Appendix 3 - FURTHER INFORMATION FOR EXPERIMENT #7
Last modified, February 2005
This lab course primarily emphasizes the scientific foundations and basic
tools required to conduct and analyze engineering tests. This appendix
describes the procedures used in these tests. Engineering testing and simulation
plays an extremely important role in aerospace engineering. Any engineering
system must be tested before production is started. In aerospace engineering
the systems are frequently developed through testing. In many cases
this is called "simulation." Simulations use both computational and physical
tests to evaluate and refine all aspects of an aerospace system. Many engineers
are employed in both industry and government in this work, and high quality,
high productivity testing and simulation is crucial to national competitiveness
in aerospace engineering. This appendix provides an introduction to aerodynamic
testing procedures. These procedures should be followed when planning any
test, and should be used as a basis for conducting the wind tunnel tests
in this course. This material should also be useful after graduation.
1. Aerodynamic Testing
Wind tunnel testing was used by the Wright brothers to develop their first
gliders and the first controllable aircraft. As such, aerodynamic testing
has a long history of contributing to the development of flight vehicles.
During this time the procedures have become well defined. Nonetheless,
conducting a successful test requires significantly more sophistication
than might at first be expected. To obtain high quality data the test typically
requires close coordination between a number of groups. This section provides
a brief introduction to the procedures and steps associated with wind tunnel
testing. Aerodynamic testing also involves flight test, flight simulators,
and more unusual testing such as hover rigs for VTOL testing, and sometimes
water tunnel tests. In recent years many aero test engineers have become
test engineers in observables departments. Automobile companies are heavily
involved in wind tunnel testing, with about half the effort devoted to
external aerodynamics and half of the effort devoted to internal flow associated
with ensuring adequate air flow for cooling.
The book by Rae and Pope (Ref. 1) is an excellent introduction
to aerodynamic testing, and should be used as a primary reference. Aerodynamic
testing is not a pure science, and apprenticeship is the standard means
of learning the methods. Many of the best documents are developed primarily
for internal group training. Examples available from the author include
2-6. This material is equally important to students who will not be directly
involved in testing. Every aerospace engineer is at least indirectly involved
in testing. Testing, and the high cost of testing, plays a key role in
aerospace systems program schedules and development cost.
Traditional reasons to test range from physics oriented investigations
of basic fluid mechanics to measurements used in establishing design loads
for use in structural design of new, or modified, aircraft. Configuration
development is one of the key reasons to test. Despite the progress in
computational aerodynamics, aircraft configuration development is still
heavily dependent on wind tunnel testing. Development of the aerodynamics
math model used in creating flight control systems and evaluating aircraft
and flight control systems in manned simulation studies requires extensive
testing to define the aerodynamic characteristics throughout the flight
regime ( = -90° to
+ 90° at a minimum, over a large range of sideslip angles,
). Aerodynamic loads definition and math model development require a large
amount of data after the configuration is completely defined and
the configuration development job has been completed.
Testing is conducted to develop testing methodology itself, develop component
technologies, e.g. wing or inlet concepts among many, and aircraft
development and modification. An emerging area of testing is associated
with acquiring experimental data to validate computational fluid dynamics
Test engineers are required to be familiar with the available wind tunnel
facilities to determine what facility is best suited for a particular test
requirement. There are a limited number of test facilities available for
some of the more important types of testing. In addition to aircraft company
facilities, the major large wind tunnels required to develop advanced vehicle
systems are located at a relatively small number of locations in the US.
In particular, the number of large transonic wind tunnels is very small
and requests for test time must be made years in advance. Aero test engineers
travel frequently, often spending weeks and months at the test site. Major
wind tunnel facilities are located at:
1.1 Where you test, what you test
In many of these facilities the huge amount of electrical power required
to move the air around the test circuit dictates that testing be done at
night. At NASA Langley testing is reduced in the summer when the temperature
is high and power must be used to satisfy residential and commercial air
conditioning demands. When this happens the aero test engineer must literally
"cool his heels" until the temperature goes down. Subsonic testing can
be conducted at numerous locations in addition to those listed above. The
"standard" subsonic wind tunnel test section is 7ft by 10ft., and many
models are sized to fit that size wind tunnel.
NASA Langley, Hampton, Virginia
NASA Ames, Mountain View, California
NASA Lewis, Cleveland, Ohio
Calspan, Buffalo, New York
AEDC, Tullahoma, Tennessee
A remarkable variety of characteristics require testing. A partial list
(each of which is also evaluated in flight test) originating with Meyer
(Ref 6) is given here. This list is in addition to the
basic configuration development testing, which might also include detailed
flow diagnostics such as flow visualization. Each test objective will have
a unique set of difficulties that must be taken into account to obtain
high quality data for that specific item.
Thrust minus drag
Drag due to lift
Drag rise (transonic drag divergence)
Stability and Control
Stall characteristics (stability and control)
Configuration sensitivity (Gaps-Overlaps)
Basic longitudinal and lateral/directional derivatives, control effectiveness
Conventional flight regime
High angle of attack & stall data (static data) (famous Re effects
on pitching moment)
Dynamic data, .i.e. Spin Tunnel, free flight tunnel, and forced oscillation
Hinge moments (The first data mechanical controls guys want in design,
not well estimated even in the tunnel tests.)
Buffet onset, growth
Critical Flight Conditions (these are always a long way from design
Figure 1 provides a case history of testing required
to develop a modern aircraft, the F-16, and is based on data contained
in the excellent paper by Hancock (Ref. 7). That paper
also describes the important role that computers play in experimental aerodynamics.
The breakdown shows the relatively small amount of testing devoted to the
basic configuration concept, compared to the large amount required to define
the characteristics after the configuration is defined.
The F-16 was developed in the early 70's. Today, the high angle of attack
characteristics would be more fully evaluated, with additional testing
compared to the F-16. This is in part due to the experiences with the F-16.
The angle of attack range where large regions of separated flow occur,
yet prior to the angle at which the flow is entirely separated (25°
< < 50°
for a typical modern fighter), is a critical flight region. In this region
wind tunnel results are found to be sensitive to Reynolds number, support
interference and possibly other effects. Results from different wind tunnels
often produce different pitching moment characteristics. This is one area
still requiring engineering judgment to interpret the results obtained
in the wind tunnel.
One area requiring extensive wind tunnel testing, but not covered in this
course, is propulsion system testing. The airframe manufacturer is responsible
for integrating the inlet into the configuration such that the engine is
supplied with a uniform flow of air at the compressor face with minimal
pressure losses. This is required for all flight conditions, and frequently
leads to variable geometry in the inlet/diffuser system. The nozzle must
also be developed. Traditionally, the engine manufacturer designs the nozzle.
However airframe-propulsion integration requires that the components be
tested extensively together.
Thrust minus drag
Inlet stability - bleed requirements
Acoustic, thermal environment
Most new engineers are involved in testing on large programs. More experienced
engineers get involved in tests of new concepts and other specialized investigations.
Because of the large amount of testing associated with aerospace design,
the following material adapted from McLaughlin (Ref. 2)
is included to help students gain an overview of the types of testing that
occurs at various stages of aircraft system development.
