note the effect of tower at left rear, which provides self-removal of snow from entrance with prevailing from left.
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Skier in wind tunnel
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Applications of Wind Tunnels
Effects of wind barriers on solar collectors being simulated in a wind tunnel Above: Efflux velocity is equal to wind speed and slack height is 1.5 times building height. Building is downstream of stack. Below: Same conditions as above except building is upstream of the stack simulated in a wind tunnel
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Wind Tunnel Layouts Open Circuit Wind Tunnel
Plan view of an open circuit wind tunnel (Daimler-Benz Aerospace Airbus, Bremen, Germany)
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Wind Tunnel Layouts Open Circuit Wind Tunnel The following are the advantages and disadvantages of an open circuit wind tunnel Advantages 1) Construction cost is typically much less. 2) If one intends to run internal combustion engines or do extensive flow visualization via. smoke, there is no purging problem provided both inlet and exhaust are open to the atmosphere.
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Wind Tunnel Layouts Open Circuit Wind Tunnel Disadvantages 1) If located in a room, depending on the size of the tunnel to the room size, it may require extensive screening at the inlet to get high quality flow. The same may be true, if the inlet and/ or exhaust is open to the atmosphere, when wind and cold/hot weather can effect operation. 2) For a given size and speed, the tunnel will require more power to operate. This is usually a factor only if used for development experiments, where the tunnel has high utilization rate. 3) In general, open circuit wind tunnels tend to be noisy. For larger tunnels (test sections of 6 m2 and more), noise may cause environ mental problems, limit hours of operation, and/or require extensive noise treatment of the tunnel and surrounding room. Because of low initial cost, an open circuit wind tunnel is often ideal colleges and universities, where a tunnel is required for class room instruction purposes and research and high utilization Is not required. Open circuit design are also frequently used by science fair participants who build their own wind tunnels.
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Wind Tunnel Layouts Closed Circuit Wind Tunnel
Plan view of a closed circuit wind tunnel (Defense Establishment Research Agency, DERA, 13x9 ft (3.9x2.7 m=10.9 m2) tunnel in Bedford, England AE 2751
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Wind Tunnel Layouts Closed Circuit Wind Tunnel Instrumentation stand
Open test section: 1 m dia.
Closed circuit (return flow) open test section wind tunnel, Fluid Mechanics Laboratory, Department of Applied Mechanics, IIT Madras AE 2751 Wind Tunnel Techniques July-Nov. 2016
Wind Tunnel Layouts Closed Circuit Wind Tunnel The following are the advantages and disadvantages of a closed circuit wind tunnel
Advantages 1) Through the use of corner turning vanes and screens, the quality of the flow can be well controlled and most important will be independent of other activities and in the building and weather conditions. 2) Less energy is required for a given test-section size and velocity. This can be important for a tunnel used for developmental experiments with high utilization (two or three shifts, five to six days a week). 3) There is less environmental noise when operating.
Disadvantages 1) The initial cost is higher due to return ducts and corner vanes. 2) If used extensively for smoke visualization experiments or running of internal combustion engines, there must be a way to purge tunnel. 3) If tunnel has high utilization, it may have to have an air exchanger or some other method of cooling.
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Wind Tunnel Layouts Open or Closed Test Section?
An open test section in conjunction with an open circuit wind tunnel requires an enclosure around the test section to prevent air being drawn into the tunnel from the test section rather than the tunnel inlet.
For closed return wind tunnels of large size with an external balance, the open test section tends to have one solid boundary, since the balance must be shielded from the wind.
This is an anomaly for aircraft experiments other than takeoff and landing, but it is a natural condition for experiments on automobiles or surface-borne marine vehicles.
Many open test-section, closed circuit wind tunnels have experienced severe flow fluctuation problems that require extensive post construction diagnostics and corrective actions.
One of the tunnels currently acknowledged to be one of the most useful tunnels in existence nevertheless has had substantial difficulties with unsteady flow and noise when running in the open test-section configuration.
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Wind Tunnel Layouts Open or Closed Test Section?
The most common geometry is a closed test section, but a wide range of tunnel geometries have provided good experimental conditions, once the tunnel idiosyncrasies have become known to the operators and users.
Slotted wall test sections are becoming more common as are test sections that can be converted among two or more configurations.
A rectangular test section is preferable in larger size wind tunnels, as it is easier to change a model when working off a flat surface.
Further, if automobile or other ground vehicle experiments are to be conducted, a flat floor is a requirement.
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Wind Tunnel Design General Layout: Open Circuit and Blower Tunnel Open Circuit: This type of tunnel is widely used for instructional purposes and for investigations of fundamental flow phenomena. The test section is closed type (a).
Blower Tunnel: This type of tunnel also is widely used for instructional purposes and for calibration of flow devices. A characteristic design problem for these facilities is choice of details of a wide-angle diffuser between the fan or blower and the settling chamber before the contraction. The test section is open type (a).
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Wind Tunnel Layouts Test Section Size
In general it may be expected that the test section should have as large a cross-sectional area as possible. Ideally, a tunnel would be large enough to handle a full-scale vehicle (aircraft, car etc.).
In fact, several tunnels were built in 1920s and through the 1940s to achieve this goal for aircraft.
However, since World War II era, and presumably in the future, the size of the aircraft have become such that wind tunnels to accommodate full-scale vehicles are not practical.
If one uses the rule of the thumb that the model span should be less than 0.8 of the tunnel width, then Howard Hughes’ Hercules, or as more popularly known, the “Spruce Goose”, which was designed and built in the 1940s with a 320 ft wing span, would require a test section 400 ft wide.
The cost of building and operating a tunnel of this size is staggering to contemplate.
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Wind Tunnel Layouts Test Section Size
The cost of building a model, transporting it, and erecting in the tunnel, as well as making changes during an experimental program, would also be an interesting, albeit expensive task.
Thus, it is apparent for the larger of modern aircraft, the concept of a tunnel to accommodate full-size aircraft is out of question based on costs and practical difficulties.
Recalling from the earlier discussions of flow similarity, it is more important to seek to obtain Reynolds numbers for the model experiments that are as nearly as possible to the full-scale values than to be concerned with size alone.
In practice most development are done in tunnels with widths from 10 to 20 ft (3 to 6 m).
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Wind Tunnel Design Basic Considerations
The first step in the design of a wind tunnel is to determine the size (small, intermediate, large or very large) and shape of the test section (square, octagonal, circular or rectangular), based on the intended uses of the tunnel.
The details given address tunnels for which the primary use is vehicle and vehicle components.
A major part of the testing will be force testing, where information is sought for performance, fuel efficiency, stability or control of a vehicle that may be an aircraft, an automobile, a racing car, a submarine, a racing yacht, or possibly others..
The cross-sectional area of the test section basically determines the overall size of the wind tunnel.
The size of the wind tunnel will be the primary factor in determining the structural or shell costs, and the power and operating hours will determine the energy portion of the operational cost.
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Wind Tunnel Design Basic Considerations
The major cost of operation is the salaries of the tunnel personnel.
But the electrical energy cost to run the tunnel and its auxiliaries is not an insignificant cost and it will increases in the long run.
The details given address tunnels for which the primary use is vehicle and vehicle components.
There must be a balance between initial costs and operating cost.
In the past many tunnels have been built with short diffusers and related features to allow short circuit length to hold down initial cost while accepting higher energy costs of operation.
This trade-off should be carefully examined with due consideration given to anticipated energy costs, which are likely to increase.
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Wind Tunnel Design Overall Aerodynamic Objective
The overall aerodynamic objective of most wind tunnels is to obtain a flow in the test section that is as near as possible to a parallel steady flow with uniform velocity throughout the test section.
Perfection is not possible so every design is bounded by constraints that include maximum cost, available space, available time, and available knowledge.
It is almost always desired to obtain the largest size of the test section and the highest speed for the available funds.
High speed and large size are of course competing demands.
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Wind Tunnel Design Basic Decisions: Unit Reynolds Number
A central issue in the sizing of a low-speed wind tunnel is the achievable Reynolds numbers for the models that can be accommodated.
The same question arise for vehicle developers who must select from available wind tunnels one which to carry out the tests in a development program. Reynolds number, Re=r V L/m .
∞ ∞
∞
A maximum Mach number of 0.3 is chosen as the upper limit of Mach number for the free stream beyond which the effect of Mach number are to be considered. Considering sea level standard atmospheric conditions, the maximum V will be about 100 m/s and the unit Reynolds number (for unit length) will be about 2.1x106 ∞ ft-1 (7x106 m-1).
These numbers and the appropriate characteristic length of the test article (or tunnel width) give a good approximation to the available Reynolds number in an atmospheric wind tunnel.
Hence there has been and continues to be much attention focused on obtaining effective results with less than full-scale articles or with various separate components tests.
An important example of the contribution of component testing is the development of airfoil profiles for various profiles that are then incorporated into three-dimensional wing designs.
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Wind Tunnel Design Basic Decisions: Unit Reynolds Number
Another widely used method is to test half models since many vehicles have a plane of symmetry.
Most important of all is the careful study of aerodynamic phenomena as they are affected by variation of Reynolds number so that useful conclusions can be obtained from tests that do not duplicate the operating Reynolds number.
For many studies it is not necessary to produce the full-scale Reynolds number, but it must be of a “reasonable” value.
For vehicles including aircrafts and racing automobiles that can operate at speeds of Mach 0.3 or greater in the atmosphere, test articles would have to be atleast full scale to achieve operational Reynolds numbers in an atmospheric wind tunnel.
This is impossible or very costly for many vehicles.
Much low speed testing involves aircraft takeoff and landing configurations, where the Mach number is typically in 0.15 to 0.3 range.
Both the lift curve slope and maximum lift coefficient are affected by Mach numbers as low as 0.2.
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Wind Tunnel Design Basic Decisions: Unit Reynolds Number
This tends to require a tunnel speed approximately equal to the full scale landing speed. In an unpressurized tunnel using air, this means that the Reynolds number ratio of model to full scale is approximately equal to the size ratio between the scale model and the aircraft. A primary decision is the choice of the minimum acceptable value of Reynolds number. Because much of low-speed testing is at high-lift conditions, the effect of Reynolds number on the airfoils at high lift must be considered. Maximum lift and lift curve shape near stall for single-element airfoils vary considerably with Reynolds number up to at least one million. For multi element airfoils, this range is much greater. The continuing need for testing facilities that allow near-full-Reynolds number of transport aircraft high-lift systems is a primary reason for serious consideration of construction of major new wind tunnels in the mid-1990s. In any case, the lower boundary for testing airfoils and wings for vehicles that will operate at higher scale values is a Reynolds number in the range of 1 to 1.5 millions based on chord.
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Wind Tunnel Design Basic Decisions: Unit Reynolds Number
At these values of Reynolds number, the model is likely to have an extensive region of laminar flow, and the possibility exists of poor simulation owing to separation of the nodel’s laminar boundary layer. It is assumed that laminar separations are likely to occur at full scale in the normal operating conditions. Therefore, flow similar to full scale is more likely to be achieved by the transition location on the model. If the Mach number is taken as 0.2, then the tunnel velocity is about 240 kmph (70 m/s) For this speed, the unit Reynolds number is a little less than 1,500,000 ft-1 (450,000 m-1). Although the minimum Reynolds number can not be rigidly defined, the above rationale has been used to define a minimum Reynolds number of between 1,500,000 and 2,500,000 based on wing chord for low-speed tunnels to be used for aeronautical development testing. There are some flight vehicles that operate at lower Reynolds numbers. There is, in fact, a whole series of airfoils for soaring gliders that are especially designed to operate at Reynolds numbers below 1,000,000.
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Wind Tunnel Design Basic Decisions: Unit Reynolds Number
There are also an increasing number of small low-speed unmanned vehicles (MAVs and UAVs).
Operational Reynolds numbers for these aircraft are often obtainable in a medium-sized wind tunnel.
For vehicles that operate in the atmosphere at speeds such that the Mach number is less than 0.3, the operational Reynolds number can be duplicated in an atmospheric wind tunnel with a scaled model.
Consider a production automobile at 60 mph (30 m/s).
The operating Reynolds number can be obtained using a three-eighth scale model with an atmospheric tunnel test speed of 160 mph (80 m/s)
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Wind Tunnel Design Basic Decisions: Test Section Size
This is commonly the starting point in the design of a wind tunnel.
The choice will follow from considerations of the desired Reynolds number capability, the budget for tunnel construction, and the costs for tunnel operation and tunnel users as reflected particularly in required model characteristics.
These characteristics are discussed in the next slide.
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Wind Tunnel Design Basic Decisions: Test Section Size Airfoil models: For a mean geometric chord of 0.3 m and aspect ratio of 8-9 for an aircraft wing, the span is about 2.5 m. The maximum span should be less than about 0.8 of the tunnel width due to effects of tunnel walls on the flow, which leads to a width of about 3 m. For rectangular solid-wall tunnels, wall correction factors for a small wing will be minimum for a width-to-height ratio of about 1.5. These considerations indicate why so many tunnels have been build in the 7x10 to 8x12 ft (2x3 to 2.4x3.6 m) size range with the maximum speed in the range of 200-300 knots (370 to 550 kmph or 100-150 m/s). In addition to the consideration of flow properties, a model for a tunnel of this size is large enough so that the smaller parts are relatively easy to fabricate.
V/STOL Aircraft: For V/STOL models in a STOL descent case, the speed will be near 70 knots due to model power limits or tip Mach number on propellers and rotors. The reduction in in test speed will require a larger model to maintain reasonable Reynolds numbers. To minimize the wall corrections due to large downwash angles from these models, the modelspan-to-tunnel-width ratio must be smaller, typically between 0.3 and 0,5. Thus the V/STOL tunnels built in the 1960s have test section that are 20-30 ft (6 to 9 m) wide.
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Wind Tunnel Design Basic Decisions: Test Section Size Automobile: A key issue for automobile tunnels is the blockage based on frontal area. The flow around automobiles is often more characteristic of “bluff bodies” than of “streamlined bodies”. This means there is almost a sizable region of separated flow. The wind tunnel test sections needs to be sufficiently long so that e separated flow regions “close”, before end of the test section and the entry of the diffuser. Otherwise the pressure in the separated region will not be correct and a large influence on drag will exist. In addition, the length-to-width ratio of automobiles is greater than for aircraft while the width-to-height ratio is much less. A wind tunnel test section sized for automobiles is therefore typically longer than a test section sized for aircraft and the width-to-height ratio approximates the width-to-height ratio of standard automobile. Ideally the blockage, the ratio of model frontal area to test-section area, will be about 5% or less.
Keels, Submarines and Sails: Submarines and surface ships have large length-to-width ratios. Their operational “leeway” is smaller than that for angle of attack or side slip for aircraft. A test section chosen for these vehicles would have a length-to-width ratio of 2 or greater. Yacht keels are sufficiently similar to aircraft wings that they fit comfortably in aircraft test sections. Test configurations that include both keels and rudders would be better accommodated in test sections with higher values of length-to-width ratio.
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Wind Tunnel Design Basic Decisions: Test Section Size Instructional and Other Small Tunnels: For small research tunnels and student tunnels at universities and elsewhere, the prospect of achieving or even approaching operational Reynolds numbers is usually beyond the available budget. In many instances the problem of building models accurately may be a critical factor. Assuming that students can hold an airfoil dimension of 0.25 mm and that it is sufficient to hold the model to 1-2% tolerance, the following results can be reached. For a 12% thick airfoil with 2% tolerance, the maximum thickness equals 1.2 cm and the chord is 10 cm. Using a mean chord of 10 cm and aspect ratio of 8. the span is 0.8 m. AS the maximum span is 0.8 of the tunnel width, the width is 1 m. Using a widthto-height ratio of 1.5 for a minimum wall correction factor, the height is 0.7 m. The cross sectional area is then 0.7 m2. A minimum test velocity would be about 30 m/s or a dynamic pressure of about 50 mm of water (or 50 kgf/m2). Because of lower speeds and corresponding dynamic pressures, instrumentation sensitivity is very low. These instruments may not be available or very costly. Many demonstration tunnels and calibration tunnels have test sections areas of 0.01 m2 or less. These tunnels are very useful for observation of basic flow phenomena and calibration of instrumentation probes. However, they are of little use for doing vehicular component studies except in cases for which Reynolds number are quite low (ex: MAV and UAV).
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Wind Tunnel Design Basic Decisions: Test Section Size For a rectangular tunnel, the width determine the model size and the Reynolds number at affixed speed. The cost of the tunnel shell and its required power tend to vary with the square of the test section width. Since funds for a tunnel are usually fixed, the largest tunnel that the funds will buy is generally built. The size of smaller tunnel is frequently determined in the final analysis by the size of the room that will house the tunnel.
Special tunnels to increase unit Reynolds number: The unit Reynolds number can be increased by building either a pressure tunnel or a cryogenic tunnel using a cold gas such as nitogen. Other working fluids such as Freon and sulfur hexafluoride have beeb used to obtain desired conditions. These are special purpose tunnels, and the need for their special capabilities must justify the cost as with any other tunnel. The time required for model changes will be long unless special provisions are made because the test section must be isolated before workers can enter to work on the model. Test productivity, as well as flow characteristics, s an important characteristic of a wind tunnel.
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Wind Tunnel Design Open or Closed Return
Another basic design consideration is whether the tunnel will be of return or non return (open circuit) type.
Almost all of the small research tunnels are of the non return type, usually because of the lower construction cost.
Power consumption for such tunnels is usually not a significant factor in overall cost.
Although there exists some larger tunnels of non return design, most of the larger tunnels are of the return types, the majority being single return.
A few of the earlier tunnels are of double return type.
However, it has been more than 70 years since the double return design has ben chosen for a new wind tunnel.
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Wind Tunnel Design Open or Closed Test Section
Open jet tunnels will have a lower energy ratio than a closed jet wind tunnel owing to the jet entraining stagnant air as it passes from the contraction cone exit to the collector inlet.
If the tunnel has an external balance, the balance usually has to be shielded from the air jet and one of the boundaries tends to be closed.
Open throats do not work for an open circuit tunnel with a propeller in the diffuser unless the test section region is enclosed in an air tight plenum or room.
Open throat tunnels offer suffer from pulsations similar to vibrations in organ pipes.
An open throat gives easy access to the model in small tunnels.
In large tunnels scaffold of some type is required to gain access to the model.
The setting up and removal of the scaffold require additional model change time.
Since the jet length is usually kept short to reduce losses, there is the possibility that highlift models may deflect the wake enough to miss the collector or that the wake of a bluff body will interact with collector.
An open jet provides easier acess for traversing devices to move instrumentation to any point in the flow.
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Wind Tunnel Design Open or Closed Test Section
In general, the advantages appear to be with the closed throat tunnel for aeronautical testing.
However, considerations of bluff body aerodynamics and overall size requirements have led to a number of open, partially open or slotted wall, and convertible jet tunnels built by automobile companies and others who engage in automobile test work or V/STOL aircraft development.
These arrangements provide greater flexibility in uses of a tunnel provided the staff are sufficiently knowledgeable about the variety of wall effects that must be understood.
Closed thrat tunnels that are vented to the atmosphere not at the test section but at another location will have the test section below atmospheric pressure.
Thus they can suffer from leaks either through holes cut in walls for probes, wires, pipes, and so on, or through the struts required to mount the model.
These tunnels usually have a sealed room or plenum around the test section. Most small open circuit tunnels are not built this way and suffer from leaks.
This makes wood an ideal material for such tunnels because it is easy to patch.
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Wind Tunnel Design General Layout: Closed Return The general layout for closed return tunnels has reached a form generally agreed upon for reasons of construction economy and tunnel efficiency. Starting with the test section and going down the common configuration includes the following elements: a) The test section, which may be closed, open, partially open or convertible. The testsection-length-to-hydraulic-diameter ratio may typically be chosen to be 2 or more, in contrast to the shorter test sections of earlier era tunnels. b) A diffuser of at least three or four test section lengths. The typical equivalent cone angle is in the range of 2-3.5O with the smaller angles being more desirable. The area ratio is typically 2-3, again with the smaller values being more desirable. c) “First corner” incorporating turning vanes. d) Second leg that may continue the diffuser or may be constant area. e) Safety screen to prevent parts of failed models or other unintended flying objects reaching the fan. This screen is usually just ahead of the second-corner turning vanes. f) “Second corner” incorporating turning vanes that may be essentially copies of the first corner vanes solely to gain a small engineering and construction cost reduction.
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Wind Tunnel Design General Layout: Closed Return g) Transition from rectangular to circular cross section to take flow into the fan. h) Fan and straightener section. Other drive devices such as ejector have also been used. i) Return or second diffuser. This will commonly incorporate a transition back to rectangular from the circular cross section at the fan exit. The transition will likely begun in the straightner section. The second diffuser should follow similar design guides as the diffuser. j) “Third corner” incorporating turning vanes. k) Third leg that may be constant area. l) Heat exchanger. m) “Fourth corner” incorporating turning vanes that may be copies of the third corner vanes. n) Wide angle diffuser with separation control screens. Typical properties are angles of about 45O and area ratios of 2-4. o) Settling area. p) Flow conditioners typically including flow straighteners and turbulence control screens. q) Contaction or nozzle. Typically area ratios are in the range of 7-12, although lower and higher values are not uncommon.
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Wind Tunnel Design General Layout: Closed Return A representative layout of a closed return tunnel with notation is shown below:
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Layout of a closed single-return wind tunnel Wind Tunnel Techniques
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Wind Tunnel Design General Layout: Closed Return
The plane of the return passage is almost always horizontal to save cost and make the return passage easier to access.
A vertical return is justified only when space is at a premium and has only been used for small sized tunnels.
A concept for obtaining economics in wind tunnel construction by producing a series of various sizes for which there are many common components was developed and used.
The first- and second-corner assemblies for one size tunnel would be the same as the third- and fourth-corner assemblies for the next smaller size tunnel.
