A two dimensional oscillating airfoil test apparatus is presented as a method of measuring unsteady aerodynamic forces on an airfoil or rotor blade section. The oscillating airfoil test rig, which is being built for use in an 11 X 11-foot transonic wind tunnel (speed range M = 0.4 - 1.4), will allow determination of unsteady loadings and detailed pressure distributions on representative airfoil sections undergoing simulated pitching and flapping motions. The design details of the motion generating system and supporting structure are presented. This apparatus is now in the construction phase

Gibson, F. L.

Hocker, A. J., Jr.

Matsuhiro, D. S.

English

A two-dimensional oscillating airfoil test apparatus is presented as a method of measuring unsteady aerodynamic forces on an airfoil or rotor blade section. The oscillating airfoil test rig, which was built for use in NASA Ames Research Center's 11x11-foot Transonic Wind Tunnel (speed range M = 0.4 - 1.4), allows determination of unsteady loadings and detailed pressure distributions on representative airfoil sections undergoing simulated pitching and flapping motions. The design details of the motion generating system and supporting structure are presented

NASA Technical Reports Server

https://ntrs.nasa.gov/search.jsp?R=19780011099 2020-03-20T12:52:34+00:00Z
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T_IO-DIMENSIONAL OSCILLATING AIRFOIL TEST APPARATUS
By Frank L. Gibson, Andrew J. Hocker, Jr.,
and Dennis S. _iatsuhiro
NASA Ames Research Center
ABSTRACT
A Two-Dimensional Oscillating Airfoil Test Apparatus is presented as a
method of measuring unsteady aerodynamic forces on an airfoil or rotor blade
section. The oscillating airfoil test rig, which is being built for use in
NASA Ames Research Center's llxll-foot Transonic Wind Tunnel (speed range M =
0.4 - 1.4), will allow determination of unsteady loadings and detailed pressure
distributions on representative airfoil sections undergoing simulated pitching
and flapping motions. This paper will present the design details of the motion
generating system and supporting structure. This apparatus is now in the con-
struction phase.
INTRODUCTION
ExF)erimental unsteady aerodynamic data are urgently needed to validate and
guide comnutational aerodynamics research in process at Ames, Lanq!ey, and the
Air Force Flight Dynamics Laboratory. Data obtained with this apparatus will
be compared with theoretical work concerning boundary layer effects on aileron
flutter. Boundary layer effects are a major reason for the large uncertainty
and consequent weight penalty associated with current predictions of control
surface or flap flutter at transonic speeds.
A helicopter blade in forward f!iqht is subjected to wide variety of
unsteady loads due to aerodynamic effects. The effects of cyclic pitch, flap-
ping _;otions, and the asymmetry gaused by advancing and retreating blades all
contribute to unsteady aerodynar_ic loads. The unsteady flow is extremely com-
plex because the speed regime may ci,ange from subsonic through transonic to
supersonic in one revolution. The importance of predictinq these forces
becomes especially important at high forward speed where transonic effects on
stability and flutter margins degrade performance. A rational theoretical
study for unsteady aerodynamics is extremely complex; therefore, a concerted
effort was directed by Ames Research Center into the design of a Two-Dimen-
sional Oscillatinq Airfoil Test Apparatus. The test apparatus was especially
designed to apply IDrogra,lrnedpitchinq and heaving motions to the test airfoil
so that different sir,lulatedblade systems could be included if, the test proqraw,_.
BASIC OPERATIO_I
r eThe Two-Di_::enqional _)scillatinq Airfoil ,.st Apparatus (see Fiqure 1) will
") / ' /¢_g<:i11,_o a __,,-IrICil :,l.:_l r:,_'ter') choYd, 54-inch (1.37 meter) span airfoil in
lTI ' ' ._l -_'_
Ames' llxll-foot Transonic Wind Tunnel at frequencies from 0 to 60 HZ. The
test objective will be to oscillate the airfoil at pitch angles to + 2 of
rotation about any point along the chord and also to vertically displace the
airfoil up to + 2-inches. The mean angle of attack of the airfoil will be
variable over the range of -5 _ to +15 _, Befo, e each test, wedqe blocks are
manually installed between the top flexure and wing to obtain the desired angle
of attack. See Figure 3.
The oscillating motions are produced by motion generators which are
hydraulic actuators with push-pull rods connected to the airfoil. Two pairs
of high performance servo-controlled linear-hydraulic actuators induce the
motion. One pair of actuators is for driving the leading edge and the other
pair for driving the trailing edge of the airfoil. Actuator cross-coupling
is provided by the lightweight airfoil structure. The actuators contain two
pistons on a common rod for the dynamic and static loads. Each actuator
drives identical graphite-epoxy push rods and flexure bearings which attach
to the four corners of the test airfoil.
TEST APPARATUS
Motion Generators
The servo-hydraulic actuator package was desiqned and built by M.T.S.
System Corp. {Minneapolis, Minnesota) to Ames' specifications and will be
driven by two 150 HP units rated at 65 GPrl and 3000 PSI. Each actuator
consists of two separate pistons on a sinqle rod enclosed in a dual chamber
cylinder as shown in Figure 2. The upper piston is used for generating
dynamic forces. The lower section is for load biasinq. The load bias system
is used to support the constant aerodynamic load thereby reducinq the size and
