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Background
Flutter is a dynamic aeroelastic
phenomenon that involves the interactions of elastic
and inertia forces of the structure with the aerodynamic
forces produced by the airflow over the vehicle.
It is a self-excited oscillation of the aircraft
structure, where energy is absorbed from the airstream.
When the elastic structure of the aircraft is disturbed
at speeds below the flutter speed, the resulting
oscillatory motions decay. However, when the structure
is disturbed at speeds above the flutter speed,
the oscillatory motions will abruptly increase
in amplitude and can ultimately lead to catastrophic
failure of the structure. In some instances, flutter
oscillations are limited to just a single airplane
component such as the wing, whereas in other instances
the oscillations may be considerably more complex
and involve coupling of wing, fuselage, and empennage
vibrations.
The process of ensuring that
an airplane is flutter free, commonly referred
to as “flutter clearance,” involves
carefully planned analytical and experimental studies
of the complex aeroelastic interactions that the
aircraft will experience throughout its flight
envelope with sufficient margins beyond the expected
flight conditions. If a significant flutter tendency
is discovered within the operational envelope,
modifications to the aircraft structure must be
made. The most common modification is to increase
the most critical structural stiffness, which usually
results in an increase in structural weight. This
increase in weight is commonly referred to as the
“flutter weight penalty.” The civil
and military aircraft industries conduct extensive
analyses of flutter characteristics by using company-owned
facilities and computer codes during aircraft development
programs. Many of the design teams consult with
Langley flutter experts and conduct proprietary
or cooperative studies using the unique transonic
aeroelastic testing capability provided by the
Langley Transonic Dynamics Tunnel (TDT).
Flutter clearance testing
in the TDT for military aircraft configurations
has become a typical element in most aircraft programs,
and the emphasis is placed along the developmental
requirements for the specific aircraft program
at that time. Unlike military airplane testing,
TDT testing of civil aircraft frequently has a
research element included in the test program in
addition to developmental requirements. Many of
these research activities include in-depth assessments
of analytical codes, developing and evaluating
advanced flutter control mechanisms, and determining
the impact on aeroelastic phenomena of advanced
aerodynamic or structural concepts such as airfoils
for next-generation aircraft. Thus, civil aircraft
flutter testing at Langley is typically more closely
coupled to fundamental research than military aircraft
studies. Tests of this type ensure that flutter
or other undesirable aeroelastic problems that
may exist for a new design are identified early
enough in the design-development cycle that a solution
(fix) can be identified in a timely manner with
minimum impact on cost and schedule. In addition,
wind-tunnel tests such as those described herein
reduce the number of more costly flight flutter
tests.
Transonic Dynamics Tunnel
Aeroelastic testing for flutter
characteristics of civil aircraft using dynamically
scaled aeroelastic models has taken place for decades
in many subsonic wind tunnels within industry and
academia. However, aeroelastic problems at higher
subsonic speeds, near the transonic speed range,
became more critical and increasingly difficult
to predict as the performance and flight envelopes
of aircraft expanded. In response to requests for
a dedicated facility for testing large flutter
models at relatively high Reynolds numbers and
transonic speeds, NACA converted its existing Langley
19-Foot Transonic Pressure Tunnel into a new aeroelastic
testing facility with a 16- by 16-ft test section
that could operate at Mach numbers up to 1.20 with
variable pressure conditions in either air or a
heavy gas known as dichlorodifluoromethane (DuPont
Freon 12 gas). Design of the new facility began
in 1954, and the new Langley Transonic Dynamics
Tunnel (TDT) began research operations under the
management of Robert W. Boswinkle, Jr. and his
assistant, D. William Connor, in early 1960. The
first research tests by Jerome T. Foughner, Jr.,
and Norman S. Land began on February 5, 1960, with
air as the testing medium. In March 1960, calibration
tests were run with Freon 12 gas. Stagnation pressures
in the TDT vary from near-vacuum to atmospheric
conditions, which simulate variations in flight
altitude. Testing in a heavy gas provides aeroelastic
model scaling advantages because the density of
Freon 12 gas is approximately four times that of
air, and it has a speed of sound of about half
that of air. As a result, scaled models for flutter
testing can be made heavier than a scale model
that would be tested in air. This approach makes
the task of building a scale model with sufficient
strength and appropriate structural frequencies
much easier. In addition, the lower speed of sound
of Freon 12 gas results in a model having lower
frequencies of vibration; thereby more reaction
time is provided for the test crew to back off
from a potentially catastrophic test condition.