The types of aerodynamic testing conducted during aircraft development
1.2 Testing related to aircraft
The scale of the test activity, including planning, model fabrication,
pretest, test and post test analysis is much more extensive than engineers
unfamiliar with the process expect. Consider also that a similar effort
is required for structures, flight control systems, propulsion systems,
and every other component. This testing is required to ensure that the
entire system performs safely and with the required performance. Unfortunately,
the extensive testing required in modern aerospace programs is one of the
key reasons for high development costs and long program development times.
|Conceptual Study Phase
develop key technologies
screen configuration possibilities
screen weapons carriage possibilities
screen propulsion system candidates
aerodynamic wind tunnel test work
component technology development
high lift system development
control surface concept investigation
powered lift concepts
high angle of attack alternatives
auxiliary control surfaces
determine parametric effects of configuration integration
component flutter tests
isolated inlet tests
|Concept Definition Phase
select propulsion system
optimize & refine configuration
define cost & schedule to develop
aerodynamic wind tunnel test work
gross refinement of configuration architecture
proof of concept
weapons carriage parametrics
|Preliminary Design (alternative systems exploration)
investigate alternate approaches
perform cost effectiveness studies
refine configuration design
identify critical design issues
aerodynamic wind tunnel test work
detailed refinement of configuration architecture
stability and control definition
support simulation activity
study spin & spin recovery characteristics
preliminary loads definition
preliminary weapons carriage definition
|Demonstration/validation phase (pre-full scale development
resolve critical design issues (technology verification)
refine configuration design
start detailed design (long lead items)
solidify data base
aerodynamic wind tunnel test work
level II aerodynamics loads definition
update aerodynamic data base
stability and control
weapons carriage definition
integrated inlet testing
|Engineering and Manufacturing Development (EMD), Formerly
known as Full Scale Development (FSD)
complete detailed design
perform test & evaluation (ground and flight tests)
begin pilot production
aerodynamic wind tunnel test work
periodically update aerodynamic data base in support of program milestones
determine dynamic derivatives
determine spin and recovery characteristics
acquire weapons carriage & separation analysis data
document performance & flying qualities of flight test article
investigate safety of flight issues to support flight readiness reviews
full configuration dynamic flutter tests
Aerodynamic Test Planning: the objective and the scope of the test must
be precisely determined before beginning to spend money on the test. Planning
is important, a typical aerodynamic test program requires 6 mo. to 1 year
and the cost can range from several hundred thousand to several million
dollars. A complete full scale development (now EMD) test program will
cover a six year period. The basic information required to determine the
scope of the test and obtain the initial cost estimate is:
1.3 Aero test procedures (based
on Ref. 2)
Close coordination with all the participants is mandatory. With wind tunnel
testing being so expensive, a well defined test plan with precise objectives
has the best chance of resisting budget cuts. A major problem is justifying
the need for as many tests as experience indicates are necessary during
Engineering and Manufacturing Development (EMD). To solve this problem
each test should be associated with a program milestone or anticipated
Division of Responsibility
Normally the project requiring the test does not run the test directly.
Within the organization a test group usually runs the test. When the test
is going to be at another facility, you will also be interacting closely
with their engineering staff. The following lists illustrate the typical
division of work, and is based on the procedures used by a major airframe
test facility choice
date of entry
Aero Test Engineer Responsibility
establish test requirement with project agreement
establish funding & funding limitations (program authorization)
designate a single point of contact for test coordination
provide configuration details
provide model aerodynamic loads information
prioritize test objectives
define required test conditions
formulate run schedule overview
approve most efficient run schedule proposed by aero test
document pretest estimates of test results
observe direction of the test, work with aero test engineer
evaluate test results during the test
analyze and document test results
Note that both the aerodynamicist representing the project and the aero
test engineer conducting the test must work closely together. With differing
responsibilities, opportunities for conflicts to arise between the aerodynamicist
and test engineer frequently present themselves.
Successful test programs begin with a well designed model. The model should
allow for easy change of parts, and provide for a means of precisely aligning
the model in the tunnel. The model should be easy to work with in the tunnel.
Attention to detail during the design will payoff during the test. Most
subsonic models are made of wood built on an aluminum frame. Most high
speed models are metal. If the model is large enough, aluminum can be used.
If the model is small, steel may be required to obtain the necessary strength
while maintaining the design contour. Model design and fabrication is a
specialized field, and in addition to the major airframe manufacturers
and government laboratories several smaller companies specialize in making
wind tunnel models. These include Micro Craft, based in Tullahoma, Tennessee,
and DEI (Dynamic Engineering, Inc.) based in Newport News, Virginia. Because
of the specialized nature of the work, these companies can sometimes make
a model faster and more cheaply than in-house machine shops.
The most important consideration during fabrication is geometric accuracy.
Unexpected errors frequently arise during fabrication. Before the days
when surface shaded computer graphics were available the first view of
the model occurred when the numerically controlled (NC) milling machine
tape checkout piece was made. The aerodynamicist would frequently have
an unpleasant surprise when examining this piece. Complex three-dimensional
surfaces are difficult to visualize, and graphics should be used to verify
the shape before releasing the lines to the NC tape programmers. Once the
model is finished, it should be carefully inspected using a coordinate
measuring machine over nearly flat surfaces, and templates cast normal
to the leading edge and blown up 20 times on a comparator machine to document
the actual model geometry. Model tolerances are usually specified in both
absolute values and relative accuracy in terms of waviness. Tolerances
should be tight around the leading edge, and are usually relaxed after
the first 5-10% of the chord. It is unrealistic to specify a tolerance
of less than .001 inches on a complex surface. Even this accuracy will
require a highly skilled craftsman. Accuracies quoted by machine shops
are frequently based on combinations of flat and exactly round surfaces.
Accurately machined wing contours are more difficult.
The Run Schedule
The run schedule is a key item in structuring the test. It is used to establish
the entry time requirement, and sequence of model changes. A run schedule
is a list of the "runs" to made during the test. A "run" is usually a fixed
configuration with data taken over a range of angles of attack (or side
slip). Each angle of attack where data is taken is called a "point." A
schedule is initially created as a "wish list," or straw man, and then,
after the initial time and cost are estimated, scrutinized again by the
facility and the originator. Wind tunnel time is frequently made available
for a specific calendar time, based on the initial run schedule. Typically,
the aerodynamicist doesn't get enough time to perform the complete set
of runs identified on the run schedule. The run schedule is negotiated
with the facility operators to define the tunnel occupancy time, and should
be "massaged" by test engineers and the facility engineers to produce the
most efficient use of facility capabilities within the time available.
Recall that wind tunnel resources are limited, and there are great demands
on national facilities from a variety of programs. The schedule is reworked
to reflect changes in priorities as the test approaches.
Once the test begins, the schedule will be altered based on the analysis
of initial results obtained, and schedule problems. Model installation
almost never goes as smoothly as expected, and normally no additional tunnel
occupancy time is made available to make up for the lost time. In addition,
initial results often mandate special unplanned runs to examine unexpected
results. The ability to maintain a productive test program which maximizes
the amount of data while reacting to surprises (problems) discovered in
the data is a challenge that requires both concentration and talent. Engineers
that can do this well are frequently called "tunnel rats." One well known
tunnel rat is Irv Waaland, who was the project aerodynamicist on the Gulfstream
II and F-14 programs at Grumman, and worked on the F-18 and B-2 programs
at Northrop. Waaland won the AIAA Aircraft Design award in 1989.