power required for the dynamic cylinder. The load biasing circuit includes an
accumulator to maintain a constant load bias force alonq with differential
control. This load biasing section is controlled by a servo-valve system with
a resonance frequency below 0.5 HZ. T_is system a]lows slow dynamic movement
yet maintains static preload. As static bias requirements change, the servo-
valve ports oil into the appropriate end of the cylinder. This essentially
changes the accumulator (nitrogen precharge) and adjusts the static force
output.
A velocity and position transducer are mounted in the center of the
actuator. They combine into a single physical unit with coils and cores in
line axially for placement within the hollow actuator rod.
The displacement or dynamic section is controlled by a high performance
servo-valve coupled thrGugh a manifold into the cylinder. The valve spool lap
is adjusted to achieve flow linearity of better than + 2 to 35 GPI]. The
manifold houses adjustable cross-port relief vaives. To keep the breakaway
friction below 8 pounds force, controlled leakage clearance fits and labyrinth
grooves are used on both piston and end caps.
A hydrau:ic service manifold between the power unit and servo-actuator
provides hydraulic filtration and suppresses line pressure fluctuation in the
172
high electrohydraulic actuator supply and return lines. This service manifold
also houses safety features which include provisions for automatic low pressure
shutdownand solenoid valves to relieve all system pressure.
The wing is mounted_o push-pull rods through flexure bearings. The
airfoil is a lightweight graphite-epoxy structure designed to withstand a
230 Gacceleration and a lO,O00 poundaerodynamic load. See Figure 3. Ex-
tensive development in the fields of engineering, fabrication techniques, and
testing was performed at the AmesResearchCenter in order to obtain an accept-
able graphite-epoxy structure of various configurations. The first natural
bending frequency is above lO0 HZ and the first torsional modeis above 60 HZ.
Forty (40) dynamic pressure transducers and forty (40) static pressure taps
are imbeddedin the skin of the airfoil. Sampleslaminated like the Final
airfoil design had^the _ollowing properties: an ultimate t_nsile stress of
130_KSI (8.96 x IO_ N/M_), a modulqsof elasticity o_ 14xlOu PSI (9.65 x lOlO
N/M_), and a density of 0.6 LBM/IN j (I.66 x IOw KG/MJ).
Support Structure
Included in the test apparatus are splitter plates with trailinq edae
flaps and side struts. The splitter plates are a practical method of support-
ing the two-dimensional model wing in the test section without extensive modi-
fication to the wind tunnel test section. All instrumentation in the model
wing is funneled down through the central part of the splitter plate. Approxi-
mately 130 static pressure orifices are imbedded in the surface of the _plitter
chordwise and exit through the top of the splitter plate. AIso, the splitter
plate serves as a mount for the drag link and stabilizes the model wing in the
lateral direction.
Trailing edge flaps were incorporated to adjust the pressure gradient in
the channel between the splitters. Adjustments of the flaps are continuous,
as opposed to discrete, and are remotely controlled. Angles between + 5 degrees
can be obtained about the center line of the flap.
The side strut is a supporting member to stabilize the splitter plates and
to eliminate excessive deflection in the lateral direction due to aerodynamic
loads imposed on the splitter plates (See Figure l). The side struts are
fastened to the splitter plates and protrude throuqh the tunnel wall to the
exterior tunnel structure.
PRETEST SETUP
A dynamic test will be conducted on the motion qenerators before installa-
tion into the ll-foot wind tunnel. A11 of these component desiqns are pushinq
the state-of-the-art and no previous test information is available to judqe
their worthiness when used in combination. Fiqure 4 illustrates the tost setup.
The test will determine what performance levels can be expected from the
hydraulic system, the push-pull rods, the flexures, and the win_I. The model
,,_ill be !)reloaded by an underside airbaq. The model winq will be oscillated
_ertically, causing a fluctuation in surface loadinq. In addition, the effect
of the cy(;li(L i;)ading on the fle/ures and push-pull rods will be deterI:;ined.
17:_
CONCLUDINGRET_RKS
The Two-DimensionalTest Apparatus was designed to apply proqrammedpitch-
ina and heavinq motions to the test airfoil so that various applications to
blade systems can be included in the test program. The results of these tests
will be used as input parameters for the dynamic analysis of existing rotorcraft
and for checking numerical and analytical schemesfor advancedrotors. This
paper has described an oscillating mechanismfor wind tunnel studies which
should provide the motions necessary for generating this aerodynamic data.
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SIDE STRUT
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WIND TUNNEL TEST
SECTION OUTLINE
/
MODEL WING
AIRFLOW
HYDRAULIC ACTUATORS
SIDE STRUT
DRIVE RODS
SUPPORT COLUMNS
Figur_ 1 WIND TUNNEL INSTALLATION OF TWO DIMENSIONAL
OSCILLATING WING TEST APPARATUS
175
TWO-DIMENSIONAL OSCILLATING AIRFOIL TEST APPARATUS
LOAD BIASING __
ACCUMULATORS
CYLINDER ROD
CYLINDER
DISPLACEMENT SERVO VALVE
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ACCUMULATORS
LOADING BIASING
SERVO VALVE
Figure 2 HYDRAULIC ACTUATOR
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Two-dimensional oscillating airfoil test apparatus