Today, the TDT is the world’s
premier wind tunnel for testing aeroelastically
scaled models at transonic speeds. Operational
since 1960, the tunnel has been the site of a wide
variety of investigations ranging from flutter
clearance tests to fundamental research on aeroelastic
phenomena for military and civil aircraft. The
scope of tests has included
-
Flutter clearance aimed at reducing
risk and defining potential flutter problems
and potential solutions through studies of aircraft
components as well as complete configurations
-
Parametric variations of aircraft
parameters in risk reduction tests to help guide
flight tests
-
Studies to analyze and solve aeroelastic
problems of specific configurations
-
Code calibration tests conducted
in conjunction with flutter clearance tests to
obtain data for validation and extensions of
flutter prediction methods
In the 1980s, environmental concerns regarding
the impact of Freon 12 gas caused NASA to convert
the test medium for heavy gas testing in the TDT
to R-134a, an inert gas with properties similar
to those of DuPont Freon 12 gas.
Several types of model mount
systems are used for tests in the TDT. Sidewall
mount systems provide the capability of testing
semispan wing models, which are generally easier
to build and less expensive than full-span models.
Conventional sting-support mount systems are also
utilized for tests of full-span models or aircraft
components. The most unique model mount system,
however, is the cable mount system developed by
Wilmer H. Reed III, which is most compatible with
aeroelastically scaled full-span model testing.
For this system, a full-span model is supported
in the test section by a system of two cables routed
through a system of pulleys. Additional cables
are also provided for a “snubber” system
that can be manually tightened during a model instability
to prevent loss of the model. In a typical test
program, it is customary to first “fly”
a cheaper rigid or dummy model on the cable mount
system to demonstrate the stability of the model
configuration on the cable mount system prior to
testing the expensive, flexible flutter model.
Langley Transonic Dynamics
Tunnel.
Plan view of TDT.
The TDT is equipped with several
features that accommodate the unique challenges
of flutter testing. For example, the tunnel is
equipped with a group of four bypass valves connecting
the test section area (plenum) to the opposite
leg of the wind-tunnel circuit downstream of the
drive fan motor. When opened, the bypass valves
cause a rapid reduction in the test section Mach
number and dynamic pressure. This feature is available
for use when model flutter instabilities result
in motions that could damage a model. In addition
to potentially saving a model from catastrophic
destruction, the rapid reduction in tunnel flow
conditions reduces the inertial force of the model
debris that might impact the tunnel drive motor
fan blades. An additional feature of the tunnel
is a model debris catch screen located upstream
of the tunnel fan blades. The tunnel is also equipped
with a pair of vanes on each side of the test section
entrance cone. These vanes can be oscillated over
a range of frequencies and amplitudes to generate
sinusoidally varying vertical flow velocities for
use in gust response studies. High-speed motion-picture
cameras and video recorders and cameras are available
to capture the rapidly occurring aeroelastic events
so they can be viewed and analyzed later in slow-motion
detail. The tunnel is also equipped with state-of-the-art
instrumentation for high-frequency measurements
and data acquisition and analysis. Finally, a sophisticated
gas pumping and handling system is available for
processing, reclaiming, liquefying, and storing
the R-134a for reuse.
Flutter testing in the TDT
requires the use of special models. A reduced-size,
dynamically scaled aeroelastic model is essentially
a mechanical analog of the full-scale aircraft.
If the appropriate model-to-full-scale scaling
relationships are used to design and construct
a model, then its behavior during wind-tunnel tests
can be used to accurately predict full-scale behavior.
For example, if the flutter velocity is obtained
for a properly designed, constructed, and tested
model, then that value can be used with confidence
to predict the full-scale flutter velocity.
For dynamically scaled aeroelastic
models, there are several nondimensional parameters
that must be identical for both the model and airplane
to produce meaningful results. First, the geometry
of the model must be the same as that of the full-scale
aircraft. For instance, the angle of wing sweep
and wing taper ratio must be identical between
model and aircraft. The size of the model (the
length scale factor) and the test conditions—fluid
density, Mach number, velocity—at which the
model is to represent the airplane are dictated
by the size and operating envelope of the wind
tunnel chosen for the test.