The following list provides some run schedule rules from another "tunnel
rat," Charlie McLaughlin:
designate a single point of contact for test coordination
provide cost estimates & monitor expenditures
establish model design, fabrication, and test schedules and monitor progress.
select appropriate test facility
interface with test facility
propose most efficient run schedule to meet test objectives
provide pretest report to aerodynamics/facility/test sponsor
monitor model fabrication, instrumentation installation and calibration
conduct or supervise test at maximum efficiency (minimum cost)
evaluate test data for measurement uncertainty and accuracy
supervise data reduction and transmittal to aerodynamics
provide post test report
Once the model is removed from the tunnel it may never go back in. An engineer's
worst nightmare is to be assembling the test results for presentation to
management and discover that a key run needed to make an engineering decision
was not made. This could be a multi-million dollar mistake. The engineer
running the test must make sure this doesn't happen. The best way to avoid
this is to work with the data on site as it's available and, in effect,
create your presentation to management during the test. You should be able
to get off the airplane from the test site and go directly into the vice-president's
office with your results in essentially final form.
Examples of several run schedules are shown in the figures. Figure
2 is from a Calspan test of the Grumman forward swept wing configuration
before it became the X-29. This is the actual test run log. The primary
information is the configuration description, a code describing the configuration,
the control settings, the angle of attack and sideslip schedules (as noted
above: a "run" usually consists of a single
or sweep, with data
being acquired at discrete "points") the Reynolds number, Mach number,
and very importantly, a remarks column. The remarks column helps explain
the purpose of the test, and any special problems. The remarks column does
not replace the test notebook kept by the test engineer. The remarks column
is the place to identify changes in the transition strip (these will be
discussed below). In this case several Mach numbers were run for each configuration,
and the run number is shown under the Mach number heading. The nomenclature
requires additional explanation.
Figure 3 shows a typical
nomenclature sheet. Pictures of each nomenclature item should also be taken.
Figure 4 shows a portion of a run schedule from a test
at AEDC. In this case the run numbers are listed on the left hand side.
This is more typical of subsonic run schedules similar to the schedule
required for testing in the aerodynamics lab class. All the "codes" refer
to the description of the configuration for use in the computer when performing
Special considerations in developing run schedules are required to obtain
the reference conditions for use in data reduction. This includes initial
runs to document the weight of the model in a particular configuration
as it moves on the support with no wind load. These are called weight tare
runs. This must be done for each different configuration which affects
the distribution of weight. The forces and moments obtained during these
runs are then subtracted from the wind-on results to find the aerodynamic
forces on the model. Another set of early runs is frequently performed
to subtract the internal flow drag, if a flow through inlet is included
in the model. This drag is found using a pressure tube rake positioned
at the exit of the nozzle. The internal drag of the duct is not part of
the external aerodynamic force, and is properly included in an adjustment
to the propulsion force.
In most cases runs should be made early to determine flow angularity effects,
and possible support interference effects. This would include testing the
model in the upright and inverted position and through use of a dummy support
system. Details are given in the reference by Rae and Pope (Ref.
Defining data requirements
Accuracy requirements must be established early in the test planning process.
They have a major impact on model design. The instrumentation and method
of model fabrication will be determined by the accuracy specified. Over-specifying
the accuracy drives the cost up, while poor accuracy could lead to results
that are not precise enough to use, thus wasting the entire effort. Engineering
judgment must be used to determine the necessary accuracy.
Always begin with the simplest configuration (this is particularly true
of a new model).
Begin with a repeat configuration if the model or design was tested previously.
Model parts availability & variability must be consistent with the
Make one model change at a time during configuration build up or modification
Make sure the resulting data comparisons are complete.
Assign a priority to each run.
When in doubt - make the run!
Flow visualization techniques
Frequently the aerodynamicist needs to understand the flowfield in more
detail than is apparent from overall forces, surface pressure distributions,
and local balances such as root bending moment gages. Flow visualization
is perhaps the most useful tool that an aerodynamicist has available to
understand the details of the flow. The term "flow diagnostic" is usually
employed to describe these types of tests. This testing is slow and currently
produces primarily qualitative results. Therefore, in aircraft projects
it's not done unless other results indicate that a problem exists requiring
this information. Flow visualization is heavily used in research oriented
Flow visualization is generally divided into two types: surface and off-surface.
For traditional, primarily attached flow aerodynamics, surface flow visualization
is generally adequate. The types of information of interest include the
transition location, separation location and shock position. Tufts are
often used, and provide gross characteristics of the flow direction, separation
lines, and indications of flow unsteadiness. Black silk thread attached
with flat Scotch Magic tape, with the model painted light gray, works well.
Yarn is also used frequently, and should be about two 1 1/2 to 2 inches
long. They should be taped to the model in spanwise rows, spaced far enough
apart that they don't interfere with each other. Four or five rows would
work well on a six inch chord model. Figure 5 shows a
typical layout of tufts on thre upper surface of a wing. The layout is
shown schematically in Fig. 5a, and a photo of a tufted
model is presented in Fig. 5b. Small silk tufts may
oscillate too fast to see with the naked eye, and high speed photography
may be required. Minitufts (Ref 1) also work well under
the right conditions. Current VHS type video camera results often lack
the resolution provided by traditional photography, and may be marginally
valuable in making a record of the flowfield using tufts. Surface oil flows
of one sort or another are useful, but extremely messy. Applying the oil,
making the run, and cleaning up is a slow process. This makes oil flow
testing thus costly (it's not uncommon to being paying a dozen engineers
and technicians on each shift. The solvents that work best are probably
no longer acceptable in most workplaces (trichloroethylene). Colored artist's
oils thinned slightly on a white surface, applied as dots, works well.
Alternating the colors can help trace the origins of the local flow. Unfortunately,
some of the most spectacular oil flow studies I have seen occurred on competitive
configuration development jobs, and aren't available for illustration of
Figure 6 provides a sample of an oil flow photograph
taken with a Polaroid camera at the NASA Langley 4-foot Unitary Plan wind
tunnel at M = 1.62. In this case the oil shows a crossflow shock wave outboard
on the wing, and a rare "open separation" associated with flow moving behind
the shock from inboard upstream, creating a zone that the flow immediately
upstream of the shock cannot penetrate.
One of the most remarkable aspects of flow visualization is how quickly
the observer may forget the details. It is extremely important to make
the most detailed possible sketches and notes at the time of the test.
It is easy to loose confidence when your boss is arguing that you "couldn't
have seen" what you say you saw. Surface oil flows are relatively easy
to document. Tuft probe work is not. No flow visualization recording technique
is equivalent to observing the flow visualization directly with your own
Off surface flow visualization is valuable when trying to understand flows
with organized vortices and wakes. Surface results only provide a limited,
and usually confusing, glimpse of the physics of the flowfield in these
cases. Today the laser light sheet is one of the easiest ways to visualize
the flow. Prior to the laser light sheet, tuft grids were used to gain
The typical instrumentation compliment on an aerodynamic test would consist
6 component strain gage main balance
surface pressure taps
data correction measurements (determined by test engineer)
model attitude measurements
surface deflection angle measurement
Additional instrumentation necessary for a component design/integration
experiment (wing, tail, canard, store installation, propulsion system integration,
auxiliary strain gage balances
concentration of surface pressure taps
flow survey rakes
dynamic instrumentation to sense separation/buffet
When running a dedicated loads test many strain gages and pressure taps
Additional instrumentation is required for a wing design investigation
(at a minimum):
6 wing spanwise pressure tap stations (up to 150 pressure taps)
wing root 3 component strain gage balance
normal force, bending moment, torsion
wing tip accelerometer to determine buffet onset determination
flow visualization: oil flows, surface tufts, sublimation
Other flow evaluation methods may be required. These include:
hot wire anemometers, laser doppler velocimeter
hot film gages (surface measurements), dynamic surface pressure measurements
survey rakes, vortex probe, boundary layer survey probes
flow direction probes
The following list summarizes the available techniques:
Details of these methods can be found in Rae and Pope (Ref
Fixing boundary layer transition
This is one of the standard issues that must be resolved. The first question
an aerodynamicist unknown to a facility is asked when he arrives at the
tunnel is "what to do about transition?" This is a "rites of initiation"
question that you had better be prepared to answer. Normally, after a lot
of discussion the facility engineers end up using "the usual procedure"
for that tunnel. The location of transition from laminar to turbulent flow
is controlled by putting surface roughness on the model.