The fan diffuser for one size tunnel would be the same as the test section diffuser for the next smaller size tunnel.
The engineering and construction drawings would be the same for all sizes except for the specified scale for each instance.
There are many innovative configurations to accommodate particular needs that are quite different from
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Wind Tunnel Design General Layout: Open Circuit and Blower Tunnel Open Circuit: This type of tunnel is widely used for instructional purposes and for investigations of fundamental flow phenomena.
Blower Tunnel: This type of tunnel also is widely used for instructional purposes and for calibration of flow devices. A characteristic design problem for these facilities is choice of details of a wide-angle diffuser between the fan or blower and the settling chamber before the contraction.
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High Speed Wind Tunnels Definition of High Speed When compressibility effects are pre dominant the flow is generally said to be of high speed. A lower limit is approximately M=0.5. Power requirements vary as cube of velocity in the wind tunnel. This does not hold into the high speed regime exactly. Because of large power requirements, high speed wind tunnels are usually of the intermittent type.
This material is taken from the course notes of Experimental Aero (Gas) dynamics by Prof. Job Kurian of Department of Aerospace Engineering, IIT Madras
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High Speed Wind Tunnels Types of High Speed Wind Tunnels 1 Continuous (for all speed ranges) 2 Intermittent 2.1 Blowdown: M > 0.5 < 5.0 2.2 Indraft 2.3 Intermittent pressure vacuum tunnel for M>5
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High Speed Wind Tunnels Pressure Driven Blow Down Wind Tunnel .
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High Speed Wind Tunnels Indraft Type Wind Tunnel
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High Speed Wind Tunnels Comparison between Indraft and Pressure Driven Wind Tunnels Indraft wind tunnels
Pressure driven wind tunnels
Stagnation temperature at supply condition Reynolds number can be varied at a particular Mach no. is constant during a run. So also is total pressure. No fluctuations as those generated by a pressure regulator. No possible contamination such as that due Cost is much less than of an indraft tunnel. to oil. Vacuum is safer to handle than pressure. Pressure regulators are not needed. The wind tunnels described above can be converted as continuous tunnels. The comparison between blow down and continuous wind tunnels are given in next slide.
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High Speed Wind Tunnels Comparison between Intermittent and Continuous Wind Tunnels Intermittent (blow down) wind tunnels
Continuous wind tunnels
Simple to design and less costly
More in control of conditions and return to a given test condition with more accuracy.
A single drive may run several tunnels
Check points are easily obtained No panic of rapid testing Test conditions can be held constant for a longer time.
Model testing is more convenient Extra power is available to start Failure of model will not result in tunnel damage
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High Speed Wind Tunnels Supersonic Wind Tunnels: Introduction .The nozzle regulates the speed of air entering the test section of the wind tunnel so that the desired Mach number is established. Mach number is uniquely determined by the area ratio of the nozzle. A well designed nozzle makes the flow parameters uniform across the cross section. The design of a suitably shaped nozzle contour to obtain the desired uniform flow at the nozzle exit is based on the method of characteristics.
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters Dynamic pressure: The local dynamic pressure 1/2ρv2 can be related to the local Mach number and static pressure.
Mass flow rate: The mass flow rate is given as m=ρvA=ρ*a*A* where * represents choked or sonic conditions (M=1)
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters .
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters Writing in terms of stagnation conditions
.
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters . flow density is expressed as a function of stagnation conditions and the area ratio. Mass For isentropic flow, A* is a constant and A/A* is a unique function of local Mach number.
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters . Test section velocity;
Maximum velocity;
or
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is maximum when cpT=0
cp-cv=R cp/cv=g cp-cp/g=cp(1-1/g)=(cpg-cp)/g=cp(g-1)/g=R cp=gR/(g-1)
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters The test section flow velocity v for a given stagnation temperature TO approaches the maximum value vmax at relatively low supersonic Mach numbers. For example, in the case of TO = 300 K R = 287J/kgK and a test section Mach number of 5.0, the ratio of v/vmax can be calculated to see that it is equal to 0.913. This means at ordinary stagnation temperatures, the velocity in the test section reaches 91% of the maximum possible velocity corresponding to the total energy of the fluid. The stagnation temperature T0 rather than the Mach number which is important to attain high velocities.
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters Free stream Reynolds Number (Re):
Re=ρvL/μ
Experimental observation is that μ is independent of pressure in the range of 0.001 to 20 atmospheres.
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters Free stream Reynolds Number (Re):
Re=ρvL/μ
If this relation is assumed then the free stream Re can be expressed as a function of M1, the test section Mach number and of the stagnation parameters, Reynolds number per unit length Re/L=ρv/μ are expressed in stagnation quantities as given below.
M=v/a
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High Speed Wind Tunnels Supersonic Wind Tunnels: Test Section Parameters .
n= 0.768 for air Simplifying Both ao and μo are functions of stagnation temperature. Both increase with temperature. Hence, appreciable changes in free stream Re/unit length for a given M can be obtained only by varying stagnation density.
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High Speed Wind Tunnels Components of a Supersonic Wind Tunnel Air storage tanks: Size of the storage will be dependent on the mass flows required . and the frequency of runs. Pressure storage tanks are available on the shelf basis – They are mounted horizontally or vertically. Tanks are painted black to absorb heat. They are provided with safety disk or pressure relief valve. As air is drawn from the storage, polytropic expansion takes place within the tank. This results in drop of reservoir temperature which is very bothersome. Fall of stagnation temperature causes resultant change in the stream temperature for a given Mach number. Change in temperature results in the change of viscosity which in turn affects the boundary layer thickness. Changes in Reynolds number and Mach number during a run are thus consequential to the fall in reservoir temperature. To maintain constancy of stagnation temperature, it is a practice to stack the reservoir volume with empty metallic cans. They serve as heat storing matrix during compression and release heat during the expansion process. Another way to maintain the constant stagnation temperature is by providing heater units in the reservoir.
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High Speed Wind Tunnels Components of a Supersonic Wind Tunnel Settling chamber /wide angle diffusers: Wide angle diffusers lead the flow to the . settling chamber. Arrangements for leading the flow to the settling chamber may be by one of the methods shown below:
a, b: Wide angle diffusers
c: Reverse entry into the settling chamber
Uniformity of flow in the test section is improved if a large area ratio contraction is provided.
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High Speed Wind Tunnels Components of a Supersonic Wind Tunnel Convergent-divergent (CD) nozzle: The CD forms the heart of the supersonic . wind tunnel .For generating supersonic flow in the test section, it is essential that there is a c-d nozzle in the tunnel circuit before the test section. The area ratio of the CD nozzle (Aexit/Athroat) uniquely decides the Mach number
Convergent-divergent nozzle and the pressure profile
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High Speed Wind Tunnels Components of a Supersonic Wind Tunnel Convergent-divergent (CD) nozzle: When the tunnel .
3
This is called the choked condition of the nozzle.
2
Isentropic flow, g=1.4
operation starts, the flow is initiated in the nozzle as subsonic and reaches the sonic Mach number at the throat when sufficient mass flow is allowed.
Under this condition, maximum mass flow rate for the given stagnation conditions takes place through the nozzle. The ratio between the upstream stagnation pressure (PO) and the downstream back pressure (Pb) corresponding to the first time choking is called the first critical pressure ratio of the nozzle. At this pressure ratio, the flow in the divergent part of the nozzle is subsonic. The exit Mach number will be the subsonic value corresponding to the A/A* shown in the figure.
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A A* 1
0 0
M1
2
3
Area ratio, A/A* vs. Mach number
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High Speed Wind Tunnels Components of a Supersonic Wind Tunnel Convergent-divergent (CD) nozzle: In this context A is the exit area and A* the . throat area of the choked nozzle.
As the value of PO/Pb is progressively increased, the flow in the divergent part of the nozzle accelerates to be supersonic but shocks are formed in the divergent part until a pressure ratio corresponding to the supersonic Mach number of the nozzle is reached. The pressure ratio corresponding to this Mach number is the third critical pressure ratio of the nozzle. Between the first and third critical pressure ratios shocks of varying strengths take place in the nozzle and outside of it as there is no other isentropic solution between the 1st and 3rd critical pressure ratios. The pressure ratio corresponding to the occurrence of a shock at the exit plane of the nozzle is the second critical pressure ratio of the nozzle.
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High Speed Wind Tunnels Components of a Supersonic Wind Tunnel . Convergent-divergent (CD) nozzle: All the shocks generated at the different
pressure ratios inside the nozzle will make the post shock Mach number subsonic and the subsonic nozzle exit pressure will be made equal to the ambient pressure in the remaining part of the diffusing divergent channel. Between the second and third critical pressure ratios, oblique shocks of varying strengths depending on the pressure ratio will be formed emanating from the nozzle lip. The physical purpose of these oblique shocks is equalization of pressures between the exit plane and the ambient. If the pressure ratio is increased beyond that corresponding to the third critical pressure ratio, expansion fans will be formed at the lip of the nozzle.
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High Speed Wind Tunnels Components of a Supersonic Wind Tunnel Diffuser-the . necessity of providing a diffuser
Take the case of a free jet facility shown in making use of a CD nozzle of Mach number 3, exiting to the ambient conditions at one atmosphere pressure.
In order to avoid shocks and expansion waves at the exit of the nozzle, pe must be pb.
(po/pe)M=3=36.7 Pressure ratio required will be 36.7 for a wave free exit flow from the nozzle
Nozzle of a free jet facility
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Wind Tunnel Techniques
July-Nov. 2016
High Speed Wind Tunnels Components of a Supersonic Wind Tunnel A constant area section is added to the nozzle exit. . The duct similar to the test section of a wind tunnel attached to the nozzle exhausts to atmosphere. pe corresponds to static pressure at the exit plane of the nozzle before the shock. Static pressure after the shock (p2) is equal to ambient pressure.
po/p2=(po/pe) (pe/p2)=36.77/10.33=3.55 In the equation above, pe/p2 represents the shock pressure ratio at M=3.0.
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Free jet nozzle with a test section
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July-Nov. 2016
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High Speed Wind Tunnels Components of a Supersonic Wind Tunnel In the third case, a divergent channel is provided after the constant area
.duct and the shock stands at the end of the constant area duct.
Free jet facility with test section and a diffuser
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Wind Tunnel Techniques
July-Nov. 2016
High Speed Wind Tunnels Components of a Supersonic Wind Tunnel .W hen M 1. If the entropy change is confined to the region between the two throats, the diffuser throat area ADT must be larger than the nozzle throat area ANT. Diffuser throat area must be large enough to accommodate the stagnation pressure loss of the strongest shock. A cooler is included prior to the compressor because compressor work is proportional to the intake temperature.
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Wind Tunnel Techniques
July-Nov. 2016
High Speed Wind Tunnels Closed circuit supersonic wind tunnel The practical operation of the closed circuit wind tunnel may be explained as follows:
.
As the tunnel is started, flow through it begins as subsonic and as the pressure ratio is increased the nozzle is choked. Further increase in the pressure ratio causes shock to be formed in the divergent section. At a pressure ratio corresponding to second critical pressure ratio, shock is formed at the exit plane of the nozzle which is same as the entry section to the test section. The formation of shocks during the starting process gives rise to fall in stagnation pressure. The total pressure after the shock is designated as p02. This necessitates that the diffuser throat is designed larger as decided by the ratio p02/p01. The ratio of diffuser throat area to the nozzle throat is in the inverse ratio of total pressures given above.
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July-Nov. 2016
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High Speed Wind Tunnels Closed circuit supersonic wind tunnel The ratio of areas for different test section Mach numbers calculated based on normal shock losses is given in the figure. This makes sure that the starting shock passes through the diffuser throat. The diffuser throat area calculated as above does not take in to account the non isentropy of frictional flows and only the shock losses are considered.
1.0
.
0.8
ADT ANT 0.6
1
3
5
7
Mach Number, M
9
Ratio of diffuser throat and nozzle throat for different test section Mach numbers
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July-Nov. 2016
High Speed Wind Tunnels Closed circuit supersonic wind tunnel
Shock movement from nozzle exit to diffuser Assuming a frictionless operation, the shock may assume any section in the constant area test section. But, the effect of friction is to make the shock unstable in the constant area duct. The shock that is generated during starting of the tunnel does not stay at the nozzle exit (entry to test section) but is moved downstream by the effect of friction.
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High Speed Wind Tunnels Closed circuit supersonic wind tunnel The starting pressure ratio minimum required to cater to the shock at the test section Mach number is corresponding to that for locating the shock at the nozzle exit. (worst shock).
rps=(po1/po2)worst hock As the starting pressure ratio is maintained, the starting shock which moves down stream can be stable only at an area equal to that of the test section. In the convergent part of the diffuser the Mach number will be less than that in the test section. With the value of starting pressure ratio which produced the shock at the test section Mach number being maintained, shock losses at that Mach number is being catered to. Hence the starting shock stabilizes only in the diverging part of the diffuser at a section where there is equal area and Mach number as the test section. The starting shock crossing the diffuser throat and remaining in its divergent part is called the ‘swallowing of the starting shock’.
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Wind Tunnel Techniques
July-Nov. 2016
High Speed Wind Tunnels Closed circuit supersonic wind tunnel
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rting
30
Sta
20
Op er at in g
DT
Therefore the higher pressure ratio is required only during the starting when a shock at the test section Mach number is necessarily to be catered to and thereafter the pressure ratio can be that corresponding to shock at the diffuser throat.
40
Pressure Ratio, rp
It has to be remembered that as the diffuser throat is larger than the nozzle throat, the Mach number there will be more than one but less than that in the test section. After the brief duration of starting, the pressure ratio may be decreased and the shock may be brought to the diffuser throat. rpo=(po1/po2)shock at M
10
0 1
2
3
4
5
6
Throat Mach Number, MT Pressure ratios for starting and operating the wind tunnel of different Mach numbers
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High Speed Wind Tunnels Closed circuit supersonic wind tunnel In summary, the pressure ratio for starting the wind tunnel is corresponding to the normal shock losses at the test section Mach number and that for operation is that corresponding to normal shock losses at the diffuser throat. Hence the power required can be considerably reduced by incorporating the well designed diffuser and by judicious control of the two pressure ratios during ‘starting ‘and ‘operating’ of the wind tunnel.
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Wind Tunnel Techniques
July-Nov. 2016
High Speed Wind Tunnels Actual flow in the supersonic wind tunnel The boundary layer thickness and the total loss of momentum increase with increasing distance from the 1st throat. The growth of boundary layer thickness with distance from first throat is predictable and can be accounted for in the nozzle design. In the steady state operation, viscous effects between the throat and test section are not of much importance. During the transient process in which tunnel is started, viscous effects are much important. So, important are these effects that pressure ratios required to start high Mach number tunnels are atleast 100% greater than the normal shock pressure ratio, po1/po2. i.e. viscous losses are almost equal to normal shock losses. Boundary layer is stable when the pressure is decreasing in the direction of its growth. When the pressure is increasing in the direction of flow, it has a tendency to separate.
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July-Nov. 2016
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High Speed Wind Tunnels Actual flow in the supersonic wind tunnel As normal shock passes through the nozzle, it imposes a severe unfavorable pressure gradient which can cause separation. If boundary layer separates, it disturbs the flow over a large portion of nozzle. If boundary layer does not separate the high pressure gain in the downstream of shock will tend to flow to low pressure boundary layer and the flow in the duct will be altered over a significant length of the nozzle. In the diffuser viscous effects are predominant during starting and steady state operation of the wind tunnel. Unfavorable pressure gradient exists always. An oblique shock from the convergence creates additional pressure gradients when they strike the opposite wall.
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Wind Tunnel Techniques
July-Nov. 2016
High Speed Wind Tunnels Actual flow in the supersonic wind tunnel As normal shock passes through the nozzle, it imposes a severe unfavorable pressure gradient which can cause separation. If boundary layer separates, it disturbs the flow over a large portion of nozzle. If boundary layer. does not separate the high pressure gain in the downstream of shock will tend to flow to low pressure boundary layer. and the flow in the duct will be altered over a significant length of the nozzle. In the diffuser viscous effects are predominant during starting and steady state operation of the wind tunnel. Unfavorable pressure gradient exists always. An oblique shock from the convergence creates additional pressure gradients when they strike the opposite wall.
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July-Nov. 2016
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High Speed Wind Tunnels Actual flow in the supersonic wind tunnel Starting a tunnel with a model in the test section It is explained earlier that
which implies that losses in total head resulting from shocks necessitate a larger diffuser throat. Hence, losses due to shocks on the model must also be provided for. So, for starting a tunnel with a model, a second throat larger than that for a clear tunnel is needed.
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Wind Tunnel Techniques
July-Nov. 2016
High Speed Wind Tunnels Actual flow in the supersonic wind tunnel Sizing of the wind tunnel model The theoretical unobstructed cross section area of the test section at the model required for starting is the same as the second throat area. In choosing the dimensions of the model, reflection of shocks should be also considered. The oblique shocks formed as shown in figure at the leading edge of the model get reflected from the wind tunnel wall. It has to be remembered that shock reflection is not specular which means that the angle of incidence of the shock at the test section wall is not same as that for reflection. The chord length of the model is so chosen that the reflected shocks do not interfere with the model.
Typical shock wave pattern from a model
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High Speed Wind Tunnels Actual flow in the supersonic wind tunnel Problems in the operation of supersonic wind tunnels Condensation: The amount of moisture that can be held by a unit volume of air increases with increasing temperature. When the air isentropically expands to higher Mach numbers in the test section, the temperature falls. It may become super cooled. Moisture will then condense.
Factors affecting condensation: a) Amount of moisture in the stream b) Static temperature of the stream c) Static pressure of the stream d) Time during which the stream is at low temperature
Effects of condensation: Condensation results in changes of local Mach number and other flow properties due to latent heat addition. The extent of changes depends on how much heat is released through condensation and may be evaluated using the two equations given below:
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Wind Tunnel Techniques
July-Nov. 2016
High Speed Wind Tunnels Actual flow in the supersonic wind tunnel Problems in the operation of supersonic wind tunnels
where dQ = heat added through condensation H = enthalpy per unit mass A = duct area When M > 1, Mach number decreases and pressure increases. When M < 1, Mach number increases and pressure decreases. Drying the working fluid is the best way to avoid condensation. Increasing the temperature by providing stagnation heaters is another solution
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High Speed Wind Tunnels Actual flow in the supersonic wind tunnel Problems in the operation of supersonic wind tunnels condensation, the components of air liquefy when proper temperature and pressure conditions are met. Liquefaction troubles might start around M=4 if high pressure air is expanded from room temperature. The figure shows the stagnation temperature required to avoid liquefaction at different Mach numbers. It can be seen that corresponding to a Mach number above 12 the temperature required will be about 2000K
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2000
Stagnation Temperature to avoid Liquefication (K)
Liquefaction: In a manner parallel to
1500
1000
500
0
4
6
8
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Mach Number, M
12
14
Stagnation temperature to avoid liquefaction at different Mach numbers
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Calibration of Wind Tunnels Introduction After a tunnel is constructed, the next step is to determine its flow characteristics and, to change any that are not satisfactory for the purposes intended. The low low-speed speed "steady" steady airstream is usually considered to be defined when we know its distribution of temperature T, pressure p, dynamic pressure q, and "turbulence" Tu. There are several assumptions embedded in this statement. We admit that the airstream is expected to be turbulent and therefore not strictly steady or time invariant. The time variability of the airstream is to be characterized by its level of turbulence, which is commonly defined to be the rms of the variation of the longitudinal component from the mean value of air speed. The pressure and dynamic pressure are the mean values for these quantities averaged over a time interval chosen as necessary to achieve the required precision of the mean. The stream temperature is similarly considered. Using the equation of state, we can then compute the density, and using the definitions of total pressure and total temperature, we can compute these quantities and flow speed as well.
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Wind Tunnel Techniques
July-Nov. 2016
Calibration of Wind Tunnels Introduction In the case of wind tunnels to be used for wind engineering, the temporal and spatial variations must be considered in considerably more detail and will include spectra and integral scales of the turbulent flow structures as well as a profile to match appropriate types of planetary boundary layers layers. We may also compute the Reynolds number for a particular model based on its chosen characteristic length. Much of our interest is centered on determining pressure, which can be measured most simply by use of liquid manometers or more commonly by using electronic data systems, including various instrumentation elements and computer systems for data manipulation, storage, and presentation. For our subsonic testing, the precision with which measurement of stream temperature are made are usually less demanding than the typical pressure measurements.
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Calibration of Wind Tunnels Test Section Speed Speed Setting Setting the speed of a wind tunnel appears straightforward as we apply our basic understanding of fluid dynamics. However, it turns our that this basic question absorbs a considerable amount of attention even for experienced aerodynamicists. aerodynamicists When there is no model in the rest section section, a measuring device, most commonly a Pitot-static tube, can be put there to determine the air speed. One cannot, however, insert a Pitot-static rube or other measuring device in the test section to measure dynamic pressure or speed along with an object under test because the test object will cause changes in the flow. These changes are referred to as "induced flow." Consider Fig. 1, which gives a schematic indication of the settling chamber, contraction and test section of a typical wind tunnel. The tunnel speed is usually determined by measuring either static or total pressure in the settling chamber ahead of the contraction cone, cone as indicated by station L, and a static pressure ahead of the test section, as indicated by station S. If honeycomb or screens are used in the settling chamber, station L will be downstream of these devices. Using the subscripts L for the bellmouth or settling chamber and S for the region before the test section, Bernoulli's equation between the two stations will be
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July-Nov. 2016
Calibration of Wind Tunnels Test Section Speed Speed Setting
Fig. 1 Typical measurement stations for a wind tunnel “q” system
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Calibration of Wind Tunnels Test Section Speed Speed Setting
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July-Nov. 2016
Calibration of Wind Tunnels Test Section Speed Speed Setting A better approach would be to survey a volume of the test section that is occupied by the model. If the Pitot-static tube used has a hemispherical nose, it can be equipped with two yaw heads at 90 90° to each other, and the distribution of the upflow and cross-flow cross flow can be obtained simultaneously. The survey of the lest section can be done with a simple Pitot-static tube assuming there is a method to position it throughout the test section or a rake of Pitot statics can be used. lf there is no method of remotely positioning the Pitot-static tube, the survey of the test section becomes a very tedious operation. In any event, this is a critical operation and the effects of the manner of holding and supporting any selected instrument must be carefully evaluated. If a volume of the test section is surveyed and there are large variations in the dynamic pressure it may be desirable to use a weighted average of the measured dynamic pressure, pressure. When the dynamic pressure calibration is completed, there is a relation established between the indicated dynamic pressure (PL - PS) and the dynamic pressure q, in the clean test section. There usually are a series of these calibrations for different testsection configurations. If the tunnel has an external balance, calibrations are often made with and without the balance struts and their fairings. If a ground plane is used often, calibrations are made for various heights of the ground plane.