For flutter models, the three
most important parameters in addition to size and
geometry are Mach number, reduced frequency, and
mass ratio. The Mach number is a speed-related
parameter that is associated with the compressibility
characteristics of the test or flight medium. If
compressibility effects are deemed important, then
the model must be tested at the same Mach number
as the airplane. Reduced frequency can be thought
of as a time-related parameter. Satisfying the
requirements of this parameter ensures that the
relative time scale for the model and airplane
are the same and that the time required for events
on the model can be correctly scaled to full-scale
time. The ratio of model natural frequencies to
full-scale frequencies depends on this parameter.
Mass ratio is an altitude-related parameter that
relates fluid density and structural mass. The
mass of the model relative to the mass of the full-scale
aircraft is determined by this parameter. The structural
stiffness of the model is determined once the natural
frequency and mass requirements are defined because
they are interrelated.
Reynolds number, related to
the fluid viscous, or fluid friction effects, is
also an important test parameter. Matching this
parameter between model and full scale is usually
impossible; therefore, customary practice is to
try to test at as high a Reynolds number as possible.
Because the value of Reynolds number is directly
proportional to model size and fluid density, the
large size of the TDT and the high density of the
R-134a test medium offer the possibility for relatively
large Reynolds numbers for tests. For instances
where the effects of the gravitational force are
important, the requirements of a parameter known
as the Froude number must be satisfied. When compressibility
effects are important, simultaneously satisfying
the scaling requirements of Froude number and the
other three parameters is generally impossible.
This can be accomplished, however, by testing an
approximately 0.25-scale model in heavy gas in
the TDT. This feature is yet another unique one
of this facility.
Flutter model of Boeing
747 aircraft with typical spar-pod construction.
Model of Boeing 767 aircraft
with continuous skin construction.
Flutter models of full-scale
aircraft fall into essentially two categories:
low-speed models and high-speed models. For low-speed
models, compressibility effects are not important
and the customary practice is to match only the
reduced-frequency and mass-ratio parameters. Froude
number simulation for these models is also possible.
Models of this type are often fabricated by using
what is referred to as spar- and segmented-pod
construction. A central metal spar is designed
to have the appropriately scaled level and distribution
of structural stiffness. This spar is enclosed
by a series of very lightweight pods, often fabricated
from wood, that provide the proper external geometry.
Because the span of a wing is segmented into a
number of such pods, each pod is only connected
to the spar at one or two points and not connected
to adjacent pods. Thus, the pods do not contribute
appreciably to the structural stiffness. Typically,
the mass of the spar and pods is not sufficient
to simulate the scaled mass of the full-scale structure
and/or fuel; therefore, metal weights are attached
inside the pods to provide proper inertia characteristics.
For high-speed flutter models,
matching Mach number—in addition to matching
the reduced-frequency and mass-ratio parameters—is
necessary. Although spar-pod construction is still
used in some instances for models of this type,
the trend is toward using continuous skin models
because differential movement between adjacent
pods creates a steplike irregularity in the surface
contour that might cause undesirable aerodynamic
interference effects for some configurations at
transonic speeds. A continuous-skin model might
be constructed by laying a fiberglass skin of varying
thickness over a lightweight honeycomb core that
has been shaped to the proper planform and airfoil
section geometries. The skin is bonded to the honeycomb
core to form a sandwichlike structure, and the
fiberglass skin primarily contributes the structural
stiffness. The thickness of the skin is varied
to account for required variations in structural
stiffness along the wingspan and chord. To achieve
proper inertia characteristics, metal weights are
imbedded within the honeycomb core.
In its operational history, the TDT has been used
for over 550 tests of military and civil aircraft.
The following examples discuss some of the more
important contributions of Langley research using
the facility for civil aircraft.
Applications to Civil Aircraft
Lockheed Electra Aircraft
The first TDT test focusing
on a particular full-scale design was conducted
in 1960 and used a model of the four-engine turboprop
Lockheed Electra commercial transport. The relatively
new Electra, having been introduced into airline
service in the fall of 1958, could carry 98 passengers
more than 2,000 miles while cruising at speeds
up to 400 mph. Because a number of accidents, some
unexplained, involving the Electra had severely
damaged the public’s confidence in airline
transportation, a virtual crisis had developed
within the air transportation industry. The TDT
began operating just in time to assist in understanding
the causes of some of the accidents that had befallen
this relatively new airplane.