Why do you fix transition?
boundary layer studies
oil flow - transition, separation, shock location, flow direction
surface tufts - separation, flow direction
sublimation - transition
laser light sheet
boundary layer separation
so that the transition position is known
so that the flight boundary layer characteristics are better simulated
at critical conditions
at various locations on all model components
There are basically two different considerations, leading to two different
methods in defining transition. The simplest is the requirement that the
model have a fully turbulent boundary layer, or at least a known state.
This allows the aerodynamicist to make an adjustment to zero lift drag
to estimate flight values from wind tunnel data with some confidence (see
2 for further discussion). The second type of boundary layer transition
is associated with proper modeling of the boundary layer at shocks and
at the trailing edge. This is associated with modeling the boundary layer
near separation, and is mainly associated with drag due to lift. This is
usually the critical consideration at transonic speeds. The first method
is usually used for subsonic and supersonic testing, while the second method
is associated with transonic testing. An extremely precise procedure might
reposition the transition location for each angle of attack. Clearly this
is impractical for most testing, and engineering judgment must be used
to decide the best approach to simulating full scale conditions.
Fully subsonic and supersonic test conditions
aerodynamics, aero test and the test facility - jointly
Mixed flow (transonic conditions)
Braslow method (NASA TN-D-3579, Ref. 7)
fix transition near the leading edge
transition strip height and streamwise location a function of test Reynolds
number and Mach number
method is generally used at all test conditions to fix transition on all
other model components
at subsonic speeds trip strip height is generally within the boundary layer
so that no drag correction for the strip is necessary
at supersonic speeds it may be necessary to determine a grit drag correction
experimentally by testing various strip heights
Special case - aft cambered wing lower surface
Blackwell method (NASA TN-D-5003, Ref. 8)
fix at a further aft wing location to better simulate the flight boundary
layer at the wing shock location and/or the wing trailing edge
transition strip height and location a function of test Reynolds number,
Mach number and shock location
computational viscous airfoil analysis used to determine location
Braslow method used to determine height
percent chord location can vary spanwise due to wing taper
trip must be located at least 5% chord ahead of the shock to avoid laminar
shock boundary layer interaction
boundary layer must be kept laminar to the trip by keeping model surface
The exact details of the procedures are beyond the scope of these notes.
The reader should consult Rae and Pope (Ref 1). However,
the essential idea is that a roughness element is selected to give a roughness
element length based Reynolds number of 600 at a streamwise location corresponding
to a running length Reynolds number of 1.x106. This assumes
that the flow at this point has not undergone premature transition due
to adverse pressure gradients. This presents a problem for subsonic flow
at high lift, where suction peaks are immediately followed by a rapid recompression
very near the leading edge.
In many standard aerodynamic configuration tests at subsonic speeds, black
electrical tape cut by pinking shears is used. The tape is positioned about
5-10% aft of the leading edge. The pinked edge is placed facing into the
flow. Sometimes, several layers are used.
The importance of pretest estimates cannot be overstated. Without pretest
estimates it is impossible to know if you have a potential problem, or
to be able to assess whether the objectives of the test are being achieved.
The aerodynamicist must be able to determine at the test if the data fails
to agree with pretest estimates because of a data reduction error, or if
an actual aerodynamic problem exists on the model. This is similar to debugging
a computer program. In either case immediate action must be taken, and
adjustments to the run schedule may be required. Pretest estimates are
frequently required to use government facilities. The government sponsor
will attend the test and plot your experimental drag polar and predicted
pressure distributions on the same sheet (or screen) with your pretest
estimates. At this point your company will either establish credibility
with the sponsor or begin a corrective action program to establish credibility
(this could mean your job).
To summarize - pretest estimates are crucial, they:
fix further aft to provide a better flight boundary layer simulation in
the region of adverse pressure gradient caused by aft camber
avoids cove separation that could occur at lower test Reynolds number
One classic example of test work based on pretest estimate discrepancies
is the story of how the Grumman concept for the forward swept wing concept
arose. The forward swept wing idea resulted from the transonic test of
an aft swept wing of a Grumman configuration which had been proposed to
satisfy the NASA HiMAT RFP (request for proposal). The winning proposal
was submitted by Rockwell, and the Rockwell HiMAT was built, flown, and
is currently in the Smithsonian Air and Space Museum.
After Rockwell was selected, NASA still expressed interest in the Grumman
wing design, and offered to test it. Initial results indicated that the
pretest drag estimate was 20% optimistic compared to test results at the
maneuver design point (MDP). The cause of the problem was traced to the
empirical drag estimating methodology. It was too approximate for the class
of advanced airfoil technology used in the wing. On-the-spot rethinking
of the estimation method, using some additional runs at the test site,
resulted in an improved method that resolved the discrepancy. Different
wing sweep angles were studied. This was accomplished easily because the
model had a variable sweep wing (other aerodynamicists have yawed models
to study slight changes in sweep). The new drag estimation method (and
understanding) provided the first clue that a forward swept wing might
be optimum for the HiMat maneuvering design goal: 8 gs sustained at .90
Mach number and 30K feet altitude. Without precise expectations based on
pretest estimates the forward swept wing idea would not have emerged. Glenn
Spacht was the aerodynamicist who did the work in the tunnel at NASA Langley.
This effort resulted in his advocacy of forward swept wing aircraft and
his career advanced quickly. He was deputy project manager of the X-29
program and is now director of engineering at Grumman.
In addition to precise estimates, more approximate estimates are required
to locate and size the balance and determine ranges for pressure gages.
Structural analysis will be based on these estimates. Model safety analysis
will depend on them. A recent development is that some facilities (AEDC)
will request the model geometry in a numerical format which allows them
to make their own computational analysis to determine if the objectives
of the proposed test can be achieved using the proposed approach.
Many issues must be considered in conducting the test. Items that must
be checked when preparing to run the test are discussed in this section.
Tunnel flow calibration issues: The flow in a wind tunnel is not
an exact simulation of flight in free air. The tunnel flowfield will vary
in velocity and direction throughout the section. Some freestream turbulence
will be present in the wind tunnel flow. Any tunnel will have a flowfield
survey describing the quality of the basic flow. The quality of the flow
will vary between wind tunnels. The Virginia Tech Stability Wind Tunnel
has extremely good characteristics (see section 5). Sometimes
the facility operators will not be particularly eager to show users the
details of the flowfield, but will specify where the model should be positioned
in the test section to obtain the best results.
Force balance calibration: One of the most important considerations
in testing is the accuracy of the balance. The balance should be check
loaded every time a model is installed. These results should be available
for use in the data reduction program.
Model Support Interference: The model support system will produce
a deviation from the flowfield that the actual airplane would encounter
in flight. This interference is minimized by wind tunnel, model design,
but is never eliminated, and must be considered during data reduction.