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July-Nov. 2016
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Calibration of Wind Tunnels Test Section Speed Speed Setting It is desirable to obtain for each test-section configuration the distribution of the total pressure, static pressure, upflow, and cross-flow throughout the region occupied by the model. model When the tunnel has an air exchanger or heat exchanger exchanger, the temperature distribution should also be measured. The pressure sources for the tunnel speed control should be, ideally, either a ring around a tunnel station or at least a portion of a ring on the two side walls or the ceiling and floor. The possibility of the model pressure field directly affecting the nearer of the pressure sources at station S should always be kept in mind and evaluated if a larger model than normal or an upstream mounting location is to be considered. The static source should consist of either a series of static sources or a sealed tube flush with the surface with many holes evenly spaced along its length. The tubes or multiple static sources of the ring should be manifolded together to yield an average static pressure at the station. If total pressure is used in the contraction cone. It is desirable to have multiple sources also.
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Wind Tunnel Techniques
July-Nov. 2016
Calibration of Wind Tunnels Test Section Speed Speed Setting The two pressure sources in the simplest case can be connected to a U-tube manometer that is used for setting the tunnel speed. It is more common to use a highquality q y differential p pressure transducer. Then the electrical output p can be transmitted to any location convenient for the tunnel operator and to the data system. Extreme care must be taken to make sure that there are no leak in the tunnel dynamic pressure system. In large tunnels, the leak checking can be a time-consuming process. It is highly desirable and strongly suggested that the system should not be disturbed once it has been leak checked. It is often useful to record the pressures at each of the two tunnel stations as separate readings relative to atmospheric ambient pressure. Additional sources for this use should be provided rather than tapping into the tunnel dynamic pressure system. It is also desirable to have a simple method of periodically checking the system for leaks. One way to do this is to use a calibration wing that can be installed easily. The wing can be run through a pitch series at several dynamic pressures. If the slope of the lift curve does not change, there is no change in the dynamic pressure calibration. This has the further advantage of also checking the tunnel upfIow. If the drag polar does not rotate, the upflow has not changed.
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July-Nov. 2016
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Calibration of Wind Tunnels Horizontal Buoyancy As the air proceeds along the test section, the boundary layer thickens. This action reduces the effective area of the jet from that of the geometric dimensions and causes an increase in the flow speed outside the wall boundary layers. If the geometric area is constant, the speed increase in produces a drop p in local static p pressure, tending g to draw the model downstream. This added turn p drag is commonly called "horizontal buoyancy" as its action is analogous to the buoyancy due to the vertical pressure gradient in the atmosphere and the ocean. If the cross-sectional area of the jet is increased enough to allow for the thickening boundary layer, a constant value of the static pressure may be maintained throughout the test section. No exact design method is available that ensures the development of a constant static pressure. For a first approximation the walls of a closed jet should diverge about to each; finer adjustments may be necessary after the tunnel is built and the longitudinal static pressure is measured. Some tunnels whose test sections have corner fillets have these fillets altered until a constant static pressure is obtained. The advantages d t off such h a flflow are enough h tto jjustify tif a moderate d t amountt off work k iin obtaining bt i i itit. The amount of "horizontal buoyancy" is usually insignificant for wings, but for fuselages and nacelles it is larger and becomes important. For large torpedo like bodies, it can be of the order of the minimum drag in some tunnels.
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Wind Tunnel Techniques
July-Nov. 2016
Calibration of Wind Tunnels Longitudinal Pressure Gradient The static pressure gradient along the test section must be known in order to make necessary buoyancy corrections. It may be obtained by reading the local static pressure with a Pitot-static tube that is progressively moved from entrance cone to exit cone. Care must be taken that the Pitot tube is aligned directly into the wind and that no extraneous static pressure is created by th bracket the b k t holding h ldi th the Pit Pitott ttube. b Thi This llastt proviso i iis nott possible ibl tto satisfy ti f exactly. tl Perhaps a more convenient method is to use a long static tube, as discussed in the next paragraph. It is not desirable to measure the static pressure along the walls of the tunnel. It turns out that small variations in wall geometry associated with joints, mounting holes, and the like cause local variations in the static pressure that are typically too large to ignore and may change over time as various mounting installations are put in and taken out of a facility. Measuring wall pressures is a good adjunct to calibration and has become a common practice to obtain input for boundary corrections, but it is not a good idea to rely on wall pressure measurements for the baseline characterization. characterization Long Static Tube: This is a long tube that extends through the test section. The tube is often suspended on a series of wires that are used to align and tension the tube. The tube is equipped with a number of static pressure rings. These rings have four or more static ports around the circumference that are manifolded together. If the flow angularity is small, then aft of the nose the flow will be practically parallel to the tube. This then yields a static pressure distribution along the length of the tube.
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July-Nov. 2016
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Calibration of Wind Tunnels Dynamic Pressure Variation The dynamic pressure is usually measured throughout selected regions of the test section by means of a Pitot-static tube. The density is calculated from a barometric pressure measurement. A test-section static pressure measurement relative to the atmosphere, and a test-section temperature determination along with the equation of state.
The local velocities are then obtained from V=(2q/)0.5(7). The velocities as calculated from the dynamic pressures or the pressures themselves are plotted, and the points are connected by "contour" lines of equal values. The variation of q in the working range of the jet should be less than 0.50% from the mean, which corresponds to a 0.25% variation in velocity. Some tunnels have been built to tighter specifications. But it is almost inevitable that custom tweaking of aspects of the tunnel circuit will be required to do better. A plot of the dynamic pressure distribution in a rectangular test section is shown in Fig. 2. Of interest is the asymmetry that is usually found, and the maximum variation in this example is well above satisfactory limits. The survey should have been carried to the walls. Fig. 2 Distribution of dynamic pressure in the test section
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Wind Tunnel Techniques
July-Nov. 2016
Calibration of Wind Tunnels Dynamic Pressure Variation The correction of an excessive velocity variation is not as serious a problem as the correction of excessive angular variation. There are more method of attack as well as less probability that the variation will change with tunnel speed. There are several minor adjustments that may be expected to improve a less than satisfactory speed distribution. distribution There may be local flow separations that must be found and corrected or improperly set turning vanes. If the velocity variation is annular, the source may be the propeller load distribution. Such local problems in identifiable parts of the tunnel should be corrected. Finally screens may be added in the largest section of the tunnel with spatial mesh densities de s t es varied a ed so tthat at tthey ey a are e more o e de dense se in tthe e sections that correspond to high-velocity regions in the jet. The improvement in velocity distribution by such screens is shown in Fig. 3. The loss in energy ratio they cause is quite small and is far outweighed Fig. 3 Effect of screens on velocity by the improvement in testing conditions. distribution in the test section
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July-Nov. 2016
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Calibration of Wind Tunnels Flow Angularities The variation of flow angle in the jet can be measured by many devices. Regardless of the device used, it is desirable to map the upflow and cross-flow in a series of transverse planes along the longitudinal axis in the region occupied by the model over the range of intended dynamic pressures. Often, when the upflow and crossflow are plotted as flow direction vectors, regions of vortex like flow can be seen in the test section. Such flow is often the result of poor velocity distribution in the return duct before the third corner or the result of improperly set fourth-corner turning vanes. A variation of upflow across the span of a wing results in an effective aerodynamic twist. A cross-flow gradient across the test section in the region of the vertical tail will change the slope of the yawing moment versus side slip or yaw angle angle. Thus it is desirable to have the upflow and cross-flow constant across the tunnel. This is difficult to achieve. It would be desirable to have the variation less than + 0.10O, but it is often necessary to accept the best values that can be achieved. The maximum variation should be held to + 0.10O.
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Wind Tunnel Techniques
July-Nov. 2016
Calibration of Wind Tunnels Turbulence Measurements Variations between results of tests made in different wind runnels at the same Reynolds number and between tests made in wind tunnels and in flight have indicated that some correction was needed for the effect of turbulence that exists in wind tunnels. It has b been argued d th thatt this thi tturbulence b l causes flflow patterns tt iin th the tunnel t l to t be b similar i il tto th the flow pattern in free air at a higher Reynolds number. Hence the tunnel test Reynold number could be said to be a higher "effective Reynolds number.“ The physics of turbulent flow is far too complex to be captured by this simple concept. There are some phenomena for which it "works" to some extent and others for which it does not. This concept is dependent on the long-known fact that spheres (and circular cylinders) li d ) h have quite it well-defined ll d fi d critical iti l Reynolds R ld numbers b near which hi h th the d drag coefficient drops rather precipitously as the Reynolds number increases.
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July-Nov. 2016
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Calibration of Wind Tunnels Turbulence Measurements It has been experimentally verified that the Reynolds number at which the drag coefficient of a sphere decreases rapidly depends strongly on the degree of turbulence in the wind tunnel. The Reynolds number at which the reduction occurs decreases with increasing tunnel turbulence turbulence. There is also a strong effect of surface roughness. roughness We are considering only aerodynamically smooth spheres. The decrease in drag coefficient with Reynolds number can be understood as the result of increasing the Reynolds number, producing earlier boundary layer transition from the laminar to turbulent state, which in turns leads to a downstream shift in the separation point and corresponding higher base pressure, a smaller wake, and less drag. Early night measurements on spheres found that in the free atmosphere the critical Reynolds number for a sphere is 3.85X105. This value is larger g than is achieved in wind runnels, although g many y lowturbulence tunnels approach this value. In the atmosphere the turbulent eddies are so large relative to the sphere that they do not affect the thin boundary layer of the sphere.
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July-Nov. 2016
Calibration of Wind Tunnels Associated Instrumentation
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Fig. 4 Turbulence sphere Drag Coeffic cient of Sphere
Before the now common use of hot-wire anemometry, a turbulence sphere was the primary way to measure the relative t b l turbulence off a wind i d ttunnel.l It remains i a very useful and easy way to characterize the turbulent environment in a tunnel and to check if there is an indicated effect following changes in the tunnel configuration or special installations that might affect the flow quality. The configuration of a turbulence sphere is sho n in Fig shown Fig. 4 4. The critical Re Reynolds nolds number for the sphere can be measured in two ways. One method is to plot the measured CD based on cross-sectional area versus Reynolds number, as shown in Fig. 5.
0.5
0.4
0.3
0.2 Critical Reynolds Number=336,000 0.1 5 2x10
5
3x10
Reynolds Number of Sphere
5
4x10
Fig. 5 Variation of drag coefficient for a turbulence sphere as a function of Reynolds number
Wind Tunnel Techniques
July-Nov. 2016
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Calibration of Wind Tunnels Associated Instrumentation Before the now common use of hot-wire anemometry, a turbulence sphere was the primary way to measure the relative turbulence of a wind tunnel. It remains a very useful and easy way to characterize the turbulent environment in a tunnel and to check if th there is i an indicated i di t d effect ff t following f ll i changes h iin th the ttunnell configuration fi ti or special i l installations that might affect the flow quality. The configuration of a turbulence sphere is shown in Fig. 4. The critical Reynolds number for the sphere can be measured in two ways. One method is to plot the measured CD based on cross-sectional area versus Reynolds number, as shown in Fig. 5. From the plot the Reynolds number in the tunnel for CD = 0.3 is read. The second method is to take the average of the four pressures on the aft surface of the sphere and subtract this value from the stagnation value at the leading edge of the sphere, sphere yielding ielding P. P A plot of P/q versus ers s Re Reynolds nolds n number mber is prepared for the sphere and the Reynolds number determined for P/q = 1.22, as indicated in Fig. 6. The pressure method has certain advantages. It needs no drag balance with the associated balance calibration and no evaluation of support tares for the portion of the support sting exposed to the airstream. Also, the sphere support sting can be stiffer as no deflection is needed by the drag balance.
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July-Nov. 2016
Calibration of Wind Tunnels Associated Instrumentation 1.4
Prressure Loss Coefficient, P/q
From the plot the Reynolds number in the tunnel for CD = 0.3 is read. The second method is to take the average of the four pressures on the aft surface of the sphere and subtract this value from the stagnation value at the leading edge of the sphere, yielding P.
P/q=1.22
12 1.2
A plot of P/q versus Reynolds number is prepared for the sphere and the Reynolds 1.0 number determined for P/q=1.22, as indicated Critical Reynolds in Fig. 6. The pressure method has certain Number=299,000 advantages. It needs no drag balance with the 0.8 5 5 5 5 5 1 5x10 2 0x10 2.0x10 2 5x10 2.5x10 3 0x10 3.0x10 3 5x10 3.5x10 associated i t d balance b l calibration lib ti and d no 1.5x10 Reynolds Number, Re evaluation of support tares for the portion of the support sting exposed to the airstream. Also, Fig. 6 Variation of pressure coefficient for a turbulence sphere as a the sphere support sting can be stiffer as no function of Reynolds number deflection is needed by the drag balance.
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Calibration of Wind Tunnels Associated Instrumentation The critical Reynolds number as defined by either force or pressure measurements is then used to define a turbulence factor for the tunnel by comparing the tunnel's critical Reynolds number to the atmospheric free-air Reynolds number: TF 385 000/R tunnel TF=385,000/Re Then the effective test Reynolds number is defined by Reeffective=TFXRNtest The use of a turbulence sphere yields what may be thought of as an average value of tunnel turbulence. It does not give any information on the magnitude of turbulence in either the axial or lateral direction. The use of a turbulence sphere may, however, prove to be a simple method of monitoring any change in tunnel turbulence. Its use requires q no p prior calibrations and the installation and running g in a runnel can be designed to be simple and quick. The relation between the critical Reynolds number of a sphere and turbulence intensity as measured by a hot wire is shown in Fig. 7.
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July-Nov. 2016
Calibration of Wind Tunnels Associated Instrumentation
Turb bulence Intensity, TI (%)
Turbulence spheres can be made from cue, duck, and bowling balls. Several sizes are needed to enable the turbulence factor to be measured over a range of tunnel air speeds. 2.4 A brief examination of Fig. 7 might lead to the conclusion th t the that th higher hi h th the tturbulence, b l th the b better tt th the ttunnel,l as th the 2.0 effective Reynolds number of the test would be higher, This correction is not exact and if the tunnel has excessive 1.6 turbulence, the model may have a premature transition from laminar to turbulent flow, which can be critical for 1.2 laminar flow airfoils. However, low-speed models are often equipped with trip strips that fix the transition point on the 0.8 model and may reduce the requirement for extremely low 0.4 turbulence. The need for low test-section turbulence is not as severe for small student tunnels as it would be for 0.0 1.0 1.4 1.8 2.2 2.6 3.0 larger tunnels that are used for research and development Turbulence Factor, TF tests. The screens and honeycomb elements are effective Fig. 7 Variation of turbulence factor with for reducing turbulence in wind runnels. turbulence intensity from hot-wire probe measurements
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Mach Number
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Mach Number
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Mach Number Whereas Mach numbers in nearsonic and transonic tunnels are usually determined only on the tunnel centerline and at the tunnel walls, Mach numbers in supersonic tunnels are usually obtained off the tunnel center lines as well. The reason is that much larger non non-uniformities uniformities of flow are possible in supersonic nozzles because they can be caused by shock waves. An average Mach number in the vicinity of a model is desired for testing and data reduction purposes and the cross-sectional area survey will give a better average Mach number than the centerline survey. If the centerline Mach number distribution is constant or varies gradually but continuously, the centerline survey is usually adequate. However, the characteristics of this distribution are not known when the calibration is started.
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July-Nov. 2016
Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Total Pressure Pitot pressures are measured by using a simple device called a Pitot probe (tube). The Pitot probe is simply a tube with a blunt end facing into the airstream. The tube normally has an inside to outside diameter ratio of 0.5 to 0.8 (larger ratio is preferred with smaller outside diameter) and a length aligned with the airstream of 15 to 20 diameters. The pressure orifice is formed by the inside diameter of the tube at the blunt end. A rake of nine Pitot probes used for calibration traverses of a test section is shown in Fig. 8. A Pitot probe is both simple to construct and accurate to use. It should always have a squared-off entry and the largest practical ratio of hole diameter to outside diameter. It is insensitive to angle of attack upto 10 deg. for an orifice diameter only 10% of the outside diameter and 15 deg. for an orifice diameter only 98% of the outside diameter. Calibration errors due to angle of attack and hole size within the above ranges are much less than actual flow deviations found in any reasonable tunnel.
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Total Pressure
Fig. 8 A traversing rake of nine Pitot tubes
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July-Nov. 2016
Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Total Pressure At this point we may make the clarification that an open-ended tube facing into the airstream always measures the stagnation pressure (a term identical in meaning to “total pressure”) is sees. Above M=1, the shock wave that forms ahead of the tube means that it sees not the freestream stagnation pressure but the stagnation pressure behind a normal shock. This new value is called Pitot pressure and in modern terminology implies a supersonic stream, although there is no error in calling the pressure so measured in a subsonic stream “Pitot pressure”. Pressures measured by Pitot probes are influenced by very low Reynolds numbers based on the probe diameter. This effect is seldom a problem in supersonic tunnels as a reasonably sized probe will usually have will have a Reynolds number above 500 or 1000, which is the range g above which the trouble starts.
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July-Nov. 2016
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Static Pressure Static pressures in a supersonic flow are much more difficult to measure than static pressures in a subsonic flow and Pitot pressures in a supersonic flow. The “static pipe” used for calibrating nearsonic and transonic tunnels is rarely used as its presence in the tunnel affects the flow in the test section. section It changes the area ratio of the nozzle by subtracting from the effective throat and test section area, and it also interferes with the expansion pattern required for the development of uniform flow. While static pressure probes are not used extensively for calibrating supersonic tunnels, a great deal of effort has been devoted to the development of accurate static pressures for other applications. The result has been the development of probes in wind tunnel calibration and use. The primary problem in the use of static pressure probes at supersonic speeds is that any probe will have a shock wave at its forward end which a rise in static pressure. If the probe consists of a cone tip followed by a cylinder, the air passing the shoulder will be expanded to a pressure below stream static pressure. Then as distance from the shoulder increased, the pressure on the probe will approach the stream true static pressure.
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July-Nov. 2016
Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Static Pressure Pressure measurements on a cone-cylinder probe with a 7-deg. included angle cone followed by a cylinder 30 diameters in length are presented in Fig. 9. These results show negligible errors in static pressure measurements for orifices located 10 diameters downstream of the shoulder shoulder. Another type of probe designed for the accurate measurement od static pressures over a large Mach number range is shown in Fig. 10. Errors in static pressures measured with this probe are presented in Fig. 11. They were obtained by reference to a static pressure calculated from a measured Pitot and total pressure. Flow angularity during these tests was of no consequence, since the tunnel employed had extremely even flow. However it was found impossible to get consistent results until the Pitot probe mentioned above and the static probe were mounted in the test section on a vertically moving support system so that either could be moved into centerline calibration position without a tunnel shutdown and with little time interval between measurements. Evidently the extreme accuracy being sought (of the order of 0.01% of q or 0.1% of p), minute tunnel changes due to controls or thermal expansion became significant.
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Static Pressure
Fig. 9 Effect of orifice distance from shoulder on measured static pressure
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July-Nov. 2016
Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Static Pressure Static pressures on the walls of supersonic tunnels are often used rough estimation of test section Mach numbers. It is noted, however, that wall pressures do not necessarily correspond to pressures on the centerline because of the possibility of compression or expansion waves between the wall and the centerline. When Mach number is to be determined from static pressure measurements, the total pressure of the stream is measured in the settling chamber simultaneously with the test section static pressure. Mach number is then calculated from the two pressures.
Fig. 10 Approximate dimensions of supersonic static pressure probe There are three interdigitated rings of holes at a, b and c
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels
Fig. 11 Error in static pressure measured pm as a fraction of the true static pressure
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July-Nov. 2016
Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Mach Number Results of calibrations to determine Mach number from tunnel wall static pressures and tunnel axis Pitot pressures are presented in Fig. 12. This figure illustrates the difference that may exist between the tunnel wall and centerline due to expansion waves between the two. two Results of a very thorough Mach number survey at one axial station of a tunnel are presented in Fig. 13 and an even more complete survey along the plane of the horizontal axis is shown in Fig. 14. The importance of calibrating over the range of Reynolds numbers (pressures if the temperature is constant) at which the tunnel will be operated is shown in Fig. 15. A change in Reynolds number from 0.06 to 0.4x106 per inch causes a Mach number deviation of 0.06 at a nominal Mach number of 5. As mentioned previously, this effect results from changes in boundary layer thickness and consequently effective nozzle area ratio. As a matter of interest, many supersonic tunnels run at constant dynamic pressure throughout their Mach number range. This is in contradiction with the test parameter, V2. Constant q helps with handling balance loads, and with data reduction.