N79-21391
TWO-DIMENSIONAL OSCILLATING AIRFOIL TEST APPARATUS
By Frank L. Gibson, Andrew J. Hocker, Jr.,
and Dennis S. Matsuhiro
NASA Ames Research Center
ABSTRACT
A Two-Dimensional Oscillating Airfoil Test Apparatus is presented as a
method of measuring unsteady aerodynamic forces on an airfoil or rotor blade
section. The oscillating airfoil test rig, which is being built for use in
NASA Ames Research Center's llxll-foot Transonic Wind Tunnel (speed range M =
0.4 - 1.4), will allow determination of unsteady loadings and detailed pressure
distributions on representative airfoil sections undergoing simulated pitching
and flapping motions. This paper will present the design details of the motion
generating system and supporting structure. This apparatus is now in the con-
struction phase.
INTRODUCTION
Experimental unsteady aerodynamic data are urgently needed to validate and
guide computational aerodynamics research in process at Ames, Langley, and the
Air Force Flight Dynamics Laboratory. Data obtained with this apparatus will
be compared with theoretical work concerning boundary layer effects on aileron
flutter. Boundary layer effects are a major reason for the large uncertainty
andconsequent weight penalty associated with current predictions of control
surface or flap flutter at transonic speeds.
A helicopter blade in forward flight is subjected to wide variety of
unsteady loads due to aerodynamic effects. The effects of cyclic pitch, flap-
ping motions, and the asymmetry caused by advancing and retreating blades all
contribute to unsteady aerodynamic loads. The unsteady flow is extremely com-
plex because the speed regime may change from subsonic through transonic to
supersonic in one revolution. The importance of predicting these forces
becomes especially important at high forward speed where transonic effects on
stability and flutter margins degrade performance. A rational theoretical
study for unsteady aerodynamics is extremely complex; therefore, a concerted
effort was directed by Ames Research Center into the design of a Two-Dimen-
sional Oscillating Airfoil Test Apparatus. The test apparatus was especially
designed to apply programmed pitching and heaving motions to the test airfoil
so that different simulated blade systems could be included in the test program.
BASIC OPERATION
The Two-Dimensional Oscillating Airfoil Test Apparatus (see Figure I) will
oscillate a 20-inch (0.51 meter) chord, 54-inch (1.37 meter) span airfoil in
177
Precedingpageblank -
https://ntrs.nasa.gov/search.jsp?R=19790013220 2020-03-20T18:54:39+00:00Z
Ames' llxll-foot Transonic Wind Tunnel at frequencies from 0 to 60 HZ. The
test objective will be to oscillate the airfoil at pitch angles to + 2° of
rotation about any point along the chord and also to vertically displace the
airfoil up to + 2-inches. The mean angle of attack of the airfoil will be
variable over the range of -5 ° to +15 ° . Before each test, wedge blocks are
manually installed between the top flexure and wing to obtain the desired angle
of attack. See Figure 3.
The oscillating motions are produced by motion generators which are
hydraulic actuators with push-pull rods connected to the airfoil. Two pairs
of high performance servo-controlled linear-hydraulic actuators induce the
motion. One pair of actuators is for driving the leading edge and the other
pair for driving the trailing edge of the airfoil. Actuator cross-coupling
is provided by the lightweight airfoil structure. The actuators contain two
pistons on a common rod for the dynamic and static loads. Each actuator
drives identical graphite-epoxy push rods and flexure bearings which attach
to the four corners of the test airfoil.
TEST APPARATUS
Motion Generators
The servo-hydraulic actuator package was designed and built by M.T.S.
System Corp. (Minneapolis, Minnesota) to Ames' specifications and will be
driven by two 150 HP units rated at 65 GPM and 3000 PSI. Each actuator
consists of two separate pistons on a single rod enclosed in a dual chamber
cylinder as shown in Figure 2. The upper piston is used for generating
dynamic forces. The lower section is for load biasing. The load bias system
is used to support the constant aerodynamic load thereby reducing the size and
power required for the dynamic cylinder. The load biasing circuit includes an
accumulator to maintain a constant load bias force along with differential
control. This load biasing section is controlled by a servo-valve system with