To assess the magnitude of these effects several experimental techniques
are used. Rae and Pope (Ref. 1) describe methods for
correcting the test results to attempt to eliminate these effects from
the data. This is a problem area where computational aerodynamics can be
used to get insight into the way in which the support is interfering with
Fouling and removal of base/cavity drag: Several rather subtle aspects
of testing can be troublesome. The first is known as fouling, and occurs
when the support system hits the portion of the model supported by the
balance. This arises because of the desire to make the smallest possible
deviation of the model to accommodate the balance and the deformation of
the model under load. This can be checked by installing a fouling circuit
strip, such that if the model touches the support an electric circuit is
completed and a warning is issued at the operators console. Temptation
to save installation time and run without a fouling strip can lead to acquisition
of questionable data and is a false economy. Frequently it is hard to determine
if fouling exists without the fouling strip.
Another important aspect of aerodynamic testing is the correction required
to account for base or cavity drag that does not exist on the actual vehicle.
This requires the measurement of a base or cavity pressure, and the determination
of the area over which this pressure acts. Cavity pressures should be checked
early during testing to make sure that the correction is being made properly.
Axis issues: Care must be taken to ensure that the axis system is
understood. If the balance is mounted on a strut the measurements will
be made in a wind axis system nominally aligned with the tunnel. If the
balance is located internally in the model the forces will be measured
in a body fixed coordinate system. The usual axis systems are the wind,
stability and body axis systems. Most aerodynamic performance analysis
requires results in the stability axis (lift and drag). Stability and control
work may be done using either axis. High angle of attack stability and
control analysis is usually carried out in the body axis. It is generally
easier to understand the effect of a conventionally oriented control device
using the body axis because they are generally designed to produce a pure
moment about a specific body oriented axis. Airframe manufactures may use
either system, and this is sometimes a source of confusion when companies
engage in joint ventures. Figure 7 from Ref.
5 illustrates the sign convention and relationship between axis systems.
Transformations from Rae and Pope (Ref. 1) are repeated
here. Note that the data should be translated to a common origin, and then
rotated using the angle of attack, ,
and angle of yaw, :
For data analysis the sideslip angle
is usually taken to be -.
The subscript "sa" denotes stability, "w" denotes wind, and
"B" denotes body based axis systems. In the body axis system lift
and drag are replaced by normal and axial force, and are used without subscripts.
Here c is the mean aerodynamic chord and b is the reference
Stability axis results are obtained from the wind axis results through:
Body axis results are obtained from wind axis results through:
CN = CLwcos
CA = CDwcoscos
The wind axis coefficients can be obtained from the body axis coefficients
CLw = CNcos
CDw = CAcoscos
CYw = CAcossin
provide an immediate indication of test procedure and/or data reduction
provide an immediate indication of model design and/or fabrication problems
provide an immediate indication of aerodynamic design problems
estimates generally assume an optimum design
the problem can be either due to the actual aerodynamic configuration or
with the estimating methodology
Numerous corrections to wind tunnel data may be required to predict the
corresponding values that would occur under actual flight conditions.
8 illustrates a few of the issues that must be considered. Several
chapters of Rae and Pope (Ref 1) contain details. One
fundamental data adjustment not shown is the weight tare, which is the
balance reading with the model moved through the test range with the wind
off. This must be subtracted from the wind off force to get the aerodynamic
force. Usually a table is created and interpolated for the combinations
of angle of attack and yaw used in the actual test. A table is required
for each configuration with a different mass distribution.
Flow angularity adjustments are required because the freestream velocity
is not precisely aligned with the tunnel centerline. As shown in Fig.
8, when the balance is mounted on the strut, the measured readings
must by transformed, and the angle of attack corrected using the following
CLtrue = cosupCLmeas
CDtrue = sinupCLmeas
Ideally the flow angularity, up,
is small. When this is the case, these equations show that the correction
to the value of lift is small, but the correction to drag can be large.
This is one of several reasons that drag is difficult to measure in a wind
tunnel. The flow angularity is best found from a careful test using a reference
wing in both the upright and inverted positions, with a dummy strut system.
Alternately, the tunnel survey with no model present can be used. Sample
survey results for the Virginia Tech Stability Tunnel are contained in
Another correction is required to account for test boundary effects. These
are different depending on the type of boundary; solid, open, or slotted.
As long as the model is small compared to the test section the corrections
should be small. Wind tunnel wall correction theory is not yet complete,
and a variety of advanced wind tunnel concepts have been proposed to minimize
Post Test Analysis
Final data reduction/analysis: As described above, the data reduction
program should be verified, and the initial data plotted together with
the pretest estimates before leaving the test. In most facilities the only
difference between the data available on site during the test and the final
data is due to the use of more elaborate averaging process and the correction
of problems identified during the test while comparing with the pretest
analysis. The elaborate averaging process usually doesn't change the results
by an amount that can be plotted on a normal report scale. Data reduction
system errors are frequent, and will likely never be completely understood
unless identified during the test. Without pretest predictions it is virtually
impossible to spot these problems. Examples include the wrong sign on a
tare correction (responsible for a major problem once when a full scale
V/STOL model, including two operating TF-34 engines, was being tested in
the NASA Ames 40x80 tunnel); incorrect removal of base drag at the sting
cavity, and failure to take into account misalignment between the model
axis and the sting axis. Sometimes a reference condition is incorrectly
entered, and the balance is assumed to be broken when in fact the data
is good. I have encountered all these problems at one time or another.
Documentation/Final report preparation: The technical requirements
are addressed in section 3, Presentation of Aerodynamic
Data. In determining the final report contents, the test engineer will
consult with the project and determine:
Data correction and reduction
Other analysis included in the post test report include plots providing
Reynolds number effects, and transition strip effects, both transition
strip on and off, but also strip size and location effects.
Wind tunnel to flight data corrections/extrapolation: This effort
is strictly speaking not part of the aerodynamic testing problem, but is
closely associated to the application of wind tunnel data to actual aircraft
predictions, and has a direct bearing on the planning and conduct of the
test. It is discussed in some detail section 2.
types of curves desired (plotting schedules)
selection of scales
type of curve fits to be used in connecting data points
comparison plots to be made
Safety is a key consideration in planning and conducting an aerodynamic
test program. People die during test operations. Not often, but I'm personally
aware of several cases. In one case the model technician was working under
a large powered Grumman model lent to NASA Langley for their use. The sting
support failed, and the technician died after being crushed by the model.
For this reason, safety must be a key consideration during model design.
A stress analysis, documented in a stress report, must be done for the
model, and the model-balance-support system under all the loading conditions.
A stress analysis must also be done for all appendages to the model. The
facility will require these reports before model installation and testing
Model scale wings and especially canards and tails are often extremely
sharp, and engineers and technicians working around models are frequently
cut. In some tunnels the model is hoisted above the tunnel and lowered
into the test section using a ceiling crane. In one case I was, luckily,
working the other shift at McDonnell when the workers hoisted the model
too fast. It started swinging back and forth. One of the technicians tried
to keep it from hitting the wall by jumping in the way to stop it. His
whole shoulder was crushed.
Occasionally an engineer stands inside the tunnel test section during tests
to probe the flow with a tuft wand. Extreme care should be taken when this
is done. Safety goggles must be worn. Finally, a video camera should be
used to monitor the engineer (unlike the Virginia Tech Stability
Tunnel, in many tunnels the operator cannot physically observe the
test section), and a hand should be kept on the emergency stop button.
Two other test procedures require special safety considerations. High pressure
air is always potentially dangerous, and care must be taken working with
it. The final item is associated with the introduction of lasers into aerodynamic
testing on an almost routine basis. In particular, laser light sheet flow
visualization is extremely valuable. However, in examining the flowfields
in laboratory conditions it is sometimes very easy to be tempted to look
directly into the sheet from the side. Special care should always be taken
to avoid eye damage. Most facilities have established strict procedures
for working with laser systems.