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Mach Number
Fig. 12 Wall and tunnel axis calibration data from M=2 nozzle
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Mach Number
Fig. 13 Contour plot of M=3 nozzle Fig. 14 Mach number distribution the AEDC M8 B tunnel
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Calibration of Wind Tunnels Mach h number deviatiion, M
0.00
-0.08
Calibration of Supersonic Tunnels M=1.5 M=2.5 M=3.5 M=4.5 M=5.0 M=5.5
-0.16
M=6.0
-0.24
0.1
0.2
-6
0.3
0.4
Reynolds number per inchX10 Fig. 15 Variation of centerline Mach number with Reynolds number
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July-Nov. 2016
Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Flow Angularity The flow angularity in a supersonic tunnel is usually determined by using either cone or wedge yawmeters. The sensitivities of several yawmeters for use in supersonic tunnels are presented in Fig. 16. The sensitivities of these yawmeters are maximum when the wedge d or cone angles l maximum. i Th They work kb below l M Mach h numbers b ffor which hi h wave detachment occurs, and are so used. The cone yawmeter is used more exclusively than the wedge yawmeter because it is easier to fabricate and more robust. A photograph of a wedge yawmeter is shown in Fig. 17. The use of the yawmeter at supersonic speeds requires a calibration to determine the aerodynamic error. A typical summary of results from tests to determine flow angularity in a wind tunnel is presented in Fig. 18.
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Flow Angularity
Fig. 16 Sensitivity of several yaw meters at supersonic speeds, pressure ratio per degree
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Flow Angularity
Fig. 17 A wedge yawmeter
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Fig. 18 Maximum up and down flow in a supersonic tunnel for about 0.7 tunnel height
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Turbulence Level Measurements with a hot-wire anemometer Table 1: Turbulence in settling chamber and demonstrate that there are high-frequency test section of a supersonic tunnel fluctuations in the airstream of supersonic Settling Chamber Test Section tunnels that do not occur in free stream. 4.5 Mach number All 2.2 These fluctuations, broadly grouped under 1% Sound, Pt/Pt Less than 0.1% 0.2% the heading, “turbulence”, consists of small Entropy, Tt/Tt Less than 0.1% Less than 0.1% oscillations in velocity, stream temperature Vorticity, V/V 0.5 to 1% Less than 0.1% (entropy) and static pressure (sound). Values obtained from one tunnel are presented in Table 1. The fluctuations arise from a variety of causes, mostly from the pressure regulator valve, the drive system, the aftercooler, and the test section boundary layer. Velocity fluctuations emanating from upstream causes may be reduced at low and moderate Mach numbers by the addition of screens in the settling chamber. At high Mach numbers, upstream pressure and velocity effects are usually less, since the large nozzle contraction ratios damp them out. Temperature fluctuations are unaffected by the contraction ratio. The existence of such fluctuations is of less interest than their effect. Here the calibration procedure has been to determine the transition Reynolds number on smooth cones and compare this value with values obtained in other tunnels.
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July-Nov. 2016
Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Determination of Transition By common usage, transition cones have either 5- or 10-deg included angles and highly polished surfaces. Various methods have been used to determine the point of transition on the cone. These have been (1) optical methods in which Schlieren pictures or shadowgraphs h d h were made d and d iinspected t d tto d determine t i th the point i t att which hi h a sudden dd thickening of the boundary layer occurred; (2) traversing along the cone a constant
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Determination of Transition By common usage, transition cones have either 5- or 10-deg included angles and highly polished surfaces. Various methods have been used to determine the point of transition on the cone. These have been (1) optical methods in which Schlieren pictures or shadowgraphs were made and inspected to determine the point at which a sudden thickening of the boundary layer occurred; (2) traversing along the cone a constant distance away from the surface with a small Pitot probe that is within the turbulent boundary layer and noting the point at which the probe pressure changes from a steady to a fluctuating pressure; (3) making similar traverses with a hot-wire anemometer and (4) measuring temperatures of the surface by using thermocouples on the inner surface of a thin walled cone. The dimensions of a cone for use with the surface temperature technique are presented in Fig. 19. Surface temperatures of this cone were used to obtain a recovery factor, Rr=(Taw-T)/(Tt-T). where Taw=adiabatic wall temperature, OK Tt=stream total temperature, OK and T =stream static temperature, OK The resulting data, which indicates the method of locating the point of transition is shown in Fig. 20. A summary of transition Reynolds numbers in several wind tunnels is presented in Fig. 21.
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Determination of Transition
Fig 19 Dimensions of JPL transition cone Fig.
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Determination of Transition
Fig. 20 Typical determination of transition Reynolds number on a cone Freestream Reynolds number per foot, 4.31x105, Freestream Reynolds number per foot, 4.31x105
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Determination of Transition
Fig. 21 Transition Reynolds number on 5- and 10-degree cones as measured in several facilities
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Test Section Noise Test section “noise” is defined as pressure fluctuations. Noise may result from unsteady settling chamber pressure fluctuations due to upstream flow conditions. It may also be due to weak unsteady shocks originating in a turbulent boundary layer on the tunnel wall Such weak shocks to which noise is attributed are illustrated by the shadowgraph wall. of Fig. 22. Noise in the test section of a wind tunnel is very likely to influence the point of boundary layer transition on a model. Although it is possible that other effects on tests results may exist, there has not been evidence showing that they do. Test section noise can be detected by either hot-wire anemometer measurements or by high-response Pitot pressure measurements. The method used to determine if the noise is coming from the test section boundary layer is to make measurements in the tunnel settling chamber as well as in the test section. It is then possible to determine if fluctuations in the two places are related. It has been found that test section noise generally increases as tunnel operating pressure increases, and, that test section noise originating in the settling chamber generally decreases as tunnel Mach number increases.
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Wind Tunnel Techniques
July-Nov. 2016
Calibration of Wind Tunnels Calibration of Supersonic Tunnels: Test Section Noise
Fig. 22 Noise emanating from the turbulent boundary layer on a missile model M=3.5; Re=2x106/inch. Note the diminution of wavelet strength as the distance from the source is increased.
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Calibration of Wind Tunnels Calibration of Supersonic Tunnels: The Use of Calibration Results
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Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t U Tube T b The basic notion off pressure that is implemented in measurement technology gy is force p per unit area. Other concepts p such as those embodied in various gas laws and theory are important in wind tunnel work but are necessary in relating pressure to other ideas and results rather than directly in the measurement of pressure pressure. One of the oldest devices for measuring pressures, and one of the easiest to build, build manometer; a term normally applied to a device used to measure differential pressure. pressure Common applications pp are measurements of the difference between a reference pressure such as atmospheric and a process pressure such as a port on a wing and measurements of the Fi 1 U-tube Fig. U t b manometer t difference between two pressures in a process process, such pressures f from th the ttotal t l and d static t ti ports t off a Pit Pitot-static t t ti tube. t b A simple i l manometer ((shown in Fig. g 1)) can be made from f two pieces off straight g tubing g made p parallel and connected byy tubing g at the bottom or by bending tubing into a U shape. shape The tubing is filled with a liquid liquid, and the difference heights in the two tubes is measured usually by an attached scale. measured, scale
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t U Tube T b The tubing g is filled with a liquid, q , and the difference heights g in the two tubes is measured,, usually by an attached scale. scale The plane defined by the centerlines of the two tubes may be inclined at varying angles to the horizontal horizontal, which provides varying sensitivity. sensitivity Setting th ttubes the b vertical ti l is i the th mostt common case. Th The diff difference in i pressures is i related l t d to t the th height difference and the parameters describing the manometer by the hydrostatic pressure equilibrium relation:
p2-p p1=h h sin g (l-g)
where h is the difference in extents of liquid q columns in the two tubes and is the angle g between the horizontal and the plane of the parallel manometer tubes. tubes Th product The d t h h sin i is i the th vertical ti l difference diff in i the th heights h i ht off the th two t columns. l Using smaller values of provides an increased sensitivity. sensitivity Here g, g is the weight per unitit mass, also l referred f d to t as acceleration l ti off gravity. it The Th value l needed d d here h is i the th "local" value,, which varies with location on Earth and with altitude. Also l is the density of liquid in the manometer (equivalently specific gravity times Also, densityy of water)) and g, is the densityy of air in most wind tunnel applications. pp
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July Nov 2016 July-Nov.
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Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t U Tube T b The manometer Th t iis the th mostt fundamental f d t l instrument i t t readily dil available il bl for f pressure measurement. Manometers are used frequently q y for calibrating g and checking g other devices,, as it is difficult to obtain a more accurate or precise p result in the range g of differential pressures commonly of interest in subsonic aerodynamic testing. In the past past, wind runnels have used a wide array of manometers. manometers These have from the simple i l U-tube, U t b similar i il to t Fig. Fi 4.1, 4 1 using i a ruler l or strip t i off graph h to t measure flfluid id heights h i ht to large g banks of 50-200 tube manometers with constant level reservoirs to maintain the reference fluid height g to manometers that can be precisely p y read using g some of the devices. The U U-tube tube and precision manometers could typically provide one or two pressures and the large multitube manometers were used to determine pressure pressures, distributions over model surfaces. surfaces The simple, simple precision precision, and small multitube manometers usually were read and recorded by an individual individual. The large multitube manometers t were photographed, h t h d and d th the fil films were read d later l t by b various i techniques. t h i The test test-section section total and static pressures were put on two of the manometer tubes. Then the pressure coefficients were merely the h of the desired pressure divided by the h of the dynamic pressure. pressure To process data from large test programs, programs the photographs of the manometers were used with optical schemes arranged to scale the images so that the pressure coefficients could be ready directly using a microfilm machine. hi Thi could This ld be b considered id d a form f off optical ti l data d t processing. i
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t M Manometer t Two useful manometers that can be built in almost anyy laboratoryy are described. The first is a simple unit with a few tubes and adjustable slant angle (Fig. (Fig 2). 2) The second is a unit (Fig. (Fig 4.3) 4 3) capable of very precise measurements with h readings of 0.0001 0 0001 in. in b i possible being ibl if an appropriate i t micrometer i t or height h i ht gage is i used d in i the th construction. t ti To achieve repeatability of such fine measurement, it is necessary that the thermal environment be quite stable. stable For example, example if the room temperature is around 75 OF, F a th thermal l shield hi ld mustt b be provided id d between b t the th person observing b i the th meniscus i through th h the optics p and the main part p of the unit. Otherwise steady y drift will be observed due to the radiant heat from the person to the manometer unit and the corresponding expansion of the material material.
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Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t Multitube M ltit b Inclination I li ti M Manometer t
Fig 2 A multi tube variable inclination manometer Fig. AE 2751
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Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t Multitube M ltit b Inclination I li ti M Manometer t
Fig 3 A hi Fig. high gh p precision i i micro i manometer t AE 2751
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Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t Pressure P T Transducers d
The most commonly used pressure transducers are of the diaphragm type, type which simply means that the basic sensing mechanism is a thin sheet of material that deforms as the pressure across it changes. changes The methods of sensing the diaphragm deformation include strain gages directly attached to the diaphragm, diaphragm circuits to sense the change in capacitance due to the geometric change change, and circuits to sense the change in inductance due to the geometric change. change The strain gage units are most numerous as they can be made smaller and d with ith currentt solid-state lid t t electronic l t i methods th d can be b made d ffor a ffew d dollars. ll Th These units it are remarkable k bl for f their th i economy although lth h they th mustt b be calibrated lib t d frequently f tl against i t more stable standard units. There are many strain gage units with stainless steel diaphragms and even some quartz diaphragms. Transducers using the capacitive sensing method tend to be the higher priced ranges and to have larger dynamic range capability as well as being more rugged than the strain-gage-based units. units These are more likely to be found the main tunnel condition sensing applications and for secondary calibration calibration. A majority of the pressure transducers used in subsonic wind tunnels are of differential type type, with ith a selected l t d reference f pressure applied li d to t the th reference f side. id Th The absolute b l t pressure versions i are almost l t always l available. il bl They Th differ diff from f the th differential diff ti l units it by b having h i one side id off the h di diaphragm h permanently l sealed l d to a fixed fi d pressure that h iis close l to vacuum. This Thi is necessaryy in order to avoid serious temperature p sensitivityy and leads to the absolute units having g diaphragms p g that can withstand least 1 atm. It turns out that this means theyy are not sufficiently su c e y se sensitive s e for o app applications ca o s in low-speed o speed wind d tunnels. u es
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t Micro Mi manometer t The output from the pressure transducers can be read by analogue or digital meters. meters A digital high precision micro manometer is shown in Fig. Fig 4. 4
Fig 4 A digital high precision micro manometer (FCO560) Fig.
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July-Nov. July Nov 2016
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Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t Scanning S i g Box B For measurement of multiple p pressures, p essu es, from o tthe e stat static cp pressure essu e taps on the models, models a scanning (selection) box (Fig. (Fig 5) or a scanivalve along with a digital high precision i i micro i manometer t can be b used. Each scanning g box can be connected to 20 nos. of static pressure taps. taps Slave units with 20 nos can be added to increase the nos. number b off static t ti pressures to t be b measured.
Fig 5 A 20-channel scanning box Fig. AE 2751
Wind Tunnel Techniques
July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t Pressure P Scanner S An alternative to the scanning box and digital hi h precision high i i micro i manometer is a pressure p scanner (Fig. 6). A pressure scanner has 32 individual pressure transducers The range of transducers. each h off th these ttransducers d can be different. Additional scanners each of 32 channels can be added to increase the number of pressures to t be b measured. d The size of the scanner is sufficiently to be mounted inside the wind tunnel model. model
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Fig 6 A pressure scanner (Scanivalve ZOC23B/32Px) Fig.
Wind Tunnel Techniques
July Nov 2016 July-Nov.
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Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t Pressure P Sensitive S iti P Paint i t Almost all of the preceding pressure pressure-measuring measuring devices require that the pressure at a point off iinterest t t be b communicated i t d to t the th sensing i element l t from f II pressure "tap" "t " or "port" " t" through th h a small tube to the sensor. There are two major j p problems with this method. The most important one is that it makes the models quite expensive to build. build The second is that the time response of the pressure measurement is limited by the presence of the tubing as a transfer mechanism. There have been a number of important p experiments p using g surfacemounted transducers, but these are by far the most expensive models and test programs. In any case case, there can never be a large enough number of pressure taps to provide high accuracy for f forces f obtained bt i d by b integration i t ti off measured d surface f pressure. Pressure-sensitive paint offers the promise of very high spatial density of measurements with a moderate impact on model cost. cost There is, is at present, present a high capital cost for the cameras needed for high-quality g q y results. This new method is being g used extensivelyy in transonic testing but has only recently been demonstrated to be capable of useful results for test speeds as low as 100 knots. knots The PSP principle is schematically illustrated in Fig. Fig 7. 7
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements P Pressure M Measurements: t Pressure P Sensitive S iti P Paint i t
Fig. 7 Oxygen Fig Oxygen-quenched quenched photoluminescence process AE 2751 Wind Tunnel Techniques July Nov 2016 July-Nov.
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Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t P Probes b A Pitot tube is used to measure total pressure. p The shape p of the tube affects its sensitivity to flows inclined to the tube axis. axis Pitots with hemispherical noses begin to show errors in total pressure at very low angles of flow inclination. inclination Pitot tubes with a sharp h square nose begin b i to t show h errors near 8O flow fl inclination. i li ti This Thi can be b improved i d by chamfering the nose. A Kiel Ki l tube t b can provide id accurate t stagnation t ti pressure for f flow fl angles l beyond b d 30O. Other Oth tip p shapes p p provide various sensitivities,, as indicated in Fig. g 8. It is a relativelyy easyy task to measure a Pitot's Pitot s sensitivity to flow angle by use of a flow of known angularity. angularity The Pitot is pitched or yawed depending on which is more convenient to determine its sensitivityy to fflow angularity. g y
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t P Probes b
Fig. 8 Variation of measured stagnation pressure with yaw for selected probe geometries: Cpa=2[p(a)-p(0)]/V2
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Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t P Probes b As previously indicated, indicated a total-head total head tube with a hemispherical tip will read the total head accuratelyy as long g as the yyaw is less than 3O. A squared-off q Pitot tube will go g to higher angles without error, error but both square square- and round-tip round tip Pitot tubes suffer errors if th are used they d att ttoo llow R Reynolds ld numbers b or ttoo close l to t a wall. ll At very low Reynolds numbers, numbers the flow regime is referred to as "creeping creeping flow flow." The difference ff between the pressure and stagnation g pressure is not the dynamic y pressure, 2 q=00.55V , which gives cpstag 1 for high high-Reynolds-number Reynolds number flow. flow In this regime the t =1 stagnation press pressure re on a bl blunt-nosed nt nosed bod body is given gi en approximately appro imatel in terms of the Reynolds y number based on the bodyy diameter as
cpstag=((pstag-pp)/q=1+6/Re 1 6/R D Corrections for Pitot tubes under these conditions are shown in Figs. Figs 9 and 10. 10
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t P Probes b 1.3
Cppsttaggnn
1.2
Square tip Hemisphere tip
1.1
1.0 10 100 1000 Reynolds number based on probe tip diameter, diameter ReD
Fig 9 Performance of Pitot probes at low Reynolds numbers Fig. AE 2751
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Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t P Probes b
Fig 10 Velocity correction for a circular Pitot probe near a wall Fig.
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t Static St ti Probes P b The most common device for determining the total pressure or total head and the static pressure of a stream is the p p pilot-static lube, an instrument that yyields both the total head and the static pressure pressure. A "standard" standard Pitot-static Pitot static tube is shown in Fig. Fig 11. 11 The orifice at A senses total t t l head, h d p+0.5 +0 5V2, and d th the orifices ifi att B sense the th static t ti pressure. This Thi discussion is limited to subsonic flows. If the pressures p from the two orifices are connected across a manometer or pressure transducer transducer. the pressure differential will be approximately pp y 0.5V2, from which the velocityy mayy be calculated provided p the densityy is available. available The density can be calculated from the equation of state based on a temperat re meas temperature measurement rement and the static press pressure re meas measurement. rement The Pitot Pitot-static static tube is easy to construct, construct but it has some inherent errors. errors If due allowance ll i made is d for f these h errors, a determination d i i off the h dynamic d i pressure within i hi about 0.1% 0 1% is possible. possible
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Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t Static St ti Probes P b
Fig 11 Pitot-static Fig. Pitot static probe AE 2751
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Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t Static St ti Probes P b The pressure sensed at the "static" static holes differs from the stream static pressure due to t two effects ff t off th the geometry. t Th The fifirstt effect ff t iis generic i for f a semi-infinite i i fi it axisymmetric i ti bodyy with flow approaching pp g the "nose" of the bodyy along g the axis of symmetry. y y The nose of the probe has a region on the upstream surface where the flow stagnates and the pressure is above the stream pressure of the approaching flow. flow At the stagnation point, point the p pressure is the total pressure p of the stream. The flow accelerates from the stagnation point around the curved surface and the local surface pressure rapidly falls through and below the stream or static pressure in the approaching flow. flow A minimum pressure is i reached h d somewhere h on the th curved d surface, f and d the th pressure starts t t to t "recover" eco e to toward a d the t e stream st ea stat staticc pressure p essu e as one o e moves o es downstream do st ea along a o g the t e surface parallel to the stream direction. direction If the probe were infinitely long and aligned with th flflow, th the then the th pressure on the th ttube b surface f would ld asymptotically t ti ll approach h the th stream p pressure with distance from the nose. This effect means that pressure taps on the probe's probe s parallel surface a finite distance f from th the nose would ld produce d a measurementt th thatt iis llower than th the th llocall stream t pressure. The amount of the difference in indicated in Fig. 12.
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Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t Static St ti Probes P b
Fig 12 Effect of static orifice distance from tip or stem Fig. AE 2751
Wind Tunnel Techniques
July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t Static St ti Probes P b The second effect is associated with the presence of the "stem stem," a cylinder whose axis is perpendicular to the approaching stream. stream A high-pressure high pressure region exists ahead of the stem and that region includes the surface of the probe itself. stem, itself In fact, fact the Prandtl design h a stagnation has t ti point i t near the th intersection i t ti off the th stem t and d the th probe b body. b d This Thi effect ff t creates t pressures on partt off the th probe b surface f that th t are higher hi h than th the th static t ti pressure in the approaching pp g stream. The two effects may cancel each other if the static hole locations are properly chosen. chosen The "standard" standard Pitot Pitot-static static tube does not employ this principle because the static hole l location ti is i so critical iti l that th t smallll deviations d i ti in i construction t ti or damage d to t the th tip ti can produce d a relatively l ti l large l error in i the th static t ti reading. di Th The P Prandtl dtl d design i (Fig. (Fi 13) is i intended to take advantage g of this cancellation. If a new Pitot-static tube is to be built, built either it may be designed as per Fig. Fig 11 and its static pressure readings corrected as per Fig. Fig 12 or the Prandtl design may be used used. The Prandtl design (Fig. (Fig 13) should require no correction but should be checked for accuracy. E Existing i ti Pit Pitot-static t t ti tubes t b should h ld be b examined i d for f tip ti and d stem t errors so that th t their constants may y be found. If a long static tube is available available, the static pressure can be determined along a longitudinal line in the test section. section Then the Pitot Pitot-static static tube can be placed on this line and the static pressure orifices in a Pitot-static Pitot static tube can be calibrated calibrated.