Soon after its introduction
to the civil transport fleet, the Electra suffered
a number of accidents, with the wreckage of two
widely publicized fatal accidents raising concerns
about the structural integrity of the airplane.
On September 29, 1959, a Braniff Electra with 5
crew members and 28 passengers enroute from Houston
to Dallas disintegrated near Buffalo, Texas without
survivors. Investigation of the aircraft wreckage
revealed that the left wing had failed and separated
from the airplane in flight. The outboard engine
nacelle on the failed wing displayed evidence that
the propeller and gearbox had twisted over 30∞
out of alignment with the wing. Although the cause
of the accident remained unidentified, catastrophic
flutter was among the candidate factors. Then,
on March 17, 1960, a Northwest Airlines Electra
carrying 6 crew members and 57 passengers from
Minneapolis to Miami crashed in Indiana with a
startling similarity to the Texas accident. Its
right wing was found over 11,000 ft from the crash
site, indicating that it had also been torn from
the airplane. Again, the outboard engine nacelle
of the failed wing panel showed the same twisted
characteristics as the earlier accident. With over
130 Electras operating in the civil fleet at the
time, immediate action was required to ensure that
these airplanes operated safely. Consequently,
cognizant authorities immediately reduced the cruise
speed of the airliners while the investigation
attempted to identify the cause of the fatal crashes.
Meanwhile, public and congressional concern over
the safety of the Electra and commercial aviation
in general rapidly increased with a demand for
an explanation of the causes of the crashes and
assurance that the safety of the flying public
was guaranteed during continued operations. The
Electra problems, therefore, became a problem for
the entire air transport industry.
As suspicions of flutter-induced
structural failure continued to grow within the
investigation groups, key managerial interactions
between industry, the Civil Aeronautics Board (precursor
of today’s National Transportation Safety
Board), and Langley concluded that flutter tests
of an Electra model in the newly commissioned TDT
would be beneficial to explaining the cause of
the accidents. The purpose of these tests would
be to determine whether the airplane might have
experienced propeller-whirl flutter as Lockheed
flutter engineers were beginning to suspect. Propeller-whirl
flutter, first identified in the late 1930s, is
characterized by a wobbling motion of the propeller
resulting from gyroscopic coupling of the rotating-propeller–engine
system with the elastic wing through flexible engine
nacelle supports. No airplane had encountered it
previously, but the advanced Electra design with
its large, high rpm (revolutions per minute) turboprop
engines might have just the combination of parameters
necessary to produce propeller-whirl flutter. Langley’s
Philip Donely, I. Edward Garrick, John C. Houbolt,
Robert W. Boswinkle, Jr., and Dennis J. Martin
led Langley’s involvement in the investigation.
An existing 1/8-size full-span
Lockheed model that had been used previously in
the Electra flutter clearance program was modified
to meet the needs of the study. Among the modifications
were the incorporation of windmilling propellers
and the facility to adjust engine mount stiffness.
The model was mounted on a Boeing-developed vertical-rod
support system that allowed for a limited simulation
of free-flight conditions. Lockheed Aircraft Corporation,
Boeing Airplane Company, and NASA engineers participated
in the tests. Nine different tunnel entries occurred
between May 1960 and December 1961. In addition
to tests of the full-span model, studies were conducted
with an isolated propeller-nacelle model and a
sidewall-mounted semispan wing-nacelle model. In
addition to the experimental activities in the
TDT, staff members Wilmer H. Reed III and Samuel
R. Bland developed mathematical methods for analysis
of the propeller-whirl phenomenon and helped validate
refined models for the prediction and elimination
of such problems in the future.
Powered model of Lockheed
Electra mounted in Langley Transonic Dynamics Tunnel
for flutter tests.
The TDT test results showed
that propeller-whirl flutter could, in fact, occur
for the Electra, but only if the engine mount stiffness
was reduced below the nominal design value (as
might be caused by a structural failure during
a hard landing or an encounter with extremely severe
turbulence in flight). During the model tests,
no flutter was observed for the model in the basic
as-designed condition. However, when the NASA and
industry team reduced the scaled stiffness of the
engine mounts on the outer engine nacelles, the
whirl-flutter mode was observed. It was predicted
that the fatal resonance could build up and tear
the full-scale airplane apart in a matter of seconds.