Engineers are required to consider safety, and must be prepared for investigations
resulting from tests they plan, design or conduct, when an accident occurs.
NASA requires a stress analysis and an aeroelastic analysis as part of
the pretest safety review before testing in their tunnels. They are concerned
for both human safety and possible damage to the tunnel if the model fails.
One joint Grumman/NASA Langley program resulted in a model/support system
combination that was dynamically unstable at high angle of attack. During
one test in the Langley Full Scale Tunnel the dynamics resulted in a sting
system failure. The oscillation built up so fast that on the video of the
test (always a good idea) you couldn't see the failure. One instant the
model was in the picture, and the next instant it wasn't. Several incidents
have occurred where a key major national facility was closed for a year
or more after a failure during testing. Due to the control on model safety,
in most cases the tunnel structure and not the model caused the accident.
Flight operations are even more rigorous regarding safety.
2. Use of Wind Tunnel Data in Aerodynamics
(based in part on Ref. 6)
Once the test is conducted, adjustments are made to the results to apply
them to the full scale vehicle at flight conditions. This is a step where
judgment is required, and it is not unusual for the aerodynamicist to err
at this stage. The consequence of a mistake here can be very serious, and
aerodynamics managers frequently demand details and caution here.
A somewhat standard list of corrections (assuming that the usual corrections
to account for wind tunnel wall effects, etc. have already been made) to
wind tunnel data include:
Frequently theoretical and computational aerodynamics methods are used
to estimate these effects. Otherwise, adjustments are made based on previous
results obtained experimentally. Since each aircraft company has a different
experience base, the adjustments will differ slightly among companies.
For example, consider the problem of extrapolating the minimum drag obtained
in a wind tunnel test to flight conditions. If the geometry is the same
for your model and the full scale vehicle (which it usually isn't, as discussed
below), then the primary consideration is the skin friction. Typically,
a theoretical estimate is made of the skin friction at both the model and
flight conditions. The difference between the theoretically estimated drag
computed at the two different conditions is then added to the experimental
results to estimate the minimum drag of the vehicle at full scale conditions.
This is one of the reasons that it is critical to know how much laminar
flow exists on the model. Otherwise, the model friction estimate may be
Model scale effect:
Skin friction (Reynolds number and transition location): This is frequently
the most significant correction made to the data.
CLMAX - this adjustment is crucial, and is difficult.
Support system interference and modifications made to the model to accommodate
the sting installation.
Airplane protuberances and surface roughness not simulated on the model.
Geometry and surface finish (includes all cooling inlets and vents, all
Propulsion/internal flow related: Airplane secondary air systems not simulated,
nozzle geometry not accurate (often arising from item 1), incompatible
or inadequate inlet spillage drag conditions.
Explicit power effects
Aeroelastic effects - more difficult with respect to aeroelastically tailored
Even with advances in computational and experimental techniques, several
aspects of wind tunnel to flight adjustments still require further research.
The following list presents some of the more notorious:
Recognized Wind Tunnel Problem Areas
There's no simple way to extrapolate from wind tunnel to flight operation
of aircraft with total confidence. Reference 10 provides
a basic primer on the subject. To get some insight into the wealth of complications
and problems that arise, look at the important NASA work on the XB-70 (Ref.
Reynolds number sensitive data
Supercritical airfoil and winglet test requirements
"Scaling" buffet intensity data
Sensitivity to model lines/accuracy
Controlled transition/grit selection
Consistency of high alpha data
Tunnel wall interference
Reflection plane models for drag
In practice the two items that are very difficult to deal with
Another complication is the reliability of initial flight test results.
The issue of flight test data accuracy further complicates the comparison
between wind tunnel and flight test data. When the flight test data first
arrives and the inevitable panic begins, experience shows that the flight
test data system itself may be suspect, although the flight test group
usually refuses to acknowledge the possibility. Two particular cases that
I had some involvement with were the discrepancy between tunnel based estimates
and flight test results for the loads on the vertical stabilizer of the
EF-111 (upper transonic, M = 1.1, windup turns), and cruise drag of the
Grumman-American Gulfstream III executive jet.
In the first case the initial flight data was over predicting the discrepancy
by about 100%, and about 50% was real. A redesign of the vertical tail
structure of the EF-111 was required (recall that the EF-111 differs from
the F-111 by the addition of a large electronic jamming antenna placed
on the vertical tail-the EA-6B has a similar pod but never flew in this
speed regime). During the ensuing "fire drill" the Air Force program office
was demanding daily viewgraph presentations of the results of the Grumman
In the second case the discrepancies were apparently traced to the use
of a reflection plane model and two separate problems. Reflection plane
models are notorious for poor drag results. The G-II reflection plane model
tunnel to flight increments (which were found to agree nicely with flight
data) were applied to the G-III reflection plane results. The first problem
was that the G-III wing was added to the basic G-II reflection plane model,
resulting in a wingspan that was too large for the tunnel. The second problem
was that tunnel to flight increments for "conventional" (actually the British
developed peaky) airfoils and the supercritical type airfoil used, at least
in part, on the G-III scale differently between tunnel and flight (this
involves the proper way to fix transition).
Cases similar to the ones described above usually aren't documented in
the literature; they're not stories the participants are particularly proud
of. In the end, when working with the data you begin to get a feel that
you understand the problem and can account for everything. i.e. that you
have "the story" (but you can't prove it scientifically).
Another famous (possibly the most famous) case involved the tunnel to flight
pressure distribution on the C-141 wing. The shock in flight was much stronger
and further aft than had been found in the wind tunnel. This led to trim
changes and additional drag. At the time the difference was attributed
to Reynolds number effects, and was used to convince congress that the
a new wind tunnel was needed which could simulate flight Reynolds number
using wind tunnel models (this requires that testing be carried out at
very low temperatures using near liquid state nitrogen). This wind tunnel
has now been built (after having been designed in large part by a VPI grad,
Blair Gloss), is located at NASA Langley, and is known as the National
Transonic Facility (NTF). Subsequently we tried to use this case to demonstrate
the capability of computational methods to simulate Reynolds number effects.
The results were disappointing, with only about half the shock movement
predicted by our computational methods. When discussing our results with
engineers familiar with the C-141 data, we were told that the Reynolds
number was only part of the story, the other being wall interference effects.
To achieve flight Reynolds number for at least marginally affordable costs
the NTF is basically an 8 foot tunnel. Our biggest transonic wind tunnels
are 16ft square. Thus the wall interference issue still exists in the NTF.
To conclude: Using "raw" wind tunnel data and making simple skin friction
adjustments to estimate flight data may easily result in a 10% difference
between tunnel and flight. The critical areas in the flight regime are
almost always more complicated than first appearances. However, once you
work at understanding the full story, you generally feel that you can account
for the differences within a few counts. Original drag estimates made at
the time that a production decision was made may not take into account
a lot of the effects that occurred after the estimates were made, or did
not turn out in hindsight to be wise interpretations of data. Engineering
judgment is still required, and experienced engineers will be required
to make tunnel to flight correlations and analysis for the foreseeable
You rarely have the same geometry in flight that you tested. It was 15
years before Grumman had a model of the F-14 that actually corresponded
to the flight article. For the F-14, as with most configurations, the configuration
was "tweaked" one last time after the last wind tunnel test. The only reason
Grumman eventually did have an accurate F-14 was that the AMRAAM program
needed one and they were willing to pay for a new model. Other typical
situations include vortex generators added during flight test programs
but not addressed in original drag estimates/tests, and the gap/step correspondence
between wind tunnel models and flight hardware.