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Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t Static St ti Probes P b If a long static tube is available, available the static pressure can be determined along a l longitudinal it di l line li in i the th test t t section. ti Then Th the th Pitot-static Pit t t ti ttube b can b be placed l d on this thi line li and the static p pressure orifices in a Pitot-static tube can be calibrated. The accuracy of a standard Pitot-static Pitot static tube when inclined to an airstream is shown in Fig. g 14. If a Pitot-static tube is p placed near a model, the model's static p pressure field will influence the pressure sensed by the static ports ports, and the reading will not be the free freestream velocity velocity. This is why tunnel dynamic pressure calibrations are made without a model d l in i the th test t t section. ti The Th same problem bl exists i t on an aircraft, i ft where h greatt care has h to be ta taken e in the t e location ocat o of o a static stat c pressure p essu e sou source ce so as to find d a location ocat o where e e the t e static pressure varies as little as possible with lift or aircraft attitude. attitude
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Wind Tunnel Techniques
July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements Fl Flow Instrumentation: I t t ti Pitot Pit t Static St ti Probes P b
Fig 13 Prandtl Fig. Prandtl’s s design for a Pitot Pitot-static static probe
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Fig 14 Performance of the standard PitotFig. Pitot static probe in yaw
Wind Tunnel Techniques
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements t The flflow velocity Th l it can be b measured d using i one off the th following f ll i techniques: t h i 1 Pressure probes 1. 2. Pressure probes p using g fast response p pressure p transducers 3 Thermal Anemometry 3. Anemometr 4 Optical Anemometry: Laser Velocimetry 4. Velocimetry, Laser Doppler Anemometry and Particle Image Vecocimetry The above techniques can measure the three velocity components in an incompressible flow For measurements in compressible flow, flow. flow additionally temperature has to be measured d using i a suitable it bl instrument. i t t Pressure probe measures time averaged velocity and its components. components Pressure probes using sing fast response press pressure re transducers, transd cers thermal anemometr anemometry and optical anemometry y measures time dependent p velocityy and its three components. p In addition the six Reynolds stresses can also be measured. measured F th For the sake k off b brevity, it lilimited it d discussion di i off the th above b techniques t h i is i presented. t d
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Pressure P Probes P b Yawheads: A yawhead is a sphere that has two or more static ports on the forward face of the sphere (Fig. (Fig 12). 12) If the ports are at exactly 45O to the centerline of the support and the flow is parallel to the support, support then (Pa-P Pb)/q=P/q=0. =0 If there th iis flflow angularity, l it then th P/q≠0, ≠0 and d the th value l off P/q will ill be b a function f ti off the th flow angle. g In practice, p a yyawhead must be calibrated. The probe p is placed p in a flow that has no angularity, g y, and the probe p is pitched p or yawed y about its center through g an angle g range g both p positive and negative. g This is done for the y yawhead in the "upright," p g , or normally used, position, and then the yawhead is rotated 180O to an "inverted" inverted position. This wilI result in two curves of P/q=0 0 (upright and inverted) versus the angle for the yawhead If the static ports are symmetrical to the support axis yawhead. axis, the two curves will lie on top of each other other. If there is an asymmetry in the static port locations or possibly in the conditions of the static ports themselves, themselves the curve will be displaced by twice the error The true curve lies halfway between the two measured curves. error. curves If the flow used to calibrate the yawhead is not parallel to the yawhead support axis at the zero angle, angle the tr e ccurve true r e will ill not pass through thro gh the zero ero angle angle. It is normally normall desired to have ha e the yaw a h d calibration head lib ti independent i d d t off dynamic d i pressure and d static t ti pressure so the th coefficient ffi i t d fi d b defined below l may be b used: d Cy=(p (pb-p pa))/[p [pc-0.5(p (pa+p pb)] It is instructive to investigate the variation of the coefficient as if the pressure distribution on the spherical yaw head were accurately given by potential flow theory for a sphere in a uniform stream stream.
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Pressure P Probes P b
Fig Fig. g 15 Yawhead AE 2751
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Pressure P Probes P b Quite Q ite often a yawhead a head has five fi e static ports rather than the three just j st described. described With such an arrangement, g , the two flow angles g required q to define the three-dimensional flow vector may be obtained although the calibration is now over the two angles and a surface must be generated rather than a single curve curve. Pitot Pitot-static static tubes that use a h i h i l nose can be hemispherical b made d to t have h two t yawheads h d built b ilt into i t the th nose. In I this thi case, one instrument at any place in the test section will measure total pressure, static pressure dynamic pressure, pressure, pressure upflow, upflow and cross-flow. cross-flow This is quite useful for determining fl flow conditions diti in i a test t t section. ti Claw p probe: A claw probe p will also measure flow angularity g y and is simple p to build. In its simplest form a claw probe consists of two parallel pieces of tubing that are bent 45° 45 away from their common axis and then bent back 90 90° toward their common axis (Fig. (Fig 16) The 16). Th two t heads h d off the th probe b are cutt off ff square about b t two t diameters di t from f the th centerline. Often a third tube is added to measure total pressure, and the two claws can be made to simultaneously measure both cross-flow cross flow and upflow upflow. The calibration t h i technique for f a claw l probe b is i the th same as that th t for f a yawhead. h d Cl Claw probes b are more delicate than yyawheads because the two tubes used to measure can be easily y bent,, thus changing the calibration calibration.
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Pressure P Probes P b
Fig 16 Cl Fig. Claw p probe b AE 2751
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Pressure P Probes P b Cone Probes: C P b C Cone probes b with i h five fi or seven boles b l are widely id l used d for f flow fl vector measurements. They y have the advantage g of simplicity p y of construction and can be calibrated more readily at higher angles. angles Gupta gives detailed treatment and examples of calibration results Gallington and Gerner et al results. al. give results that allow calibration for compressible flow. flow Probes for Measurements in Reverse Flow: A number of developments p have been carried ca ed out o on probes p obes to allow a o measurements easu e e ts to be made ade in flow o with t angles a g es from o almost a ost any direction. direction Cogotti describes and has used a 14 14-hole hole probe extensively extensively. This probe is essentially two 7-hole 7 hole cone probes in a back-to-back back to back configuration configuration. Yamaguchi g et al. and Rediniotis have developed p spherical p probes p with 13 holes arranged g as multiple 5 5-hole hole yaw probes. Gupta gave procedures for a spherical probe with 11 holes arranged in the pattern of the faces of a dodecahedron with the probe stem emerging from the location of the 12th face center. center Other Pressure Devices: Sometimes a simpler p version of a y yawhead is used. One device consists of five tubes arranged in a cross configuration with one tube in the center and two pairs of tubes attached to it at right angles. angles The center tube is cut off with its end approximately perpendicular to the flow. flow The other tubes are cut off at a 45O angle. angle A second dd device i is i similar i il but b t consists i t off two t parallel ll l tubes t b cutt off ff att an angle. l Both B th off these th devices are calibrated in a manner similar to a y yawhead or claw probe. p
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Devices D i R Responding p di g to t Aerodynamic A dy i Forces F The devices described here are small airfoils wedges airfoils, wedges, and vanes. vanes They provide a response based on aerodynamic forces over surface f areas llarger th than a typical t i l pressure port. Airfoils Ai f il and d Vanes: V Fl Flow angles l can be b measured using g small wings g attached to sting balances balances. The flow angle calibration Fi 17 Vane-type Fig. V t angularity l it probe. b U Uses a fivefi is obtained by the rotation of the drag component p balance and can resolve about polar. l S Smallll vanes off various i configurations fi i 0 006O. Owing to its natural high frequency 0.006 frequency, can also be used (Fig. 17). Often these probe can make continuous motion surveys are attached to low-friction low friction potentiometers t read to d th the angle l or to t balances. b l
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Thermal Th l Anemometry A ty Hot wires and hot films are used to obtain fast-response velocity measurements in turbulent flows as well as mean velocities and, and with multiprobes, multiprobes flow angularity. angularity The probe or sensor is a fine wire (diameter p ( of a few micrometers)) or a cooled quartz q fiber attached between two supporting needles on the probe (Fig. (Fig 18). 18) Current passed through the wire or film raises its temperature above the adiabatic recovery temperature of the gas. The Th hot h t wire i responds d to t changes h in i total t t l temperature t t and d mass flux fl (To and d U). U) IIn subsonic applications pp where the densityy is high g and the fluid temperature p is low and constant the problem of beat transfer through the support needle (end losses) and constant, radiation effects can be ignored, ignored and the wire's ire's response basically basicall is a function f nction of velocity elocit alone. Under these conditions and using g appropriate pp p calibration and measurement of the voltage across the bridge bridge, both mean and turbulent velocities are obtained. obtained Electronic El t i hot-wire h t i circuits i it include i l d a feedback f db k system t to t maintain i t i the th wire i att a constant t t wire resistance. This is a constant constant-temperature temperature anemometer. Feedback can also be used to maintain a constant-current anemometer anemometer. Constant-temperature anemometers are easiest i t tto use iin subsonic b i or iincompressible ibl flow, fl while hil the th constant t t currentt anemometers t are p preferred in compressible p flow. Frequency q y response p to 50 kHz is easilyy obtained.
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Thermal Th l Anemometry A ty Hot wires and hot films are used to obtain fast-response velocity measurements in turbulent flows as well as mean velocities and, and with multiprobes, multiprobes flow angularity. angularity The probe or sensor is a fine wire (diameter p ( of a few micrometers)) or a cooled quartz q fiber attached between two supporting needles on the probe (Fig. (Fig 18). 18) Current passed through the wire or film raises its temperature above the adiabatic recovery temperature of the gas. The Th hot h t wire i responds d to t changes h in i total t t l temperature t t and d mass flux fl (To and d U). U) IIn subsonic applications pp where the densityy is high g and the fluid temperature p is low and constant the problem of beat transfer through the support needle (end losses) and constant, radiation effects can be ignored, ignored and the wire's ire's response basically basicall is a function f nction of velocity elocit alone. Under these conditions and using g appropriate pp p calibration and measurement of the voltage across the bridge bridge, both mean and turbulent velocities are obtained. obtained Electronic El t i hot-wire h t i circuits i it include i l d a feedback f db k system t to t maintain i t i the th wire i att a constant t t wire resistance. This is a constant constant-temperature temperature anemometer. Feedback can also be used to maintain a constant-current anemometer anemometer. Constant-temperature anemometers are easiest i t tto use iin subsonic b i or iincompressible ibl flow, fl while hil the th constant t t currentt anemometers t are p preferred in compressible p flow. Frequency q y response p to 50 kHz is easilyy obtained.
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July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Thermal Th l Anemometry A ty
Fig 18 Three Fig. Three-element element hot-film hot film probe
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Optical Opti l Anemometry A ty A laser l velocimeter l i t uses a tracer t method, th d which hi h depends d d on detection d t ti off an optical ti l effect from a p particle that is carried with the flow. The most common laser Doppler pp velocimeter (LDV) uses optics to split the laser beam into two parallel beams that are focused to cross at the point where measurements are to be made (a dual-beam dual beam s stem) O system). Owing ing to wave a e interference interference, a fringe pattern in an ellipsoid ellipsoid-shaped shaped volume ol me at the beam intersection is formed. A second lens assemblyy (the ( receiver)) with a small aperture is focused on this fringe region to collect light from seed particles crossing the fringes fringes. This light is fed to a photodetector that is used as the input to the sophisticated electronic signal processor that measures modulated frequency frequency. In a dual-beam system y the frequency q y from the scattered light g from the two beams are superposed on the photodetector's photodetector s surface surface. The mixing process in the photodetector then gives the difference in the frequency from the two beams beams. All other th frequencies f i are too t hi high h to t detect d t t (thi (this iis called ll d an optical ti l heterodyne). h t d ) If is i the angle g between one of the beams and the bisector of the two beams and is the wavelength of the laser light light, then it can be shown that when the particle velocity is much less than the speed of light the modulator frequency is fD=2us sin
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Wind Tunnel Techniques
July Nov 2016 July-Nov.
Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Optical Opti l Anemometry A ty where us, is the velocity parallel to the plane of the two beams and perpendicular to the bi bisector t off th the b beams. Th Thus th the relationship l ti hi between b t the th flow fl velocity l it and d fD is i linear li and a function of half the beam angle g and the wavelength g of the laser light. g Perhaps p it is easier to think of the system working in the following manner: When a p particle moving g with the flow passes p through g each fringe, g it is illuminated, which causes a series of pulses from the photomultiplier. photomultiplier As the distance between the fringes is known, known the time to cross the fringes is measured, measured and this yields the particle velocity. velocity It should h ld be b noted t d that th t th the ffringe i spacing i iin micrometers i t is i equall to t the th calibration lib ti constant in meters per second per megahertz. When a large number of samples are taken the signal processor and computer calculates both the average velocity and taken, i t t instantaneous velocity, l it which hi h can be b used d to t obtain bt i the th turbulence t b l or velocity l it variation. i ti
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Total T t lH Head d Rake R k Airfoil Ai f il profile fil drag d has h often ft been b measured d by b the th use off a drag d wake k rake. k In I this thi t h i technique the th momentum t lloss iin the th wake k is i determined d t i d by b measurements t with ith a bank b k of total head tubes. The rake also should have two or more static tubes offset from the total head tubes to obtain the local static pressure. p The rake size must be adequate q to encompass p more than the width of the wake. Often,, the tube spacing p g is greater g at both ends of the rake than in the center by a factor of 2. The spacing of the tubes must be known with precision so that the momentum profile can be accurately determined. determined Generally the total tubes are made of 0.0625 0 0625-in in. thin-wall thin wall tubing (Fig (Fig. 19). 19) The static tubes must be offset from the plane of the total tubes to avoid interference effects on the static pressure pressure. Their purpose is to determine the static pressure in the wake. wake Because the static pressure can be affected by both the total head tubes and the base of the rake they must be carefully calibrated so that the error in static pressure is known. rake, known Since the static ttubes bes m must st ha have e a hemispherical nose shape, shape Krause Kra se has shown sho n that it i possible is ibl to t adjust dj t this thi shape h to t reduce d the th error to t zero. As an alternative to a wake rake, a mechanical traversing mechanism can be used. Using various encoders, the location of a probe can be determined with a high degree of accuracy accuracy. The sensor can be a Pitot and, and preferably, preferably a static tube, tube a multihole probe, probe a hot wire wire, or a thin film or even a fiber-optic laser velocimeter head head. When measuring the momentum loss in the wake by any method, method care must be taken to ensure that the whole width of the wake is measured. measured
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Total T t lH Head d Rake R k
Fig 19 T Fig. Total t lh head d rake. k N Note t single i gl static t ti p pressure tube. t b
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t B Boundary d yL Layer y M Mouse Measurements in the boundaryy layer y are often made to detect the transition between laminar and turbulent flow or to find the local skin skin-friction friction coefficient coefficient. Obtaining knowledge of the velocity profile in the boundary layer is important in attempts to determine these quantities. titi Withi Within an attached tt h d b boundary d llayer the th static t ti pressure is i essentially ti ll constant t t while the total pressure varies. There are several ways in which the velocity profile can be obtained. obtained The oldest method is byy use of a boundaryy layer y mouse (Fig. ( g 20). ) This device is a series of total head tubes tubes, often with oval or flat inlets. inlets To obtain the velocity profile with adequate resolution at the surface requires the total head tubes to be spaced closer together than their diameters. Thus, the total head tubes are placed on an inclined plane to obtain the required close vertical spacing. The boundary layer mouse often has a static orifice to measure the static pressure or the static pressure can be measured by a surface port. port D i use the During th mouse is i attached tt h d to t the th model. d l The Th boundary b d layer l mouse measures the th velocityy profile p over a finite span p of the model,, rather than a single g spanwise p station.
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t B Boundary d yL Layer y M Mouse
Fig 20 Boundary layer mouse Fig. AE 2751
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t T Traversing i g Probes P b The velocityy profile p can also be measured byy using g a traversing g mechanism whose position off the surface can be quite accurately determined by a digital optical encoder. encoder The traverse mechanism can carry a single total head probe, probe a hot wire or a split film. film Very good d agreementt h has been b shown h between b t Pitot Pit t probes, b hot h t wires, i and d thin thi films fil when h supported by a traverse mechanism. IIn general,l it is i better b tt to t supportt the th boundary b d layer l mouse or the th traverse t mechanism h i from the model rather than the tunnel walls. This avoids two problems. p First,, when the walls are used for support, support the probes must be moved when the model is pitched and then reset to obtain a very close and accurately known proximity to the surface. surface The second problem with a wall support is that most models tend to move slightly g y and often f oscillate when under loads owing to balance deflections. Or simply structural flexibility. If the probes are being used to detect transition between laminar and turbulent flow, flow extreme care mustt b be taken t k to t ensure that th t th the probe b it itselflf is i nott causing i transition t iti prematurely. t l
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Transition T iti Manyy methods are in use to determine the location of the transition region. g Theyy include the following: 1 Pl 1. Plott the th velocity l it gradient di t in i the th boundary b d layer l and d determine d t i whether h th the th flow fl is i laminar or turbulent by the slope of the gradient, as illustrated in Figs. 21 and 22. 2 Determine 2. D t i the th b beginning i i off transition t iti as the th point i t where h the th velocity l it as a function f ti off streamwise distance at a fixed small height g above the bodyy surface is a minimum,, as illustrated in Fig. Fig 23. 23 3. Read the static p pressure at a small height g above the surface, determine the transition by a slight dip in the plot of static pressure versus percent chord. chord 4 R 4. Read d the th velocity l it att a smallll height h i ht above b the th surface f with ith a hot-wire h t i anemometer t and d note the transition as a region of unsteadiness in the output. 5 R 5. Read d the th velocity l it att a smallll height h i ht above b the th surface f with ith a hot-wire h t i anemometer t or thin-film gage g g and note the start of transition as the point p of minimum velocity. y
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Transition T iti
Fig 21 Velocity distribution in laminar and turbulent boundary layers Fig. AE 2751
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Transition T iti
Fig 22 Boundary layer six chordwise stations along an airfoil Fig. AE 2751
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Transition T iti
Fig 23 Velocity in boundary layer at a constant small height above surface Fig.
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Transition T iti 6. Carefullyy emit smoke from flush orifices and note the transition byy the dispersal p of the smoke stream (may be difficult at high velocities) velocities). 7. P 7 Paint i t th the model d l with ith special i l chemicals h i l that th t evaporate t slowly. l l The Th evaporation ti will proceed most rapidly where the flow is turbulent. 8. Li 8 Listen t tto th the b boundary d llayer with ith an ordinary di doctor's d t ' stethoscope t th connected t d to a flat total head tube,, moving g the total head tube p progressively g y along g the surface from the beginning of the boundary layer in the downstream direction. direction As long g as the flow f is laminar, a soft f sh-sh-sh-sh can be heard. When it is turbulent, a distinct roar is heard. This same input fed into a transducer becomes quite graphic on an oscilloscope or amplified and fed to an audio speaker. speaker
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Wind Tunnel T nnel Measurements Meas rements V l ity M Velocity Measurements: t Preston P t Tubes T b This device is used to experimentally p y measure the wall coefficient of skin friction byy measuring both a static pressure and total pressure at the same chordwise location. location The total pressure is measured by a Pitot tube that touches the surface surface. This can be done b because ffor unseparated t d tturbulent b l t flflow th there iis a region i near the th wallll on the th order d off 10% of the boundary layer thickness in which the flow depends on the local wall skin friction w, the density , the kinematic viscosity and a length. length Preston took one-half of th Pit the Pitot-tube t t b di diameter t as the th length. l th Dimensional Di i l analysis l i leads l d to t th the equation: ti
Other methods using thermal anemometry and optical anemometry are available to measure skin friction and its temporal variation on the airfoil surfaces and other surfaces surfaces.
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements t The purpose p p of load measurements on the model is to make available the forces and moments so that they may be corrected for tunnel boundary and scale effects and utilized tili d in i predicting di ti the th performance f off the th ffull-scale ll l vehicle hi l or other th device. d i The loads may be obtained by at least the following four methods methods, which are listed in the order d off ffrequency off use: (1) measuring i the h actuall forces f and d moments on the h complete l model or on parts of the model with one or more balances; (2) measuring the stress distribution over the model by means of orifices connected to pressure measuring devices or other means such as pressurep or shear-sensitive coatings; g ; (3) ( ) measuring g the effect that the model has on the airstream by wake surveys and tunnel wall pressures; and d (4) measuring i the th motion ti off the th model d l under d the th action ti off the th aerodynamic d i forces f and computing the forces from equations of motion. In this course, we will consider onlyy the first of these, as force balances are universallyy used in all wind tunnels. tunnels
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t R Reference f Frames F The two most used reference frames are body y axis frame and wind axis. A third reference frame is referred to as stability axes. axes Any reference frame is determined by its orientation relative to some other frame or a basic physical reference and the location of th origin. the i i A reference f frame f iis a sett off th three orthogonal th l axes, by b convention ti always l labeled in a right right-hand hand sequence. Wind Wi d Axes: A F For wind i d tunnel t l applications, li ti we first fi t consider id the th wind i d axes. We W illustrate ill t t this in Fig. g 24. The wind axes have xw p pointing g into the wind,, zw p pointing g down,, and yw pointing to the right looking into the wind. wind If the test section is not horizontal, horizontal then an appropriate local convention must be adopted adopted. We show vectors indicating components off fforces that are used in wind axes. Note that drag g is in the negative g xw direction and lift f is in the negative zw, direction while the side force is in the positive yw direction. We note W t th thatt a perfectly f tl aligned li d wind i d tunnel t l would ld have h the th wind i d axes exactly tl parallel ll l to t the wind tunnel axis. In reality, y, there are angularities, g , and these lead to the wind axes for any given model that may be angularly offset from the tunnel axes. axes
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t R Reference f Frames F
Fig 24 Wind and body reference frames Fig. AE 2751
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t R Reference f Frames F Bodyy axes: The bodyy axe are fixed f to the model and move with it. The exact alignment g with anyy particular p model must be specified p as part p of test planning. p g The xb-zb p plane is frequently a plane of symmetry or approximately so. so The force components on body axes are sometimes referred to as axial force, force side force, force and normal force for the xb, yb, zb components respectively, components, respectively or sometimes as body drag, drag body lift and body side force. force This multiple lti l terminology t i l can lead l d to t confusion f i and d requires i attention tt ti to t avoid id errors in i communication. We will use the former set to refer to bodyy axis components. p Moments and Reference Frame Origins: The moment components on the xx, yy, z axes are referred to as rolling moment, moment pitching moment, moment and yawing moment, moment respectively respectively. If context t t iis nott sufficient ffi i t tto di distinguish ti i h between b t what h t is i intended, i t d d then th body b d roll, ll body b d pitch it h and body y yyaw or wind roll,, wind pitch p and wind yaw y must be used. Note that the origins g of the reference frames must be carefully specified in every case since the moments are directly and critically dependent on this choice and there is no universal standard. standard Model Attitude: The standard wayy to specify fy model altitude is to use an Euler angle g sequence q going g g from wind axes to bodyy axes of “yaw”, y , , about the zw axis,, "pitch“, p , , about an intermediate axis axis, and "roll" roll , about the xb axis. axis However, However it is the aerodynamic angles angle of attack, angles, attack , and sideslip, sideslip , that are the preferred independent variables for writing aerodynamic functions. functions
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t R Reference f Frames F
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Moment M t Transfers T f
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l We have W h been b treating t ti relations l ti between b t force f and d momentt components t on different diff t reference frames. A wind tunnel balance is expected to separate these force and moment components and accurately resolve what is almost always small differences in large g forces. A complicating p g factor is that the various force and moment components p vary widely in value at any given air speed and each varies greatly over the speed range ffrom minimum i i tto maximum. i B Balance l d design i and d use are problems bl th t should that h ld not be deprecated; in fact, it might truthfully be said that balance design is among the most trying problems in the field. field Concept of a Six Six-Component Component Balance: In an attempt to picture the situation most clearly a conceptual but impractical wire balance based on spring scales is shown in clearly, Fig. g 26. The model,, supposed pp to be too heavyy to be raised byy the aerodynamic y lift,, is held by six wires. wires Six forces are read by scales A, A B, B C. C D, D E, E and F. F The wires attached t A and to d B are parallel ll l to t the th incoming i i air i velocity l it vector t and d define d fi a plane l that th t can be b taken as a reference plane for the balance. We will designate this the xx-y y plane.