The problem was solved by increasing the structural
stiffness and damping designs of the engine mounts
and nacelle. Based on these results from the TDT,
the engine mounts on all Electra aircraft were
modified and strengthened, and the modified Electra
(known as the Super Electra II) entered service
in 1961. The modified Electra, and its military
derivative Navy P-3 Orion patrol aircraft, have
since operated successfully without flutter issues.
Electra model following
catastrophic flutter.
The Electra test was an extraordinarily
difficult challenge for a new facility with a newly
assembled staff. The challenge required marshalling
the efforts of practically everyone associated
with the facility to ensure its success. The dramatic
investigation of the Electra in the TDT had many
impacts. First, the airline industry and Lockheed
were provided a graphic and accurate definition
of the causes and cures for the crashes. The solution
to the problem provided a badly needed boost to
the public’s confidence in air travel and,
no doubt, played a key role in the continued expansion
of commercial aviation. The success of the TDT
investigation, conducted in the international spotlight
with intense scrutiny, properly portrayed the professionalism,
dedication, and value of the Langley staff and
facilities.
Boeing 747 Aircraft
A flutter model of the Boeing
747 was tested twice during separate entries in
a cooperative NASA and Boeing study in the TDT
during 1967–1968. The investigations, led
by Langley’s Moses G. Farmer and Irving Abel,
were focused on determining the effects of the
relatively large engine cowls of the high-bypass
engines on flutter characteristics of the aircraft.
The 4.6-percent scale, full-span model was tested
on two different mount systems in the TDT—the
Boeing-developed vertical-rod mount system that
had found extensive use in low-speed flutter model
tests, and the newer Langley-developed cable mount
system that had become the system of choice for
TDT tests. Model parameters for the tests included
nacelle aerodynamics, engine-pylon stiffness, mount
system, and mass ratio. The aerodynamic effects
of the nacelles on flutter characteristics were
determined by replacing the engine nacelles with
“pencil nacelles” that simulated the
inertia and center-of-gravity characteristics of
the engine nacelles.
Researcher Irving Abel with
Boeing 747 flutter model.
Results from the tests indicated
that the aerodynamic forces for the simulated high-bypass-ratio
engines reduced the flutter speed by about 20 percent.
The flutter characteristics were greatly dependent
on the outboard engine lateral frequency. The effects
of mount systems on flutter speed were small.
Lockheed L-1011 Aircraft
In 1969, Langley and Lockheed
collaborated for cooperative tests of the Lockheed
L-1011 in the TDT. The objectives of the tests
included flutter clearance as well as general research
on the effects of supercritical airfoils on flutter
characteristics. Five tests used L-1011 models,
the first being conducted in January 1969 and the
last in October 1969. The first four were flutter
clearance tests and utilized full-span cable-mounted
models. The fifth test was of a research nature
and used semispan sidewall-mounted L-1011 wing
models to evaluate the effects of the supercritical
airfoil shape on flutter characteristics. The actual
L-1011 aircraft did not incorporate a supercritical
airfoil, but Lockheed’s interest in future
applications of supercritical technology led them
to furnish a conventional-airfoil wing and a supercritical-airfoil
wing for the test. This investigation was the first
such experimental study of the impact of supercritical
airfoil on flutter. Several TDT engineers played
important roles in the L-1011 tests, including
Moses G. Farmer, Rodney L. Duncan, Perry W. Hanson,
and Robert V. Doggett, Jr.
Lockheed engineer with L-1011
flutter model in TDT.
DC-10 Aircraft
The TDT was used in November
1969 and July 1970 to evaluate the flutter characteristics
of the unique split-rudder empennage configuration
of the DC-10. The proposed rudder design was split
into two spanwise sections, rather than being the
single section rudder found on most airplanes.
This test, directed by Maynard C. Sandford, used
an empennage and aft-body model of the DC-10 on
a sting mount. Specific objectives of the program
were to determine the flutter characteristics of
the vertical tail with a split rudder configuration
and compare those results with the flutter characteristics
of a vertical tail with a conventional single section
rudder design. The results of this TDT study showed
that the split rudder had a beneficial effect on
flutter by requiring less structural stiffness
to prevent flutter than that needed for a conventional
single section rudder.