There always seems to be a disagreement over exact performance of the engine.
This makes accurate flight test drag values difficult to obtain. Many,
many corrections are made to the engine data. Usually there's a lot of
money at stake. Both the airframe and engine must meet performance guarantees,
with large penalties if they aren't met.
3. Presentation of aerodynamic data
The format established for the course should be used. The following comments
are specific to aerodynamic testing, and may not be generally applicable.
Text (two aspects of the test report are critical)
1) In the documentation of the test enough detail must be included from
the test to settle any question that comes up after the test is over. This
includes the test run schedule, configuration nomenclature description,
sign convention for deflecting surfaces, and the so-called "tab data" from
the test. The test engineer must keep a detailed notebook during the test,
and include extreme detail in the test report. Many well annotated photos
should be included. Many times questions arise (sometimes years later)
where the documentation is insufficient to determine with certainty exactly
what happened. Few of us can remember particular test details even after
a few months, and particularly when being grilled because something doesn't
"look right" (this is the situation as soon as the flight test data arrives,
as described above). Two typical personal examples include exact details
of transition fixing, and sign convention for deflection of surfaces (at
high angle-of-attack it may not be at all obvious what effect a "plus"
or "minus" deflection would produce on the aerodynamic results). Since
a typical test might cost several million dollars, and the data might be
examined for effects that weren't of specific interest during the test,
good documentation is crucial. An unfortunate, but frequent, occurrence
in practice is that the time and budget expire before the test report is
completed. Since it's done last, budget overruns frequently result in poor
final documentation. It's best if the report can be put together while
the test is being conducted. This approach can minimize the problem.
2) In the data analysis portion of the report: When writing the report
provide specifics, not generalities, i.e., rather than "greater than,"
say "12% greater than." What do the results mean? In large organizations
the test engineer will write a test report precisely documenting the test,
while the project aerodynamicist will write the report analyzing the data.
When writing the analysis, do not simply provide tables of numbers and
demand that the reader do the interpretation. The conclusion to be drawn
from the each figure must be precisely stated.
Plots and Graphs
Use real graph paper. For A size plots this means K&E2
Cat. No. 46 1327 for 10x10 to the half inch, and an equivalent type for
10x10 to the centimeter. There is an equivalent catalog number for B
size paper. Wind tunnel data (especially drag polars) are often plotted
on B size paper. This is Albanene tracing paper. It's what's actually
used in engineering work, and it's expensive. The University Bookstore
stocks this graph paper. You should use it carefully, and not waste it.
With high quality tracing paper, where the grid is readily visible on the
back side, you plot on the back. This allows you to make erasures and also
produces a better looking plot. Orange graph paper is standard, and generally
works better with copy machines. The tracing paper also allows you to keep
reference data on a set of plots and easily overlay other results for comparison.
Remember to allow for overlay comparisons by using the same scale for your
Always draw the axis well inside the border, leaving room for labels inside
the border of the paper. Labels should be well inside the page margins.
Data plots should contain at least:
Use proper scales. Use of "Bastard Scales" is grounds for bad grades in
class and much, much worse on the job. This means using the "1,2, or 5
rule". It simply says that the smallest division on the axis of the plot
must be easily read. Major ticks should be separated by an increment that
is an even multiple of 1, 2 or 5. For example, 10, 0.2, 50. and 0.001 are
all good increments between major ticks because it makes interpolation
between ticks easy. Increments of 40, 25, 0.125 and 60 are poor choices
of increments, and don't obey the 1,2, or 5 rule. The Boeing Scale Selection
Rules chart in included as figure 9. Label plots neatly
and fully. Use good line work. In putting lines on the page, use straight
edges and ship's curves to connect points, no freehand lines. Ship's curves
and not French curves are used by aeronautical engineers, and some catalogs
call them aeronautical engineering curves. The University Bookstore stocks
at least the most common ship's curve size. As a young engineer, I was
told that if the wind tunnel data didn't fit the ship's curve, the data
was wrong. More often than not this has indeed turned out to be the case.
Drag polars are traditionally plotted with CD on the
abscissa or X-axis, and CL on the ordinate or Y-axis.
Moment curves are frequently included with the CL -
curve as shown in Figure 10. The moment axis is plotted
from positive to negative, also shown in the figure. This allows the engineer
to rotate the graph and examine Cm-Cl in a
"normal" way to see the slope.
More comments on proper plots and graphs are contained in your engineering
graphics text, by Giesecke, et al. (Ref. 12). The engineer
traditionally puts his initials and date in the lower right hand corner
of the plot. An example of an acceptable plot is included as figure
Can you use your computer to make plots and graphs? Of course. But they
must be of engineering quality. To achieve this you certainly have to understand
the requirements given above for hand plots, and have made enough graphs
by hand to be able to identify problems in the computer generated graphs.
Often it's easier to make plots by hand than to figure out how to get your
plotting package to do an adequate job. Typical problems include poor scale
selection, poor quality printout, and inability to print the experimental
data as symbols and the theory as lines. Another problem that arises is
the use of color. While color is important, it presents a major problem
if the report is going to be copied for distribution. Most engineering
reports don't make routine use of color - yet. A final problem frequently
arises with the labeling. In reports, the figure titles go on the bottom.
On view graphs and slides the figure titles go on the top. Many graphics
packages are oriented toward placing the figures on the top. This is unacceptable
in engineering reports.
reference area, reference chord and span as appropriate (include units)
moment reference center location
Reynolds number, Mach number, and transition information
4. Strain Gage Balances
Aerodynamic forces must be accurately measured. Initial aerodynamic testing
used simple scales. Today most force measurements are made using strain
gage balances. At large facilities the strain gage balances are designed,
built and maintained by an instrumentation group that usually consists
of mechanical engineers. However, you should understand the principles
of operation. Balances must be selected for a particular test based on
the physical size and loads demanded on the balance. Each balance should
have a drawing indicating the size and on the same drawing you will usually
find a table containing the maximum loads. Sizing and placement of the
balance requires use of the pretest loads estimates made by the aerodynamics
group. Typically the strain gage balance is located inside the model at
a position near the aerodynamic center to minimize the possibility that
the pitching moment limit will be exceeded. This can be a little challenging
if the same model is to be used for both subsonic and supersonic testing,
and in cases where component build-up testing may lead to situations where
the pitching moment becomes large (also variable sweep wings). In some
cases provision for two balances might have to be made.
Frequently aerospace companies and government agencies borrow balances
to obtain the correct size. Balances are delicate and can be easily damaged
if handled improperly. Their installation in models requires precision
machining, and therefore, the balance is pinned in place. The portion of
the model that produces loads measured by the balance is known as the metric
portion of the model. Accounting for the metric and non-metric parts of
the model are sometimes tricky, especially in powered model testing, where
the forces over portions of the model are desired. Powered model testing
is one of the most challenging areas of aerodynamic testing. Frequently
the propulsion group needs more wind tunnel test time than the aerodynamics
Because the balance must reflect the loads accurately, care must be taken
when "bridging" the balance. This means that a minimum of cables and pressure
tubes should cross from the metric to the non-metric part of the installation.