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component t Balances B l p These wires point p in the x direction. The wire attached to F is perpendicular to the A and B wires and is in the x-y x y plane plane. This i points i t in i the th -y di ti Th i wire direction. The wires attached to C and D are in a plane that is perpendicular to the xx-yy plane plane, which we d i designate t the th y-z plane. l Th The C and d D wires i are p perpendicular p to the x-yy p plane. Wires A and C are attached at a common point on the right wing wing. Wires B, B D, D and F are f attached to a common point on the left wing. Finally the wire attached to E is parallel to C and D and is in a plane parallel t C and to d 0 and dh halfway lf between b t them: th
Fig 26 Di Fig. Diagrammatic g ti wind i d tunnel t l balance b l AE 2751
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l 1 Since the hori 1. horizontal ontal wires ires A, A B, B and F cannot transmit bending, bending the vertical ertical f force perpendicular di l to t Y Y, th the lift, lift iis obtained bt i d from f the th sum off the th forces f i the in th vertical wires: L=C+D+E. 2 The drag is the sum of the forces in the two horizontal wires parallel to the 2. direction of V: D=A + B. B 3. The side force is simply p y Y=F. 4 If there is no rolling moment, 4. moment that is no moment component in the direction of the x axis axis, scales C and D will have equal readings. readings But more generally a rolling moment will ill appear as I=(C I (C - D) X b/2. b/2 Note carefully caref ll that this i with is ith reference f to t a point i t halfway h lf between b t the th two t wires i C and d D through th h which the line or action of F passes, p and in the plane p defined as containing g E. 5 Similarly a yawing moment, 5. moment that is a moment component in the direction of the z axis axis, will result in nonequal forces in the wires A and B and the yawing moment will be given by n=(A - B) X b/2. b/2 Here also note that thi iis a momentt with this ith reference f to t a point i th halfway lf between b t A and d B and d through g which the line off F passes. 6 The pitching moment is given by m=E X cc. This is a moment about the 6. line containing F. F
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l Exactt perpendicularity E di l it between b t the th wires i mustt be b maintained. i t i d For F instance, i t if the th wire to scale F is not exactly perpendicular to wires A and B, a component of the drag will appear (improperly, (improperly of course) at scale F and will be interpreted as side f force. A similar i il situation it ti exists i t in i regard d to t lift and d drag d and d lift and d side id force. f Si Since the lift is the largest g force by y far in typical yp aircraft complete p model wind tunnel work,, extreme care must be taken to ensure that it is orthogonal to the other components. components To illustrate the situation in more detail, detail consider a planar subset of lift, lift drag and pitching moment moment, as indicated in Fig. Fig 27. 27 We will assume that we can determine precisely i l the th direction di ti off th the x and d z axes and d apply l loads l d along l those th axes in i order to explore p the reaction in the wires. The expressions p for equilibrium q in the X and Z directions are Z+A sin -C C cos -E= E= 0
X+A cos -C C sin =0
m+Ec=0
(11 )
Assume Ass me the act actual al loads are m=0 0 and Z Z=10X. 10X What will ill o ourr balance read with ith the improper p p alignment g as indicated byy Fig. g 27. Solving g for A,, C,, and E gives g E=0
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Fig. Fig 27 Effect of balance component skew
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l Four types yp of external balances have been in general g use. These balances are named for their main load-carrying y g members-wire,, platform, p , yoke, y , and pyramidal-and py are discussed in the following g paragraphs. p g p Wire Balances: One or the earliest types of wind tunnel balance was the wire balance, balance similar in principle to Fig. Fig 26. 26 Spring scales were not used for the balance output since their deflections would change the model attitude. attitude The model usually was mounted inverted so that aerodynamic lift added to the weight to prevent unloading the wires as the resulting tension can never be allowed to diminish to zero. zero With this type of balance there was a large tare drag on the wires that was difficult to assess accurately. accurately The wires i tended t d d to t break, b k which hi h could ld lead l d to t the th lloss off th the model. d l Wi Wire balances b l turned t d outt tto b be much h lless robust b t and d versatile til than th the th alternatives lt ti and d have h nott been b used d extensively t i l since i the th very early l days d off aeronautics. ti Platform,, Yoke,, and Pyramid y Balances: Currently, y, most external balances provide p struttype yp mounting g of models. These balances provide p mechanisms for changing g g the angle g of attack and yaw and transmit the model loads down into a system of linkages that separate them into force and moment components. Such an apparatus is shown diagrammatically in Fig. Fig 28 28, and a linkage system is shown in Fig. Fig 29. 29 The general massiveness of a balance structure may be seen in Fig Fig. 30 30.
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l
Fig. 28 Greatly simplified diagrammatic sketch of balance components
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Fig 29 Balance linkage Fig. linkage. Lift linkage (not shown) is beneath the roll table
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l
Fig 30 Massiveness of a wind tunnel balance is well illustrated by this photograph of a Fig. balance designed for a 150 150-mph mph wind tunnel with a 9 9-ft-diameter ft diameter test section. section D ring this early During earl set setup p in the factory factor the load members have ha e been dropped in place l without ith t going i through th h their th i respective ti windshield i d hi ld supportt bases. b As A shown, h th the balance b l has h approximately i t l 45O negative ti yaw
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l Tracing g the p pathwayy followed byy the loads from model to measuring g unit (Fig. ( g 28), ) we see first that the model as illustrated is supported pp on two front-load members or "struts" and a tail strut. The struts,, in turn connect to the inner part p of a floating g ring g frame that is free to turn (model yaw) and a mechanism is provided to raise or lower the tail strut to produce model pitch pitch. The outer pan of the floating frame is held in place by a system of struts that are specially designed to be strong in tension and compression but very weak in bending. bending These struts separate the components of the load by means of a linkage system and feed them into the measuring unit or output transducers. transducers Above the floating frame is a fairing turntable on which are mounted the windshields that minimize the direct aerodynamic forces on the support struts struts. The load turntable is tracked by the f i i turntable fairing t t bl through th h the th use off a servomechanism h i arrangement. t And, A d as the th fairing f ii t t bl rotates. turntable t t Th The windshields i d hi ld are rotated t t d iin the th opposite it sense by b a gear-driven di mechanism h i so that th t th they remain i parallel ll l to t the th airstream. i t In I some balances b l the th tail-strut t il t t fairing g is moved up p and down to keep p the exposed p length g of tail strut constant as the angle g of attack is changed. g The windshields are connected electricallyy so that upon p contact with the load members they y activate fouling g lights g and/or audible signals g so that the malfunction may be noted and corrected. The linkage system by which the force and moment components are separated have gradually worked into three different fundamental types. types These are named platform, platform yoke, k and d pyramidal, id l according di to t th the manner iin which hi h the th main i system t is i assembled. bl d
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l Platform Balance: The platform balance (Fig. (Fig 31) utilizes either three or four legs to supportt the th main i frame. f For F the th three-legged th l d type, t the th forces f and d moments t are L=-(a+b+c)
D=d+e I=2(a-b)(w) n=2(e-d)w S =-f
m=cx
Platform balances are widelyy used,, rugged gg and naturallyy orthogonal, g , theyy mayy be constructed and aligned with a minimum of difficulty. difficulty But they also have disadvantages: (1) the moments appear as small differences in large forces, forces an inherently undesirable arrangement; (2) the h b balance-resolving l l i center is i not at the h center off the h tunnell and d the h pitching moments must be transferred; and (3) the drag and side force loads put pitching and rolling moments on the load ring. ring These interactions must be removed from th fifinall data. the d t Some S off the th disadvantages di d t are ameliorated li t d by b the th ease off computation t ti today y as compared p to several decades ago. g Also,, applications pp such as automobiles and surface marine vehicles will be mounted on or near the floor which is near the resolving center for a yoke balance. balance
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Fig. 31 Basic layout of a platform balance
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l Yoke Balance: The yoke balance (Fig. (Fig 32) offers an advantage over the platform b l balance in i that th t the th moment-resolving t l i center t is i near the th center t off the th tunnel. t l However, H the inherent design of the yoke leads to bigger deflections than the platform balance, Particularly in pitch and side force. force Because the balance frame must span the test section ti iin order d to t gett th the ttwo upper d drag arms iin their th i positions, iti the th yaw lever l arm is i exceptionally p y long. g The high g supporting pp g pillars p are subject j to large g deflections. Once again the final forces must be summed up: The drag is the addition of three forces and the lift is the sum of two forces in the variant shown forces, shown. The yoke balance as shown h here h brings bi out the h pitching i hi moment in i the h drag d system instead i d off in i the h lift. For the yoke balance, the forces and moments are L ( +b) L=-(a+b)
D +d+ D=c+d+e
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Fig 32 Basic layout of a yoke balance Fig.
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l Pyramidal P id l Balance: B l Th The complaints l i t usually ll heard h d about b t the th platform l tf and d yoke k balances b l are largely g y overcome byy the ingenious g engineering g g of the pyramidal py type. yp However,, as usually happens, happens additional difficulties are added. added These are the advantages: The pyramidal balance reads the moments about the resolving center, center and the six components are inherently i h l separated d and d read d directly di l by b six i measuring i units. i No N components need be added, subtracted, or multiplied. The difficulties involved in reading the small differences in large forces are eliminated, eliminated and direct reading of the forces and moments t simplifies i lifi the th calculations. l l ti Note N t that th t this thi is i less l off an advantage d t today t d than th it was several decades ago g when these systems y were being g intensivelyy developed. p Several criticisms of the pyramidal balance are warranted. warranted The alignment of the inclined struts is so critical that both the construction and the calibration of the balance are greatly complicated. Furthermore (and this appears quite serious), deflections of the inclined struts may so change their alignment that the moments are not accurate. accurate This effect ff t mustt b be th thoroughly hl iinvestigated ti t d d during i th the calibration lib ti off th the b balance. l Th The manner in which the pyramidal py balance separates p the moments is not simple p and it behooves the engineer and the student to approach the setup using an elementary truss system. system
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component p t Balances B l Consider a truss in which two legs arc jointed (Fig. (Fig 33). 33) The force D, D acting through the pin i joint j i t O, O produces d only l tension t i in i OE and d compression i in i OF. OF No N force f is i registered i t d att A. However, the force G, not acting through O, produces bending in OE, and OE would collapse unless the force A=aG/b were present. present If G and b are known, known the size of the f force A determines d t i the th point i t off action ti off G. G IIn thi this manner, if G were a known k drag d f force, its pitching p g moments about the resolving g center 0 would be determined byy the force A. Though the previous example illustrates the principle of the pyramidal balance in actual practice i a considerable id bl revision i i is i required. i d In I order d to prevent the h llegs off the h pyramid id from being in the airstream. They are cut off at what would be c in Fig. 33. The truncated legs are then carefully aligned so that their extensions pass through a common point. point The complete l t setup t iis illustrated ill t t d in i Fig. Fi 34. 34 The Th forces f and d moments t are Lift=total Lift total weight on lowest table
Drag=D Drag D
Side force=force C
Pitching g moment=- PXf
Rolling g moment=RXf
Yawing g moment=YXa
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Wind Tunnel T nnel Measurements Meas rements F Force M Measurements: t Six Si C Component t Balances B l p
Fig. 33 Two Two-dimensional dimensional schematic showing the principle of the pyramid balance
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Fig. 34 Pyramidal or virtual center balance
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Wind Tunnel T nnel Measurements Meas rements I t Internal l Balances B l IInternal t l balances b l are used d extensively t i l for f complete l t model d l workk and d even more extensively for measuring loads on parts of models. A six-component six component internal must address all the problems faced by a six-component external balance, balance but it must accomplish li h its it purpose within ithi a tightly ti htl specified ifi d and d highly hi hl restricted t i t d volume l and d shape. p Even the best internal balance cannot match the overall dynamic y range g and component independence of most external balances balances. But the option of matching the properties of an internal balance closely to a specific need greatly expands the options available il bl when h planning l i an experimental i t l investigation. i ti ti The two transducer types that are most widely used for internal balances are strain gages and d piezoelectric i l t i elements. l t Of these th two, t strain t i gages far f outnumber t b piezoelectric devices. The reason for this is that although p g piezoelectric p devices are extremely stiff and have outstanding frequency response response, they do not measure steady loads well. well In practice, practice the term internal balance and internal strain gage balance are effectively ff i l synonomous.
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Wind Tunnel T nnel Measurements Meas rements I t Internal l Balances B l Basic B i Aspects: A t There Th are ttwo basic b i types t off elements l t in i an internal i t l balance: b l force f elements and moment elements. In any six six-component component system there will be three force units and three moment units units. Force-measuring g elements employ p y either a cantilever beam or a column arrangement. g An eccentric column provides greater sensitivity but also allows more deflection, deflection as does a single cantilever. cantilever The choice might well depend on the particular balance size and d arrangement needed d d ffor a specific ifi model. d l Th The axial i l force f “ “gage" " shown h in i Fig. Fi 35a 35 is one of the most common types. This unit can be made very sensitive by sizing the flexures but since the model is attached to the cage, flexures, cage it is subjected to the relatively l large normall forces. f One O might i ht thus th expectt (see ( Fi 35b and Fig. d 36) an obvious b i interaction i t ti to occur because of the deflection of the flexures. A unit designed g at David Taylor y Model Basin (DTMB), (DTMB) shown in Fig. Fig 35b, 35b has reduced this kind of interaction to a minimum. minimum All forces except axial are carried by webs as shear or direct tension. tension A rod transmits the axial force f to a cantilever beam mounting g the gages. g g The arrangement g for f normal force f readout is shown typically in Fig. 37. In this case, the wiring is arranged so that the difference of the two moments electrically. electrically
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Wind Tunnel Measurements I t Internal lB Balances l
Fig 35 Axial force gage Fig.
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Fig 36 Normal force interaction on axial force gage Fig.
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Wind Tunnel Measurements I t Internal lB Balances l
Fig. 37 Normal force and pitching moment gage arrangement: (a) differential circuit for normal force: (b) summing circuit for pitching moment
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Wind Tunnel T nnel Measurements Meas rements I t Internal l Balances B l Since the Si th normall force f iis equall to t the th difference diff off the th two t moments t divided di id d by b the th distance between gages, g g the unit may y be calibrated directlyy in terms of applied pp normal force. o ce It is s important po a that a bo both gage sstations a o s have a e the e sa same e sec section o p properties, ope es, Ily, a and d matched gages gages. The greater the gage spacing, spacing the more accurate the normal force readout If Mf and Mr, are the front and rear moments readout. moments, then the normal force N is given by (Mr - Mf)/d, )/d where d is the spacing between gages. gages It is noted that the same arrangement may be b used d to t measure side id force. f Pitching or yawing moments may be measured by the same gage arrangement discussed above b except, t as shown h in i Figure Fi 7.48, 7 48 the th b bridge id is i connected t d as a summing i circuit. i it The Th differential circuit employed p y for normal force will also yield y the moment if the moment reference point is between the two gage stations, for, then, Mref=Mf+Xref ((Mr-Mf))/d where h Xref is i measured d from f th the ffrontt gage station. t ti A Another th way iis indicated i di t d in i Fig. Fi 38, 38 where the p pitch gages g g are "stacked" and located between the normal force gages. g g For rolling moment a torque tube or a double double-beam beam type with gages mounted on the side faces of the beams can be utilized utilized. There are many mechanical variations in internal balance design but the basic arrangement of strain gages attached to flexures is common to all. design, all
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Wind Tunnel Measurements I t Internal lB Balances l
Fig 38 Three component strain gage balance (NASA design) Fig.
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Flow Visualization Introduction It is difficult to exaggerate the value of flow visualization. A reasonable mental image of a flow about a body is almost always necessary for a person to have a useful understanding of an aerodynamic or hydrodynamic problem. This is true whether the approach pp is strictly y theoretical,, mainly y experimental, p , mainly y computational, p , or some combination as is always the most effective. The ability to see flow patterns on and around a device under investigation often gives insight into a solution to an aerodynamic problem. This sometimes happens when the pattern of flow exhibited by the experiment or produced by a computation is in some significant way different from the mental image that the aerodynamicist had formulated. Or, perhaps, the aerodynamicist realized there were two or more possibilities and the experimental evidence resolved the uncertainty. An important reason for the wide appreciation of computational t ti l flfluid id d dynamics i iis th thatt th the processes required i d tto reach h any solution, l ti whether or not it is consistent with physical reality, also strongly support very visual presentations of very detailed results. These visual representations of detailed data sets are very memorable, as are many types of flow visualizations in physical experiments, and both are thereby quite useful for similar reasons.
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Flow Visualization Introduction Classic flow visualization for low-speed flows is done by putting something that is visible into the flow at convenient locations and watching how the something, called a tracer, moves. The flow is inferred by the motion of the tracer. Tracer methods are the oldest and most commonly applied means of flow visualization. It seems likely that every human being has seen naturally occurring smoke or cloud formations being convected by movement of air and thereby conveying information about the motion of the air. We must, however, take care to relate how these "pictures" imprinted on our consciousness are related to our mathematical equations, which we use to quantify flow phenomena. In addition, it is necessary to investigate to what degree various sizes and types of tracer particles actually follow the flow. For the moment, assume that it is possible to put "tracer" particles into the flow at any desired location and that these particles are then convected perfectly along with the flo or that we flow, e simply simpl have ha e the capabilit capability to "see" an any small element in the flo flow that we choose. We have the following definitions. Pathline: The path of a point or particle convected with the flow is called a pathline. If we could release a tracer particle at any selected point and record its subsequent path this would be a pathline. If we knew the functions indicated by (4), we could construct pathlines by the parametric relations.
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Flow Visualization Introduction
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Flow Visualization Direct Flow Visualization Techniques We consider two broad categories of flow visualization. The first is surface flow visualization and the second is flow field visualization. Strictly speaking surface now is also a flow field, but we will use these terms to distinguish between on body and off-body fields. fields The methods considered for direct-surface flow visualization include tufts, tufts oil now, ink dot, china clay, and liquid crystals. The methods considered for off-body visualization include smoke injected in several ways, helium bubbles, and streamers. Methods of Recording Direct Visualizations: There are basically four methods of recording direct visualizations. The first, historically most important but least permanent, method is for the engineer to observe with his or her eyes. Because of depth perception, one can see three-dimensional patterns and one always sees an evolution of the phenomena,, not jjust the final average p g state. However,, there is no direct p permanent record in a form that can easily be put into a report or paper. It is possible, however, to sketch the patterns as they are observed. To do this efficiently, one needs to prepare in advance a basic drawing of the model on which streamlines or separated regions will be sketched when the tunnel is running. An advantage of doing sketches is that the mental process increases the likelihood of forming cognitive maps that capture the flow patterns.
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Flow Visualization Direct Flow Visualization Techniques Other common methods of recording the results of flow visualizations are by film, either still or movie, by video recording using a standard VCR format and by digital recording of digitized video, either still or movie. These methods produce a two dimensional projection of a three-dimensional phenomenon. phenomenon In principle, principle two or more cameras can be used to obtain multiple two-dimensional projections that contain sufficient information to reconstruct the three-dimensional image. In practice this has rarely been achieved and is not, at the present time, available for routine use. The state of the art in the needed technology indicates that such capabilities will be available in the near future. A situation in which such a system would be of great use would be when using smoke or helium bubbles to trace flow streamlines past a model. The photographic methods while requiring more time or developing and printing, yield higher resolution. Video has the advantage of instant replay. The available resolution for video equipment is rapidly increasing and is adequate for most aerodynamic work today.
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Flow Visualization Surface Flow Visualization Techniques: Tufts Information about the flow on the surface of an object being studied is usually most critical. Many times, the flow off the body is of interest primarily in order to understand the flow features on the surface. Key aspects of surface flows that may be investigated using visualization techniques include stagnation point location location, separation lines lines, location of boundary layer transition, characteristic unsteadiness, extent of separation zones, and types of critical points and their locations. The simplest and most frequently used method for surface flow visualization is to attach tufts to the surface of interest. The tufts must be of light, flexible material that will align itself with the local surface flow as a result of direct aerodynamic force. The most commonly used material is light yarn with weights and lengths chosen according to model size and test speeds. p Very y small monofilament has also been used. There are also polyester and cotton sewing threads, such as Clark's O.N.T. mercerized cotton No. 60, which can be treated with a fluorescent material. The thread is a multiple-strand material and tends to unravel with time. Tufts do affect the aerodynamic forces to some extent as we will show, but there are many situations in which the method is so easy and economical that it is the first choice.