Maynard C. Sandford and
empennage and aft-fuselage flutter model of McDonnell
Douglas DC-10.
Gulfstream American III Aircraft
A model representative of
the Gulfstream III configuration with a proposed
supercritical wing was tested twice in the TDT
in 1978 by a team of NASA and Gulfstream American
Corporation engineers led by Charles L. Ruhlin.
This joint NASA and Gulfstream program used a 1/6.5-scale
semispan model of the Gulfstream III wing. The
wing was mounted to a rigid half-body fuselage
so that the effects of fuselage aerodynamics could
be properly accounted for. The wing could be configured
with three different wingtip arrangements: a nominal
wingtip, a wingtip with a winglet, and a normal
wingtip ballasted to simulate the winglet mass
properties. Specific test objectives were to determine
the effects of the winglet on the flutter characteristics
of a supercritical wing, compare the results with
analytical predictions, investigate possible angle-of-attack-induced
flutter issues, and examine the effects of aeroelastic
deformations on aerodynamic characteristics of
the supercritical wing. Transonic flutter characteristics
were measured for each configuration over a Mach
number range from 0.60 to 0.95. Although some tests
of the winglet effects on flutter had been performed
previously with simple research models, this was
the first such study of those effects using a wing
model of an advanced civil transport airplane.
The shape of the transonic
flutter boundary observed for all three configurations
tested was consistent with what has been observed
for many other configurations in the past, that
is, a gradual decrease in flutter speed with a
minimum occurring at a transonic Mach number (0.82
in this instance) followed by an increase in flutter
speed as Mach number increases. However, the magnitude
of the flutter speed was different for the three
configurations. The flutter speeds of the ballasted
tip and winglet configurations were lower than
those of the normal tip configuration throughout
the test Mach number range. A comparison of these
data showed that most of the reduction resulting
from the addition of the winglet was due to winglet
mass effects rather than winglet aerodynamic effects.
Calculated flutter boundaries obtained by using
doublet lattice unsteady aerodynamic theory correlated
well with the experimental results up to Mach 0.82.
Patricia Cole inspects semispan
model of Gulfstream III in TDT.
Boeing 767 Aircraft
Two 1/10-scale models representative
of the Boeing 767 were tested in the TDT in August
1979. This cooperative Boeing and NASA study was
led by Charles L. Ruhlin of Langley. Both a low-speed
model and a high-speed model were tested. A low-speed
flutter model is designed without consideration
of Mach number, or compressibility, effects; a
high-speed flutter model is designed with Mach
number effects taken into consideration.
The objectives of the low-speed-model
test were to study the effects of mass-density
ratio (an altitude-related parameter) on flutter
and to provide experimental data for correlation
with and evaluation of analytical flutter prediction
methods. This model was tested with and without
a wingtip-mounted winglet. The analytical data
correlated reasonably well with the experimental
results, although there was some discrepancy in
the conditions at which a change in flutter mode
of vibration occurred.
Stanley Cole with representative
Boeing 767 model mounted for general research on
limit cycle oscillations in TDT.
The objectives of the high-speed-model
tests were to determine the effects of Mach number
on the wing flutter boundary, to determine the
effects of a wingtip-mounted winglet on the wing
flutter boundary, and to correlate the experimental
data with analytical results. The wing was tested
without a winglet, with a winglet, and with a tip
mass equal to the mass of the winglet. Results
were obtained for variations in simulated wing
fuel loading and engine mount stiffness. The results
showed that the addition of the winglet significantly
reduced the wing flutter speed and that winglet
aerodynamic effects caused most of this reduction
in flutter speed. This result was different from
the Gulfstream III results previously described,
where winglet mass effects were the primary cause
of the reduction in flutter speed.
Subsequently, the high-speed
flutter model was used in a cooperative effort
between NASA and Boeing to investigate and understand
the various aeroelastic phenomena associated with
advanced high-speed transport configurations and
to provide a database to evaluate unsteady aerodynamic
theories and aeroelastic analysis methods. Key
Langley participants in these studies were Stanley
R. Cole and Donald F. Keller. In addition to flutter
and buffet, the study specifically addressed the
nonlinear phenomenon known as limit cycle oscillations
(LCOs). LCO, which is related to flutter, is a
limited-amplitude, self-sustaining oscillation
that may occur at transonic speeds when shock waves
are present on the airfoil. These waves move back
and forth and interact with the boundary layer
in complex manners to force the wing to vibrate.