In some cases this means that a scanivalve or its equivalent should be
placed inside the metric model to allow surface pressure and force and
moment results to be acquired simultaneously. The scanivalve measures pressures
from a number of orifices and allows many pressures to be transmitted to
a single pressure transducer. Thus, rather than having 48 pressure tubes
bridging the balance, the only lines that crosses the balance are the scanivalve,
control cable, and possibly a reference pressure tube. Frequently this
results in a severe temperature and vibration environment for this instrumentation,
and possible problems should be considered during pretest planning.
A strain-gage balance measures forces and moments by sensing deformation
of a beam-like element that has a strain gage attached to it. The balance
is based on the principle that the electrical resistance of a conductor
changes when subjected to mechanical deformation. A strain gage is a small,
thin printed circuit type electrical resistance element which changes its
resistance when elongated or compressed (Fig. 11a).
This gage, which may be on the order of 1/4 inch square or smaller, which
is glued onto a structural element which will be affected by the force
to be measured.
A simple strain-gage balance to measure drag can be made using two strain
gages on a single beam as shown in Fig. 11b. As a drag
force is applied, the upstream gage element feels tension and the downstream
element compression due to the bending in the beam. Since the signs of
the resistance changes on the two elements are opposite, taking their difference
by placing them in a circuit such that they subtract will result in a doubly
strong resistance change. Any change due to a pure "upward" or "downward"
force through the other beam will thus cancel out. This is important because
temperature induced effects arising from thermal expansion of the material
would result in misleading indications of an externally applied force.
The change in resistance can be detected by either using a constant voltage
supply and measuring a current change or by using a constant current and
measuring the voltage change. The resulting reading can then be calibrated
by comparing the output signal under a range of test loads. Unless the
gage or beam is over-stressed, the readout will be a linear function of
the force applied.
Only one problem exists with the balance design in Fig.
11b, the gage output is really determined by the bending moment, which
is a function of the distance between the application point of the force
and the gage element location, as well as the force magnitude, instead
of being a function of the force alone. A remedy to this is shown in Fig.
11c where 4 gage elements are used. Now, the circuitry can be set up
to read two individual moments M1 and M2 due to the applied force and subtract
them. The difference between these two moments is a function only of the
force and the fixed distance between the two sets of gage elements. Since
this distance is constant, the drag can be directly determined by a single
circuit regardless of the location of the applied force on the beam. The
circuitry for such an arrangement is shown in Fig. 11d
where gage elements 1 and 2 are in tension and 3 and 4 are in compression.
A wind tunnel strain-gage balance has at least one such "bridge" to measure
each force and moment. For more information on strain gage balance design
see Rae and Pope (Ref. 1).
5. Virginia Tech 6' X 6' Stability Wind
The full description of the Stability Wind Tunnel is located on the World
Wide Web at the following URL http://www.aoe.vt.edu/research/facilities/stabilitytunnel/index.html.
Rae, William H., Jr., and Pope, Alan, Low-Speed Wind Tunnel Testing,
Wiley & Sons, New York, 1984.
McLaughlin, C.B., "Experimental Aerodynamics and Configuration Design,"
Grumman Aerodynamics Lecture Series, May, 1985.
Toscano, E.J. "Aero Test Notes - Basics of What a Wind Tunnel Test Engineer
Should Know About Internal Strain Gage Balances, or What a Model Designer
Must Know About Internal Strain Balances and Other Things," Grumman internal
Nark, T., "Powered Lift Testing," First Atlantic Aeronautical Conference,
Williamsburg, Virginia, March 26-28, 1979. (Boeing Company Viewpoint)
Hedrick, I.G., "Using Wind Tunnel Test Results," Lecture XII, Introduction
to Aerospace Engineering, MIT, May 3, 1982.
Meyer, R., "General Information on Wind Tunnel to Flight Correlation,"
Grumman Memo EG-ARDYN-80-150, December 15, 1980.
Hancock, G.J., "Aerodynamics - the role of the computer," Aeronautical
Journal, August/September, 1985, pp. 269-279.
Braslow, A.L., Hicks, R.M., and Harris, R.V., Jr., "Use of Grit-Type Boundary-Layer-Transition
Trips on Wind-Tunnel Models," NASA TN D-3579, Sept. 1966.
Blackwell, J. A., Jr., "Experimental Testing at Transonic Speeds," in
Aerodynamics, ed. by Nixon, D., AIAA Progress in Astronautics and Aeronautics,
Vol. 81, AIAA, New York, 1982, pp. 189-238.
E.J. Saltzman and T.G. Ayers, "Review of Flight to Wind Tunnel Drag Correlation,"
Journal of Aircraft, Vol. 19, No. 10, October 1982, pp 801-811.
"Wind-Tunnel/Flight Correlation Study of a Large Flexible Supersonic Cruise
Giesecke, F.E., Mitchell, A., Spencer, H.C., Hill, I.L., Loving, R.O.,
and Dygdon, J.T., Principles of Engineering Graphics, Macmillan
Publishing Cop., 1990, pp. 591-613.
Ahn, Seungki, Choi, Kwang-Yoon, and Simpson, Roger L., "The Design and
Development of a Dynamic Plunge-Pitch-Roll Model Mount," AIAA Paper 89-0048,
Part I, NASA TP 1514, James C. Daugherty, Nov. 1979.
Part II, NASA TP 1515, John B. Peterson, Jr., Mike Mann, Russell Sorrells,
Wally Sawyer and Dennis Fuller, Feb. 1980
Part III, NASA TP 1516, Henry Arnaiz, John B. Peterson, Jr. and James C.
Daugherty, Mar. 1980
1 This appendix was reviewed at
Grumman under the direction of Casper Catalanotto, Group Lab Manager, Aerodynamics
Test, Ground Testing Technology, Grumman Aircraft. John McAfee and Howard
Jarvis contributed. Revisions were made to incorporate their comments.
2 This paper is very high quality
paper. With computers replace hand plotting, this paper is being discontinued
by K&E. Most art supplies stores (sometimes erroneously claiming to
be engineering supply stores also) don't stock good graph paper. Cheap
paper will not be transparent, allowing easing tracing from one plot to
Figure 1 Example of the testing required to develop a modern fighter
aircraft (Ref. 7).
Figure 2. Typical run schedule: transonic, run numbers under Mach
Figure 3. Typical nomenclature chart associated with run schedule.
Figure 4. Typical run schedule with run numbers on the left hand
side. Defining data requirements
a) Typical methods of attaching tufts. Only A is suitable for high
speed work. (Ref. 1)
b) Typical tuft layout on the upper surface of a wing (Ref.
Figure 5. Illustration of the use of tufts for flow
visualization. from Grumman Aero Report 393-81-1, October, 1981, "Experimental
Pressure Distributions and Aerodynamic Characteristics of Flat and Cambered
Conceptual Wing-Body and Wing-Body-Canard Models at M=1.62," by W.H. Mason.
Figure 6. Oil flow photograph of an open separation at M = 1.62
Figure 7. Relation between body and wind axis (Ref.
Figure 8. Example of sources of wind tunnel effects requiring data
Figure 9. Boeing scale selection chart (AIAA Student Journal, April,
1971) from Grumman Aero Report No. 393-82-02, April, 1982, "Experimental
Pressure Distributions and Aerodynamic Characteristics of a Demonstration
Wing for a Wing Concept for Supersonic Maneuvering," by W.H. Mason
a) lift and moment
b) drag polar
Figure 10. Examples of wind tunnel data plots.
from Grumman Aero Report No. 393-82-02, April, 1982, "Experimental Pressure
Distributions and Aerodynamic Characteristics of a Demonstration Wing for
a Wing Concept for Supersonic Maneuvering," by W.H. Mason
Figure 11. Strain gage balance data