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Flow Visualization Surface Flow Visualization Techniques: Tufts Two basic methods of attaching tufts to a surface are by scotch tape or by glue. When tape is used, the tufts are usually made on a "tuft board." The tuft material is strung back and forth around pins, then the tape is applied to the tufts and the tuft material is cut at the edge g of the tape. p This g gives a length g of tape p with tuft attached that is applied pp to the model (Fig.1). The model surface is cleaned with naphtha or other solvents to remove oil so that the tape will hold under the adverse conditions of high-speed flow. When tufts are glued to the model, a nitrocellulose cement such as Duco is used, thinned 50% with acetone or methyl ethyl ketone. Often 10% pigmented lacquer is added both to obscure the portion of the tuft under the glue and to make the glue dots visible by using a contrasting color. The glue dots are kept as small as possible. Tufts readily show where flow is steady and where it is unsteady. Regions of complete separation and buffeting flow are readily identified. The resolution of the determination is of the order of the spacing of the tufts. The possibility of significant influence of the tufts themselves on the flow is very high and must always be kept in mind. This can be investigated by removing tufts upstream of indicated flow separation. An example using No.6 floss (crochet yarn) is shown in Fig. 2. An example using No. 60 thread tufts is shown in Fig. 3 for a transport aircraft wing.
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Flow Visualization Surface Flow Visualization Techniques: Tufts
Fig. 1 Methods of taping tufts to model and a tuft board. The pattern shown as A is used for high-speed tests
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Flow Visualization Surface Flow Visualization Techniques: Tufts
Fig. 2 No. 6 floss (crochet yarn), white light source: =27.3O. Tufts taped to wing as in method B of Fig. 1.
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Wind Tunnel Techniques
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Flow Visualization Surface Flow Visualization Techniques: Tufts
Fig. 3 No. 60 thread tufts glued to wing, UV source: =27.3O
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Flow Visualization Surface Flow Visualization Techniques: Minitufts The glue technique is used for minitufts. These tufts have the least effect to the aerodynamic data and thus are often left on the model. The tuft material is mono filament nylon that has been treated with a fluorescent dye. Two sizes are used: 3 denier (diameter 0.02 mm, 0.0007 in.) and 15 denier (diameter 0.04 mm, 0.0017 in.). The dye used is Leucophor EFR liquid in a concentration of 1% in water with 2% acetic acid added. The tuft material is wound on an open wire reel and immersed in the dye for 15 min at 82.2°C (180°F) with frequent agitation. After drying for at least 1 hr the tuft material is wound onto small spools. During this step the material should be wiped with tissue pads to remove loose fluorescent powder that can transfer to the model surface in irregular patterns. When gluing tufts to a model, a square grid is used (typically about 0.75 X 0.75 in.). The tuft material is taped to the wing undersurface and then wrapped around the wing in a chordwise direction direction. The material is in the chordwise direction on the top and moves diagonally across the wing on the bottom surface. As an alternate, the tuft material can be taped at both the leading and trailing edges. After the tuft material is applied, it is glued using a hypodermic syringe with a fine needle (a coarse needle can be partially closed with pliers). As the desired size drop of glue forms on the needle, touch it to the surface and pull away quickly.
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Wind Tunnel Techniques
July-Nov. 2016
Flow Visualization Surface Flow Visualization Techniques: Minitufts After the glue dries the tufts are cut just ahead of the glue spot of the next tuft. The model surface should be cleaned before the tufts and g glue are applied pp using g Freon or chlorinated hydrocarbons. The monofilament nylon minitufts acquire static charges. These can be neutralized by the use of antistatic solutions or the antistatic material used in home dryers. The minitufts are viewed and photographed in ultraviolet light as this is the way to make the small monofilament most visible. Threads can be similarly treated and can be photographed in either ultraviolet or white light. Minitufts provide the same type of information as larger tufts. They can provide Fig. 4 Fluorescent minitufts, UV source: =27.3O. greater resolution and have less influence on Wing is marked with a fluorescent felt marker pen the flow. An example of surface visualization Compare stalled region near tip with Figs. 3, 5 & 8. using fluorescent minitufts is shown in Fig. 4.
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Flow Visualization Surface Flow Visualization Techniques: Oil Oils and other viscous fluids are used to show the surface flow. The selected material is usually spread on the areas of interest with a paint brush. It will then flow under the influence of shear stress from the air stream and gravity. Since inclined surfaces are almost always of interest, the mixture needs to have viscosity sufficient so that it will not flow rapidly under the influence of gravity. The flow speed of the air must then be sufficient to impress shear stress large enough to cause the oil to flow and reveal the surface patterns within an acceptable time. This is typically of the order of 10's of seconds after the tunnel is brought to speed. It is difficult to use oil flow on vertical surfaces at air speeds less than 100 mph and 150 mph is much better. The most common material for oil flow is petroleum lubricating oils. These materials are messy to clean up afterward, both on the model and more so in the t tunnel. l Another A th material t i l that th t works k as wellll as oilil when h t t d with treated ith a fluorescent fl t dye is polyglycol. At high CL's or high surface velocities this material may have too low viscosity, making it difficult to use. This material can be cleaned up with soap and hot water. When cleaning the tunnel after extensive oil flow runs, a portable set of ultraviolet fluorescent tubes is most useful. If the oil flows too slowly, it is thinned with naphtha, and if it is too thin, 60-70W oil is added.
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Wind Tunnel Techniques
July-Nov. 2016
Flow Visualization Surface Flow Visualization Techniques: Oil The viscosity of the mixture is adjusted by trial and error for each application. The color of the oil needs to contrast with the color of the model surface. A widely used method is to add a fluorescent dye to the oil and illuminate it with ultraviolet lights, as is done for the fluorescent mini tufts. In this case the model color is not too critical so long as it does not reflect strongly under the ultraviolet illumination. illumination A light blue works very well with a dye that provides fluorescence in the yellow region of the spectrum. This is the most commonly used combination. An example result of oil flow visualization using 40W motor oil treated with a very small amount of fluorescent dye is shown in Fig. 5. Oil can be made white by mixing titanium dioxide into it. This can be applied to a black model and ordinary light used for viewing and photography. This is sometimes preferable to the installation of black lights and the subsequent requirements on light management. Two examples are presented. Figure 6 shows a low aspect-ratio wing at high angle of attack. Figure 7 shows the upper surface of an automobile. Photographs can be taken after the tunnel is turned off, off but the available time is short even on horizontal surfaces as the oil will flow under gravity. China Clay China clay is a suspension of kaolin in kerosene. The fluid is applied with a paint brush, usually with the model set at the desired attitude. The tunnel is started as quickly as possible after the model L painted. When the mixture has dried, photographs can be taken after the tunnel is shut down because the pattern does not change rapidly with time. This is similar to the Fales method with the kaolin substituted for lamp black. An example of visualization using china clay is shown in Fig. 8.
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Flow Visualization Surface Flow Visualization Techniques: Oil
Fig. 5 Oil applied at =0O, tunnel started and brought close to speed, model pitched to =27.3O
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Flow Visualization Surface Flow Visualization Techniques: Oil
Fig. 6 Oil flow on a low aspect ratio rectangular wing (Clark Y-14 wing, aspect ratio 4, =25.4O, Re=360,000)
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Flow Visualization Surface Flow Visualization Techniques: Oil
Fig. 7 Oil flow on an automobile, yaw 0O
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Flow Visualization Surface Flow Visualization Techniques: China Clay Technique
Fig. 8 China clay applied at =0O, tunnel started and brought close to speed, model pitched to =27.3O
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Flow Visualization Effects of Tufts, Minitufts, China Clay, and Oil Tufts can affect the aerodynamic loads on a model. In Fig. 9 a lift curve near stall shows the effects of various tufts on the data. The g glued minitufts and No. 60 thread tufts consisted of about 900 tufts. The two taped tufts consisted of about 300 tufts. The data are an average of five runs for each set of tufts. The mini tufts and the glued No. 60 thread have the minimum effect on lift. The effect of the tape can be seen by comparing the two sets of No. 60 thread tufts. The six-strand floss tufts are similar to the tufts made out of yarn. The three different tuft types can b seen att =27.3 be 27 3O in i Figs. Fi 24 2-4.
Fig. 9 Effect of various tufts, china clay and oil flow on lift curve near stall
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Wind Tunnel Techniques
July-Nov. 2016
Flow Visualization Sequencing Attitude and Speed Setting The usual procedure for china clay visualization is to set the model at the desired angle of attack and bring the tunnel up to speed. On a wing with a slotted leading-edge and/or trailingedge flaps. This can result in erroneous aerodynamic data and flow visualization due to flow separation in the slots at low Reynolds numbers during the tunnel acceleration. This is shown by the oil flow and china clay data points y =27.3O in Fig. g 9. Similar data beyond were obtained on the clean wing. Figure 10 shows a china clay flow visualization for this test method.
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Fig. 10 China clay applied at a=27.3O, tunnel started and brought to close to speed. Leading edge slotted flap is stalled. Compare with Figs. 4 and 8
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Flow Visualization Symmetry and Hysteresis An example of a very interesting flow phenomenon is given in Figs. 11 and 5.12. Although it is common to assume that symmetric boundary conditions produce symmetric flows, there are many counter examples. Any flow that includes large regions of separation may well exhibit asymmetry of the mean flow as well as asymmetry of the instantaneous flow even if the solid boundaries are sensibly symmetric. This can lead to results such as that illustrated here in which the integrated forces and moments exhibit random switching or it can lead to hysteresis in which the forces and moments will be dependent on the time history of the altitude. Flow visualization methods are often important tools in identifying the flow structures associated with such events. Flow Topology: Topological concepts are slowly gaining recognition as important to the study of complex flow phenomena such as bluff body flows and flows about wings and aircraft at high angles of attack. The concepts are providing ways to understand the structure of flow fields and to tie the structure of a flow field to the topology of the surface flow on the test article. As oil flow techniques provide very fine detail of the surface flow, this is a preferred technique for studies of surface flow topology.
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Wind Tunnel Techniques
July-Nov. 2016
Flow Visualization Symmetry and Hysteresis
Fig. 11 Oil visualization of flow on an airplane at =14O, that has flow on right wing attached and on the left wing separated
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Flow Visualization Symmetry and Hysteresis
Fig. 12 Here the flow on the left wing attached and on the right wing separated. The model is not moved. The flow state alternated between the two conditions shown at random intervals in the range of a few seconds at a chord Reynolds number of about 2x106. This produces a very strong rolling moment that changes very abruptly and randomly.
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Wind Tunnel Techniques
July-Nov. 2016
Flow Visualization Some Procedural Details Correlation of Balance Data and Flow Visualization: The generally accepted practice when using the older, larger, six-strand floss for tufts is not to take balance measurements, or at least not to consider them to produce good data when tufts are applied to the model model. Figure 9 explains why this is the practice practice. Both the lift curve slope and maximum lift are greatly reduced. An advantage of minitufts is that their effect on the data is minimal; hence they can be left on the model. Oil and china clay also show minimum aerodynamic effects. During this comparison test of tufts at the model's minimum drag (a lift coefficient of about 1.0). The model drag decreased as tufts from the mini to No.6 floss were added to the left wing with the horizontal tail on and the reverse happened with the tail off. The tufts apparently change the wing's span load distribution. However, old practices die slowly, so force data are often not taken during surface flow visualization. This, then, can lead to an improperly established flow field and the possibility of misleading now visualization, especially near stall and, oddly enough, near minimum drag. If, however, force data are taken before and during the flow visualization run, the error may be detected and thereby possibly avoided.
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Flow Visualization Effects of Tufts, Minitufts, China Clay, and Oil Tufts can affect the aerodynamic loads on a model. In Fig. 9 a lift curve near stall shows the effects of various tufts on the data. The g glued minitufts and No. 60 thread tufts consisted of about 900 tufts. The two taped tufts consisted of about 300 tufts. The data are an average of five runs for each set of tufts. The mini tufts and the glued No. 60 thread have the minimum effect on lift. The effect of the tape can be seen by comparing the two sets of No. 60 thread tufts. The six-strand floss tufts are similar to the tufts made out of yarn. The three different tuft types can b seen att =27.3 be 27 3O in i Figs. Fi 24 2-4.
Fig. 9 Effect of various tufts, china clay and oil flow on lift curve near stall
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Wind Tunnel Techniques
July-Nov. 2016
Flow Visualization Relative Advantages of Tufts Tufts have a great advantage in terms of productivity. Once the tufts are installed, the model can be repositioned and the indications studied visually and photographed for as long as desired. The model can then be simply p y moved to a new condition and the p process continued. An example of a series of tuft photographs for a sequence of conditions for a powered tilt wing half model is shown in Fig. 13. The oil and china clay method produce patterns of limited duration in time for a single operating condition. Then the tunnel must be stopped and the material reapplied for each new condition to be visualized. Tufts are sometimes used for broad looks over a wide range of conditions with the more detailed techniques then applied for a smaller and selected set of conditions that have been found to be most critical.
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Flow Visualization Relative Advantages of Tufts
Fig. 13 Visualization of flow on a powered wing model in a series of conditions using tufts
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Wind Tunnel Techniques
July-Nov. 2016
Flow Visualization Boundary Layers and Surface Shear Stress Often the most important information being sought by flow visualization methods is a definition of the locations of transition from laminar to turbulent boundary layers and the locations of any separation regions. In the previous section, we have shown several results that illustrate separated flows flows. The location of transition between laminar and turbulent now cannot be determined by tufts and is difficult with china clay. However, oil flow, sublimation methods, infrared images and liquid crystals can be used to locate transition. Oil Flow Detection of Transition: The basis for detecting boundary layer transition by viewing oil flow patterns is the increase in wall shear stress when a boundary layer transitions from laminar to turbulent. The result is that the oil is swept away faster in the region where the boundary layer is turbulent. Transition indicated by oil flow visualization is shown in Fig. 14. The shear stress at the leading edge of a surface is high even under a laminar boundary layer so a typical t i l pattern tt is i th thatt the th oilil iis sweptt away rapidly idl att th the lleading di edge d with ith a gradual d l lessening of the scrubbing as the laminar layer thickens and then severe scrubbing downstream of transition. The oil pattern given by a laminar separation bubble with turbulent reattachment can be seen just downstream of the leading edge of the wing in Fig. 6. In some cases a subcritical bump in the oil can cause a wake, which can confuse the transition location. This does not occur with sublimation.
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Flow Visualization Boundary Layers and Surface Shear Stress
Fig. 14 Oil flow visualization showing natural transition at approximately 40% chord; CL=0.28, Re=1,26,000 based on average chord
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Flow Visualization Boundary Layers and Surface Shear Stress Sublimation: In one sublimation technique a mixture of naphthalene and a carrier such as fluorine, acetone, or methyl ethyl ketone is sprayed on the model using a standard air spray gun. Note that the last two can remove many paints. The operator must wear a respirator p mask when doing g this. The mixture will leave the model surface white and therefore works best on a black or dark surface. The turbulent boundary layer will scrub the mixture off. A natural transition is shown on a wing using naphthalene in Fig. 15. Liquid Crystals: Liquid crystals that undergo changes in reflective properties as they are exposed to shear stress can be used as detectors of transition. This method has recently been developed as a quantitative method as well as a method of visualization. Infrared Thermography: The basis for this technique is that a surface at a temperature different from the tunnel stream will have faster heat transfer from the region under a turbulent boundary layer than from the region under a laminar layer. layer The resulting small temperature differences can be observed using commercially available infrared cameras. The results will vary with model structure and heat transfer characteristics so it is advisable to check results for a particular model type by observing artificially tripped layers. It should be kept in mind that it is the relative temperature, not the specific value of temperature that gives the indication of transition.
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Flow Visualization Boundary Layers and Surface Shear Stress
Fig. 15 Naphthalene-fluorine flow visualization showing natural transition at approximately 40% chord; CL=0.28, Re=1,26,000 based on average chord
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July-Nov. 2016
Flow Visualization Boundary Layers and Surface Shear Stress Ultraviolet Fluorescence Photography: Ultraviolet fluorescence photography is used when the medium used for visualizing the flow has been treated with a dye that radiates in the visual spectrum when excited by ultraviolet light. The two common cases already illustrated are the fluorescent minitufts and fluorescent oil oil. The wavelength of the ultraviolet light is 320-400 mm, and it is transmitted by optical glass. There are three sources that are typically used to produce the ultraviolet light. They are special fluorescent tubes (black lights), mercury vapor lamps, and photo flash units. In wind tunnel use, the first two light sources are used enable the test engineer to observe the flow, and the flash units are used to take still photographs of the flow when desired. Fluorescent tubes and mercury vapor lamps in general do not have a high enough light intensity to allow photographs without a very long exposure. Because fluorescent material emits light in the range, the tunnel test section must be shielded from visible light. Both mercury vapor lamps and the flash units also emit visible light. Thus they must be equipped with an exciter filter that will transmit ultraviolet light and absorb visible light. Glass filters that accomplish this are Kodak Wrattan filter No. 18A or Corning No. 5840.
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Flow Visualization Boundary Layers and Surface Shear Stress For larger tunnels the flash lamps are usually studio units marketed for commercial photographers. The flash lamps should be able to handle 2000 W/s per flash as rated by photographers; the units come with power supplies that can store energy in capacitors and have the necessary trigger circuits circuits. The reflectors for the flash units should be 10-14 in. diameter to be efficient. The Corning glass exciter filter comes in 6.5-in. squares, four of which can be glued together and built into a frame to cover the reflector. As an alternative, one glass filter can be used with the flashbulb without a reflector. This would require approximately one additional f stop. For research tunnels where the camera-to-subject distances are small, standard flash units and a Wrattan 18A exciter, which can be obtained in a 3.5-in. square can be used. Photographic film is sensitive to blue and ultraviolet light. The light reaching the camera will contain both the visible fluorescent radiation and reflected ultraviolet radiation. To prevent the ultraviolet from exposing the film, a barrier filler is attached to the camera lens. The barrier filter can be a Kodak Wrattan filter No. 2A, 2B, 2C, or 2E. These can also be obtained in 3.0-in.-square gelatin sheets.
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July-Nov. 2016
Flow Visualization Boundary Layers and Surface Shear Stress Focusing the camera with a fluorescent light source is usually no problem because this can be done using either the black-light fluorescent lights or the mercury vapor lamp with an exciter filter as a light source. It should also be noted that black lights that can be fitted to standard fixtures are several orders of magnitude cheaper than the mercury vapor light with its power supply and exciter filler. It is also possible to photograph the fluorescent material with some video cameras during the flash from the light source.
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Flow Visualization Flow Field Visualization Tuft Wands and Tuft Grids: The least expensive and at the same time a very versatile method is using a tuft wand, a long tuft on a pole useful for tracing now near the test article. If it is necessary that a person be in the runnel. He or she must wear goggles to protect the eyes from dust dust. The person should disturb the flow in the vicinity of the test article as little as possible. Fig. 16 shows a tuft wand in use. To make the flow pattern downstream of the test article visible, a tufted wire grid is useful, as shown in Fig. 17. Helium-Filled Soap Bubbles: To trace pathlines, which are also streamlines if the flow is steady, helium-y filled bubbles that have neutral buoyancy can be used. The bubbles are inserted ahead of the model and are photographed with a high-intensity light that passes through g the tunnel as a plane of light. g With proper photographic g exposure time the bubbles appear as streaks. Maximum tunnel speed for use of helium bubbles is about 60 mph (30 m/s). The bubble generators are available commercially.
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Flow Visualization Flow Field Visualization
Fig. 16 Investigating the flow structure of an automobile with a tuft wand
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Flow Visualization Flow Field Visualization
Fig. 17 Flow visualization by the grid and tuft method, yawed delta wing
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Flow Visualization Flow Field Visualization: Smoke Methods of Smoke Production: The most common method of flow field visualization after tufts is smoke, which can be produced in a number of ways, although a universally accepted "best" way has yet to be devised. Burning damp straw, rotten wood and tobacco to produce smoke is generally unsatisfactory, even though historic work was accomplished with smoke from such materials. The same is true of pyrotechnic smoke devices (smoke bombs). Chemical methods of producing smoke include both titanium tetrachloride and tin tetrachloride which produce smoke when brought into contact with moist air. These compounds produce the very best dense smoke filaments but the products are corrosive and can be used only in an open circuit facility that can be well vented to an appropriate location. A mixture of anhydrous ammonia and sulfur dioxide produces dense white smoke, odors, and, if the air is damp, sulfuric acid. A significant advantage g of the system y is the fact that the smoke can be started and stopped pp in a fraction of a second, which is not true of the alternatives. This allows the problems associated with it to be minimized if photographic records rather than real-rime human visual inspection can be paramount. Steam and liquid nitrogen produce a good dense smoke with no aftereffects but are very difficult to control and therefore seldom used.