Data from tests such as this one provided important
insight to aid in the understanding of LCO and
to assess analytical methods being developed to
predict its occurrence.
NASA Propfan Test Assessment
Airplane
Although this airplane was
a one-of-a-kind vehicle, its TDT test is noteworthy
and significant. As part of the NASA Propfan Test
Assessment (PTA) Program, a 1/9-scale flutter model
of the flight test vehicle, a Grumman Gulfstream
II airplane with an advanced turboprop mounted
on the left wing, was tested in August 1985. The
full-span model was supported on the two-cable
suspension system to provide for the simulation
of rigid body freedoms. Results using this model,
which was built by Lockheed-Georgia, showed that
the flight test airplane was safe from flutter
throughout its intended operating boundary; this
ensured that it would not be necessary to restrict
the airplane research flight envelope because of
flutter. In 1987 the National Aeronautics Association
awarded the Collier Trophy to the Advanced Turboprop
Program, of which this test was an important part.
Several members of the TDT test team were included
as members of the award team in recognition of
their significant contributions to flutter clearance
of the flight test research airplane.
Propeller Test Assessment
configuration in TDT with Langley’s Michael
Durham.
Boeing 777 Aircraft
A dynamically scaled semispan
aeroelastic model of the Boeing 777 wing was tested
in the TDT in late July 1992 as part of the flutter
clearance program. The primary NASA test engineers
were James R. Florance and Moses G. Farmer. A rigid
fuselage half-body simulated the interactive airflow
and the effects between the fuselage and wing.
The model engine nacelle was designed to simulate
air mass flow through the aircraft engine, and
the model was designed so that the amount of simulated
fuel in the wing and the stiffness of the engine
pylon could be changed remotely; thus, testing
was expedited. Model angle of attack could be remotely
controlled as well.
Ten configurations were tested
throughout the simulated flight envelope of the
aircraft without obtaining flutter. Parameters
that were varied included wing fuel, engine pylon
stiffness, and the stiffness of the structure that
attached the wing to the fuselage. To create a
configuration for which flutter would occur within
the TDT operating envelope, a mass was installed
in the wingtip and results were obtained for correlation
with predictions. The measured flutter boundary
agreed fairly well with preliminary analytical
predictions for this configuration. The test results
were used to calibrate Boeing’s analytical
flutter codes used during the flutter safety certification
of the 777 aircraft.
Boeing 777 flutter model
mounted for TDT tests.
Gulfstream V
In 1994, Langley and Gulfstream
Aerospace Corporation teamed for a cooperative
flutter study with emphasis on code calibration
and the effect of winglets on flutter characteristics.
Three TDT entries were made under the Langley leadership
of Donald F. Keller. A semispan model of the Gulfstream
V was used, with a wing structure consisting of
an aluminum plate of varying thickness to which
balsa wood was bonded and contoured to form a supercritical
airfoil. A winglet was mounted at the wingtip,
and a fuselage half-body fairing was used to provide
more realistic wing-root aerodynamics. The baseline
configuration consisted of the wing, root fairing,
and winglet. The model was also tested without
the winglet and without the winglet but with a
tip boom in place to simulate the winglet mass
with negligible aerodynamic effects. The test program
also included an advanced design winglet.
Flutter results for the three
configurations indicated that the winglet effects
on flutter were mostly caused by the mass of the
winglet rather than its aerodynamic effects, a
result similar to that found previously for the
Gulfstream III design but different from that found
for the Boeing 767 model.
The approach in this test
differed from the usual flutter clearance studies.
The model was essentially geometrically correct
but had a simple structural representation. Consequently,
it was considerably less expensive to design and
fabricate as compared with most flutter clearance
models. The primary purpose from the flutter clearance
point of view was to acquire high quality experimental
flutter data for correlation with analytical results
obtained by using advanced analysis methods. The
experimental data were used to “calibrate”
the analysis that was used in the design and certification
of the airplane. This approach clearly proved quite
successful, as the Gulfstream V turned out to be
an outstanding design that would go on to be awarded
the Collier Trophy in 1997.
Donald Keller with Gulfstream
V model during winglet flutter investigation.