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Flow Visualization Flow Field Visualization: Smoke A method reported by Shindo and Brask that works at velocities of over 100 mph is a vaporized petroleum product called Type 1962 Fog Juice, which is used in theatrical productions. The smoke generator consists of about 75 in. of 0.060-in. outside-diameter (wall = 0.01 in.) stainless steel tubing. For a probe the tubing is placed inside a 0.375in diameter steel tube and held by a collet about 8.0 in.-diameter 8 0 in. in from the end of the tubing, tubing which is bent 90° about 3 in. from the end. Ceramic beads are used to insulate the stainless steel tubing from the outer tube. To vaporize the fluid, 10-15 A is applied to the stainless steel tubing from the collet to a point about 60.0 in. away inside a non-heatconducting handle. This allows the stainless steel tube to expand. The power unit consists of a variac whose output is connected to the 230-V windings of a 1.5-kVA 115:230 transformer. The variac is used to control the temperature on the stainless steel tubing by applying 0-50 V. The tubing has about 3 resistance, so the current is limited to a maximum of 15 A. A
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Flow Visualization Flow Field Visualization: Smoke The fuel reservoir is airtight and has a pressure regulator used to set plant air pressure at about 30 lb/in2 to feed the fuel to the probe and a needle valve to control the fuel flow. Plastic tubing connects the reservoir to the probe. It takes some experience to obtain the desired volume of smoke which is affected by the air pressure fuel flow and voltage. pressure, voltage The following values arc approximate. approximate With the power switch off, set the variac at 70%, apply 30 lb/in2 to the reservoir, and crack open the needle valve. When a small stream of fluid comes out the end, turn on the power. When turning off, cut the power, and when a stream of fluid leaves the tube, shut the needle valve off. This is done to reduce carbon formation in the stainless steel tubing. If hot fluid is emitted, the temperature is too low; either increase the voltage or reduce the fuel flow. If the smoke pulsates, it is generally a sign that the air pressure is too low. The most widely used method for producing smoke at the larger wind tunnels today is a wand system similar to that described above with polyethylene glycol as the "fuel." A useful alternative to air pressure to drive the fuel through the wand is a peristaltic pump. These are sold by suppliers to chemistry and life science laboratories and have very tine volume flow control. An example of such a system is shown in Fig. 18.
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Flow Visualization Flow Field Visualization: Smoke The fuel reservoir is airtight and has a pressure regulator used to set plant air pressure at about 30 lb/in2 to feed the fuel to the probe and a needle valve to control the fuel flow. Plastic tubing connects the reservoir to the probe. It takes some experience to obtain the desired volume of smoke which is affected by the air pressure, fuel flow and voltage. The following values are approximate. With the power switch off, set the variac at 70%, apply 30 lb/in2 to the reservoir, and crack open the needle valve. When a small stream of fluid comes out the end, turn on the power. When turning off, cut the power, and when a stream of fluid leaves the tube, shut the needle valve off. This is done to reduce carbon formation in the stainless steel tubing. If hot fluid is emitted, the temperature is too low; either increase the voltage or reduce the fuel flow. If the smoke pulsates, it is generally a sign that the air pressure is too low. The most widely used method for producing smoke at the larger wind tunnels today is a wand system similar to that described above with polyethylene glycol as the "fuel." A useful alternative to air pressure to drive the fuel through the wand is a peristaltic pump. These are sold by suppliers to chemistry and life science laboratories and have very tine volume flow control. An example of such a system is shown in Fig. 18.
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Fig. 18 Smoke generator and delivery wand used at the Glenn L. Martin wind tunnel, University of Marylan
Wind Tunnel Techniques
July-Nov. 2016
Flow Visualization Flow Field Visualization: Smoke F N M Brown developed a system of smoke generation in which kerosene is vaporized and the vapor then cooled to the temperature of the airstream before being emitted into the stream. The systems are designed to be used with special purpose flow visualization wind tunnels rather than in larger general-purpose wind tunnels. They have been used in many studies. studies Using Smoke: Smoke filaments can be used to find key locations such as stagnation points. Smoke can easily show the extent of separated regions and the size of a separation bubble. Smoke can also be used to show and track strong features such as tip or leading-edge vortices from lifting surfaces that may impinge on downstream parts of a vehicle with deleterious effects. Strong stable light sources are required. If a periodic phenomenon is being investigated, then a stroboscope will be needed that can be synchronized to the period of the phenomenon being studied. Common subjects requiring such an arrangement are propellers and rotors. rotors Regions of separated flow can be detected in two opposite ways. Smoke filaments can be introduced upstream of the model and located in a series of positions so that the boundary of essentially undisturbed flow can be detected. Any volume into which smoke does not go readily is likely a wake region. The smoke wand can then be introduced directly into the apparent wake region. The smoke will fill the wake "bubble," thereby defining its extent.
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Flow Visualization Flow Field Visualization: Smoke Wire Smoke Wire: Another method of producing small discrete filaments of smoke at low velocities is described by Batill and Mueller. This is the smoke wire technique, and it appears to be limited to flows where the Reynolds number, based on wire diameter, does not exceed 20 or at velocities from 6 to 18 ft/s (2 to 6 ft/s). The limit is based on preventing the wake from the wire from disturbing the flow behind the wire, wire and the limit has been determined by experiment. This method uses a small-diameter wire that is coated with an oil. The best results were obtained using Life-Like model train smoke, which consists of a commercial-grade mineral oil with small amounts of oil of anise and blue dye added. The liquid-coated wire has 40-80 V AC or 40-60 V DC impressed across it. As the wire is heated, fine smoke streaklines form at droplets on the wire (approximately 8 lines/cm for a 0.003-in.-diameter wire). As the wire is heated. It expands and thus sags. This can cause problems with the accurate placement of the streaklines To avoid this, streaklines. this the wire as prestressed to about 1.5X10 1 5X105 lb/in2, which is near the yield point of type 302 stainless steel wire; thus the wire must be handled carefully. As the oil-treated wire produces smoke for periods of up to 2 sec. a timing circuit was used for the lights and cameras used to photograph the smoke. Since the smoke wire is limited to low-Reynolds-number tests, it is probably best suited to use in Small tunnels as it is difficult to run large tunnels at the required low velocities. An example of a smoke wire visualization is shown in Fig. 19.
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Flow Visualization Flow Field Visualization: Smoke Wire
Fig. 19 Streamlines past an airfoil visualized by smoke wire technique
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Flow Visualization Data Driven Visualization All the techniques in which information is processed by either analog or digital methods and produce flow images of some sort are placed in this category. The data may be a set of measurements such as the voltages from a set of pressure transducers attached to multihole probes along with the geometric data that give locations corresponding to each set of voltage data. Many other types of measurements are used in similar ways. Another example would be the luminosity data from a CCD camera recording the image of a model coated with pressure- or temperature-sensitive paint. The data may also come from a set of computations based on some flow model such as potential flow, Euler codes, or Navier-Stokes solvers. The demarcation between the class of methods addressed here and those that have been termed "direct visualization" is rapidly blurring as every image shown earlier already exists in digitized form and can be manipulated as a set of numerical data. The evolution is being driven by the continuing advance in our capability to acquire, manipulate, and present in various forms larger and larger data sets in smaller and smaller times.
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Flow Visualization Optical Methods of Flow Visualization Optical methods of flow visualization can be divided into the following methods: 1. 2 2. 3. 4. 5. 6.
Shadow method Schlieren method (parallel or focused, focused gray or color) Interferometry (classical, holographic) Electronic speckle interferometry and stereography Laser Doppler anemometry Particle Image Velocimetry
The optical methods are mainly used for visualization of compressible flows. flows Reference: Ristic Slavica, “A view in the invisible”, Theoretical and Applied Mechanics, TEOPM7, Vol. 40, No.1, pp. 87-119, Belgrade, Yugoslavia, 2013.
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Flow Visualization Optical Methods of Flow Visualization
Airflow around aerodynamically models, in optical sense, is a transparent environment with complex light refraction index. Light refraction index n in each point is the function of air density, which, on the other side, is the function of speed, pressure and air temperature. The relation between air density ρ(x,y,z) and refraction index n(x,y,z) is the Gladstone Gladstone-Dale: Dale: n=1+K 1 Kρ. The constant K has dimension of ρ-1 and it is different for each gas. According to Snell's law, a light ray, passing through inhomogeneous refracted field, is deflected from its original direction and a light path is different from that of undisturbed ray. If recording plane is placed in front of light ray, after disturbing media, three quantities can be measure: the vertical displacement of disturbed ray, the angular deflection of disturbed ray with respect to the undisturbed, the retardation of deflected ray, i.e. the phase shift between both rays [1]. Optical visualization methods are based on the recording one of these three quantities, or a combination of them. Shadowgraph used the first phenomenon, the Schlieren the second, and interferometry the last. The shadowgraph is sensitive to changes of the second derivative of density or refractive index ∂2n/∂y2, Schlieren to changes of density or first derivative ∂n/∂y, and the interferometry is capable to measure absolute density n changes. If, using the optical method, light refraction index n(x,y,z) in flow is determined, another physical parameter of tested environment, significant to aerodynamic testing, can be indirectly determined as well.
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Flow Visualization Optical Methods of Flow Visualization: Shadowgraph Method The oldest and the simplest of all optical methods for flow visualization is shadowgraph. Figure 20a shows the bow shock wave ahead of sphere in supersonic wind tunnel T-36 at M∞=1.86. The trace of the shock wave on the photo is a band of absolute darkness bounded on the downstream side by an edge of intense brightness. brightness The exact geometrical position of the shock front is the other edge of the dark zone. The shock wave represents a jump of the refractive index. The air density increases after the shock and the incident ray deviates to inside edge. Since the density in the disturbance is lower than in the surrounding field, (Prandtl-Meyer expansion fan at the sharp end of the nozzle) the bright band appears at the beginning of the shadow. The same result is obtained when the compressible boundary layers is visualized. Figure 20b is typical shadowgraph g p showing g flow around a spherical p tipped pp cylinder y mounted on flat p plate [[5]. ] Shadowgraph methods with short duration light pulses can be used for fine visualization of turbulent compressible flow.
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Flow Visualization Optical Methods of Flow Visualization: Shadowgraph Method
Fig. 20 Shadowgraph visualization of (a) flow around a sphere and (b) typical shadowgraph images showing spherical tipped cylinder mounted on a flat plate
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System As mentioned before, Schlieren method is sensitive to change of the first derivative of density ∂n/∂y, (or refractive index) and it can record the angular deflection of the disturbed ray with respect to the undisturbed in a transparent medium with local inhomogeneities. The Schlieren method is the most frequently used in aerodynamic laboratories, laboratories since it is relative simple and very useful method. If a parallel beam of light passes through air with density gradient normal to the direction of the beam, the beam is refracted towards the region of greater density. The most simply is the Schlieren system with parallel light through the test section of the wind tunnel. Töepler system in hypersonic wind tunnel, as the base of all other systems, is illustrated in Fig. 21. The modern Schlieren system uses color filter or phase optical elements instead of the knife-edge, g and have several p parallel, transparent, p colored strips p ((most often three colored sheets, red-blue-yellow or blue-green-red). The color filter can be consisted of four differently colored strips arranged in a square filter for visualize the grad n in two directions. If the flow is axis symmetric, complementary colors appear for the same event (compression or expansion) above and below the flow axis. The recorded pure colors and color combinations are a measure for the local direction of density gradient in the test section. Figure 21 shows components of Schlieren systems in T-34 hypersonic wind tunnels in MTI.
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System
Fig. 21 Schlieren system components, models in the test section of a hypersonic wind tunnel and TV camera with monitor
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System Figure 22 shows black and white color Schlieren images around a bullet at transonic Mach numbers and muzzle blast from a 22-caliber rifle. Attempts to increase the amount of information extractable from Schlieren images, various opaque filters with different geometries, as well as transparent phase and color filters are used. Figure 23 shows color Schlieren images around a blunt body and thin protruding probe mounted in front of a blunt body, used to reduce the drag and the rate of heat transfer, for M∞=1.86.
Flow visualization in two dimensional model of the supersonic rocket nozzle without and with vertical, different height barriers is tested by Schlieren method and the effects are presented in Fig. 24. The nozzle is designed for an exit Mach number of M∞=2.6. 2.6.
The classical Schlieren photos obtained with color Schlieren system are presented in Figs. 25 and 26. The flow around cone with tip angle of 15° and sphere with 40 mm diameter is tested in a supersonic wind tunnel for different Mach numbers and positions of color filters.
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System
Fig. 22 Black and white Schlieren images in a wind tunnel at (a) M∞=0.86 and (b) M∞=1.1 and (c) instantaneous image of bullet and muzzle blast from a 22-caliber rifle
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System
Fig. 23 Colour Schlieren images around (a) a blunt body, (b) a blunt body with a thin protruding probe and (c) a blunt body with a thin protruding probe at an angle of attack
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System
Fig. 24 Visualization of flow in a supersonic nozzle by Schlieren system (flow is from left to right): Continued_
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System
Fig. 24 Visualization of flow in a supersonic nozzle by Schlieren system (flow is from left to right): Concluded_
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System
(a)
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Fig. 25 Colour Schlieren images obtained in T-36 wind tunnel for around a cone with 15 degree tip angle at (a) M∞=1.02 and (b) M∞ =1.56
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System
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Fig. 26 Colour Schlieren images obtained in T-36 wind tunnel for around a 100 mm dia. sphere at (a) M∞=1.02 and (b) M∞ =1.56
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System The combined holographic interferometer and Schlieren devices, have been designed, made and tested for small supersonic and large trisonic wind tunnels. The device can be included in tests either as Schlieren system or interferometer. Improvements to the basic Schlieren system include the Rainbow Schlieren, Schlieren where a colored bull’s eye filter is used rather than a knife edge to quantify the strength of the refraction. The other variety of Schlieren methods is obtained by using laser as a light source. Figure 27 illustrates results from a Schlieren system with HeNe laser as a light source.
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Flow Visualization Optical Methods of Flow Visualization: Schlieren System
Fig. 27 Schlieren system with laser as light source in T-36 and Schlieren image around a cone at M∞=1.1
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Flow Visualization Optical Methods of Flow Visualization: Interferometry Interferometry is based on the fact that a change in density not only results in a refraction of the light, but also in a phase shift. In an interferometer parallel light is split into two beams. One of the beams enters the flow field, the other beam does not enter the flow field. When both beams are merged and projected on the same photographic plate, interference occurs when the phase of one of the beams is shifted by a change of density in the fluid flow. The most used classical interferometer in the wind tunnel tests is Mach-Zehnder interferometer (MZI). MZI can be applied to any case of gas flow investigations, where density difference becomes noticeable as: thermodynamic data, thermal conductivity of gases, dissociation, aerodynamic application, turbulence, wave or sonic booms.
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry Holographic interferometry is an optical method that makes possible complete flow field testing. The method is non-contact and it does not disturb flow field. It is used for testing of different objects and phenomena. The greatest advantage of holographic interferometry, in relation to Schlieren method, is the fact that it provides complete information in a single plate, allowing a postponement selection of specific types of flow visualization. The base of this method is holography, developed in last forty five years. If, on the some plate, the image of one object is recorded two times in different moments, in the process of reconstruction both images appear simultaneously and on the same place in the space. Object waves interfere because of mutually coherence (they originate i i t from f th same light the li ht beam b th t illuminate that ill i t the th hologram) h l ) and d the th interference i t f effects can be observed in the reconstructed object image. If no change occurs on object between first and second exposition, then there is no difference in images and there are no interference fringes. If certain difference appears, then the reconstructed image contains the system of interference fringes N that indicate that change.
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry Quantitative flow testing, using holographic interferograms is performed by determining the number of fringes N(x,y) in the field image with respect to a reference point of known density. After that, the index of light refraction n(x,y) and the air density ρ(x,y) can be calculated. calculated For the isentropic flow, flow relations between N, N n, n ρ, ρ pressure P, P temperature T, velocity V, and Mach number M exist, which van be used to determine required information. The simplest case for analysis is the 2D flow. For the processing of interferograms of axi-symmetrical phase objects, the method of inversion, based on the Abel transformation, is used. The experiment geometry is usually selected in order to simplify the mathematical representation of flow and changes occurring at the path of the laser light beam through the test section. Computer tomography is an important technique for reconstructing 3-D fields from holographic interferograms. Several techniques have been developed for computer tomography as: implicit methods (series expansion, discrete element representations), explicit methods (convolution method), and Fourier transform method. The choice of the best algorithm depends on structure of the density field, the amount and format of available data.
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry Holographic interferometer with parallel beams is at the same time Schlieren and shadow device. Figure 28 shows the schematic diagram of the experimental setup. During the experiments synchronized measurements were performed. Double exposition technique was used for holographic interferograms recording: wind off (when homogeneous flow field exists) and wind on (when there is complex flow field for testing). Stagnation pressure (P0), atmospheric pressure (Pa), and Mach number (M∞) were measured by the primary measurement system (PMS) in the wind tunnel, at the moment of recording hologram, shadow and/or Schlieren results.
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry
Fig. 28 Schematic of holographic interferometry/ Schlieren system and shadowgraphy in a supersonic wind tunnel
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry In order to demonstrate advantages of holographic interferometry in complex flow field testing, and compared with other classical methods, the series of experiments were performed in MTI supersonic and trisonic wind tunnel at flow velocity from M∞=0.7 to 3.24. The photos of holographic interferograms illustrate this method. Figure 29 shows some interferograms of different flows. The use of classical methods of nozzle flow field testing comprises the introduction of probe within the expansion region and holes perforation on nozzle surface. These interventions would significantly change the flow field. Optimization of this measurement is made by the holographic interferometry. In order to demonstrate and to compare complementary possibilities of optical methods in quantitative flow visualization, Prandlt-Mayer expansion tested by three optical methods is presented. Figures 30 a and b show the flow visualization around 90° corner end edge for supersonic nozzle M∞=1.56. The interferogram is recorded by double passing, collimated, object beam trough the wind tunnel test section. The shadowgraph is recoded on a holographic plate, because of collimated beams. The color Schlieren is recorded in the some time with holographic interferogram.
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry
Fig. 29 Holographic interferogram of flow around a (a) missile for M∞=1.56, (b) 90° cone at M∞=0.86 and (c) 2D cylinder at M∞=0.76 (flow is left to right)
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry
Fig. 30 Visualization of supersonic flow (left to right) around 2D 90O nozzle edge (Prandtl-Mayer expansion) M∞=1.56 a) shadowgraph and b) interferogram
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry The holographic interferograms were used for numerical calculation of flow field parameters in the vicinity of nozzle edge where the expansion fan is formed (Fig. 30c). The fringe number N was read from this hologram. Points in front of expansion fan have N=0, since the last fringe has N=17. The theoretical and experimental values of Mach number in the expansion area are in good agreement Mexp=2.15, Mthe=2.13. The photos in Figs. 31 a and b present holographic interferograms of flow around a sphere for M∞=0.8 (without shock wave) and M∞=1,06 (bow shock wave is in front of model).
Figure 31 b is combination of holographic interferograms (upper part) and Schlieren photo for the same flow. On the interferometric part of photo easily seen are: the stagnation point, the detached bow wave, the vortex sheet generated past sphere and so on.
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry
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Fig. 31 Holographic interferogram of flow around a sphere for (a) M∞=0.82 and (b) mixed hologram and Schlieren for M∞=1.06
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry Very interesting is example of flow visualization around tunnel wall perforations. Many transonic tunnels are operated with performed walls in the test section. A number of investigations have been performed to determine how the flow in the test section is affected by the presence of the perforation. perforation The next photos (Fig. (Fig 32) reports on test performed in T-36, with a single slanted slot in the bottom plate of the test section. The disturbances originating from the slot are expressed by distortions of the parallel fringe system. A concentration of fringes indicated the formation of a pressure wave. The slanted slot was used because it had been reported that such geometry would considerably reduce the perturbation of free flow. The interferogram however shows that the disturbance from the slot is not at all negligible and reaches even beyond the axis of the test section (to about 60 % of the test section height). The perturbation has the influence on the model sting in the central line of the test section (Fig. 32).
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry
Fig. 32 (a) Test section, (b) holographic interferograms of flow (flow is left to right) in the empty wind tunnel test section with wall perforations (slanted slot) and (c) and with cone for M∞=0.83
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry The interferograms of several supersonic rocket nozzle configurations (Fig. 33a) without and with different barriers are recorded in order to provide a good insight in the physical processes (Figs. 33 b and c). The theoretical value of Mach number in the output plane of the nozzle is estimated to be M=2.6. Using the data for pressure measurements, it is obtained M=2.46 and by means of holographic calculations, Mach number is M=2.56. The placing of barriers in the supersonic flow, leads to the appearance of the stagnation zone, shock and expansion waves. Visualization of the flow field made in the experiment indicates strong interaction of the turbulent boundary layer with the oblique shock wave in the divergent part of the nozzle. Beside two-exposition method, there are used the real time method, the average or sandwich methods,, the speckle p interferometry, y, refraction interferometry, y, differential interferometry and so on. Optical holography use laser light in visible spectrum, and interferential effects are recorded on photo or thermosensitive emulsions. Electronic holography uses CCD cameras. In some specific cases acoustic and microwave holography, with electron beam, X–rays, or computer holography can be used. With similar possibilities speckle interferometry, Moiré interferometry and stereography are used nowadays.
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Flow Visualization Optical Methods of Flow Visualization: Holographic Interferometry
Fig. 33 Holographic interferograms for 2D supersonic nozzle without and with three barriers
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Flow Visualization Optical Methods of Flow Visualization: Infrared Thermography Thermographic systems have been considered to analyze fluid-dynamic phenomena thirty years ago. Nowadays high resolution and differential infrared thermographic measurement systems open up new possibilities in its application. Temperature field that can be measured by a thermographic system on the surface of a solid body invested by a flow is determined by many combined effects. effects Very important effects are: conversion of kinetic energy of the flow into thermal energy, flow temperature variation in time and space, convection heat transfer phenomena between flow and body, conduction phenomena inside the body and radiation heath exchange of the body surface with surroundings. By correspondence between convective heat transfer coefficient and local turbulence, it is possible to carry out information about the boundary layer. In addition to the laminar-to-turbulent transition boundary, the infrared camera was able to detect shock waves and present a time dependent view of the flow field. Figure 34 shows thermograms of tests have been performed using an high resolution thermographic system for fluid-dynamics analysis of a known test case, case a wing profile, profile in a wind tunnel under variable and constant temperature condition at different air flow velocities. A time dependent heat transfer code was developed to predict temperature distributions on the test subject and any necessary surface treatment. A commercially available infrared camera can be adapted for airborne use. Readily available infrared technology has the capability to provide detailed visualization of various flow phenomena in subsonic to hypersonic flight regimes.
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Flow Visualization Optical Methods of Flow Visualization: Infrared Thermography
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Fig. 34 Black Aluminum airfoil with (a) incidence of 7.5O clockwise and (b) incidence of -7.5O clockwise
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