Cessna Citation X Aircraft
Business jet aircraft must
be designed so that flutter will not occur within
the flight envelope with a 20-percent safety margin.
Traditionally, wind-tunnel model tests have played
an important role in the flutter certification
process of new designs. The objective of the cooperative
study with Cessna Aircraft Company was to provide
wind-tunnel flutter data for use in ensuring that
the wing of the Citation X would be safe from flutter.
Flutter models of the Cessna
Citation X business jet were tested on three occasions
in the TDT during 1993 and 1994. In the first test
program during February 1993, a NASA and Cessna
team led by Langley’s José A. “Tony”
Rivera, Jr., and Moses G. Farmer conducted tests
of a 1/4-scale semispan aeroelastic model of the
Citation X wing. A rigid fuselage half-body and
flow-through nacelle were used to simulate their
interference effects on flow over the wing. The
wing model included an aileron that could be tested
undeflected or deflected. Eight configurations
were tested to obtain data to correlate with flutter
and aileron-reversal analyses. A wingtip-mounted
aerodynamic exciter was used extensively during
the test to track frequencies and estimate damping
as the flutter boundaries were approached. For
nominal aileron actuator stiffness, flutter analyses
by Cessna predicted that the flutter boundary would
be outside the full-scale aircraft flight envelope
with a 20-percent margin. The experimental flutter
points obtained for this configuration in the TDT
correlated well with the analytical results, and
they indicated that the Citation X wing would be
safe from flutter.
Cessna’s Craig Mundt
with Citation X model during wing flutter tests.
In January and February of
1994, additional testing of a 1/7-scale sting-mounted,
full-span aeroelastic model of the Citation X was
conducted in the TDT for a range of Mach numbers
from 0.60 to 0.97. The Langley leaders for this
series of tests were Donald F. Keller and Moses
G. Farmer. Based on the analyses and results of
the previous flutter tests of the semispan-wing
model, the empennage surfaces of the Citation X
configuration were considered the critical structural
components for flutter. The rotational stiffness
and mass balance of the elevator and rudder were
changed to simulate hydraulic-system-failure conditions
and control-surface center-of-gravity variations,
respectively. Eight configurations were tested,
and flutter or lightly damped oscillations were
only encountered outside the scaled operating envelope.
In general, flutter analyses performed before the
test correlated well with the experimental results.
Full-span flutter model
of Citation X.
The test results were used
by Cessna to provide guidance during the flight
flutter clearance program for the Citation X and
aided the NASA and Cessna team in understanding
aeroelastic phenomena encountered during the clearance
program. The experimental flutter results were
also useful for the validation and calibration
of analytical flutter codes.
Learjet Model 45 Aircraft
In May 1995, Langley and the
Learjet Corporation teamed to conduct flutter model
studies of the new Learjet 45 business jet. These
tests were part of the airplane flutter clearance
program. Learjet furnished a 1/6-scale sting-mounted,
full-span aeroelastic model that had flexible lifting
surfaces and a rigid fuselage. This model was tested
twice in the TDT. Stanley R. Cole, James R. Florance,
and Elizabeth Lee-Rauch were the Langley project
engineers for the first test; Langley’s José
A. “Tony” Rivera, Jr., and Donald F.
Keller led the test team for the second test. The
primary objectives of the tests were to ensure
that flutter would not occur within the scaled
flight envelope with a flutter safety margin of
20 percent, evaluate the impact of free play and
jammed control surfaces on flutter characteristics,
measure the transonic flutter conditions for a
modified wing configuration, and obtain data to
evaluate the application of linear flutter prediction
codes for transonic Mach numbers. The test was
very successful with all these objectives being
met. The nominal vehicle was shown to be flutter
free up to speeds 20 percent beyond the dive speed
and dive Mach number. Component-failure configurations,
such as excess control-surface free play and control-surface
mass balance variations, were shown to be flutter
free up to the dive speed and dive Mach number.
In addition, transonic flutter was measured for
a modified wing configuration to evaluate linear
flutter prediction codes. Compared with the experiment,
the linear flutter predictions were approximately
10 percent conservative. These valuable data were
used by Learjet to minimize the risk and flight
test time for flutter clearance flight tests of
the Learjet 45 aircraft.
Stanley Cole with flutter
model of Learjet 45 in TDT.
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