a resonance frequency below 0.5 HZ. This system allows slow dynamic movement
yet maintains static preload. As static bias requirements change, the servo-
valve ports oil into the appropriate end of the cylinder. This essentially
changes the accumulator (nitrogen precharge) and adjusts the static force
output.
A velocity and position transducer are mounted in the center of the
actuator. They combine into a single physical unit with coils and cores in
line axially for placement within the hollow actuator rod.
The displacement or dynamic section is controlled by a high performance
servo-valve coupled through a manifold into the cylinder. The valve spool lap
is adjusted to achieve flow linearity of better than + 2% to 35 GPM. The
manifold houses adjustable cross-port relief valves. To keep the breakaway
friction below 8 pounds force, controlled leakage clearance fits and labyrinth
grooves are used on both piston and end caps.
A hydraulic service manifold between the power unit and servo-actuator
provides hydraulic filtration and suppresses line pressure fluctuation in the
178
high electrohydraulic actuator supply and return lines. This service manifold
also houses safety features which include provisions for automatic low pressure
shutdown and solenoid valves to relieve all system pressure.
The wing is mounted to push-pull rods through flexure bearings. The
airfoil is a lightweight graphite-epoxy structure designed to withstand a
230 G acceleration and a I0,000 pound aerodynamic load See Figure 3. Ex-
tensive development in the fields of engineering, fabrication techniques, and
testing was performed at the Ames Research Center in order to obtain an accept-
able graphite-epoxy structure of various configurations. The first natural
bending frequency is above I00 HZ and the first torsional mode is above 60 HZ.
Forty (40) dynamic pressure transducers and forty (40) static pressure taps
are imbedded in the skin of the airfoil. Samples laminated like the final
airfoil design had the _ollowing properties: an ultimate t_nsile stress of
130_KSI (8.96 x IOS N/M_), a modulus of elasticity o_ 14xlO u PSl (9.65 x I0 I0
N/M_), and a density of 0.6 LBM/IN _ (1.66 x I0 KG/M ).
Support Structure
Included in the test apparatus are splitter plates with trailing edge
flaps and side struts. The splitter plates are a practical method of support-
ing the two-dimensional model wing in the test section without extensive modi-
fication to the wind tunnel test section. All instrumentation in the model
wing is funneled down through the central part of the splitter plate. Approxi-
mately 130 static pressure orifices are imbedded in the surface of the splitter
chordwise and exit through the top of the splitter plate. Also, the splitter
plate serves as a mount for the drag link and stabilizes the model wing in the
lateral direction.
Trailing edge flaps were incorporated to adjust the pressure gradient in
the channel between the splitters. Adjustments of the flaps are continuous,
as opposed to discrete, and are remotely controlled. Angles between _ 5 degrees
can be obtained about the center line of the flap.
The side strut is a supporting member to stabilize the splitter plates and
to eliminate excessive deflection in the lateral direction due to aerodynamic
loads imposed on the splitter plates (See Figure I). The side struts are
fastened to the splitter plates and protrude through the tunnel wall to the
exterior tunnel structure.
PRETEST SETUP
A dynamic test will be conducted on the motion generators before installa-
tion into the ll-foot wind tunnel. All of these component designs are pushing
the state-of-the-art and no previous test information is available to judge
their worthiness when used in combination. Figure 4 illustrates the test setup.
The test will determine what performance levels can be expected from the
hydraulic system, the push-pull rods, the flexures, and the wing. The model
will be preloaded by an underside airbag. The model wing will be oscillated
vertically, causing a fluctuation in surface loading. In addition, the effect
of the cyclic loading on the flexures and push-pull rods will be determined.
179
CONCLUDING REMARKS
The Two-Dimensional Test Apparatus was designed to apply programmed pitch-
ingand heaving motions to the test airfoil so that various applications to
blade systems can be included in the test program. The results of these tests
will be used as input parameters for the dynamic analysis of existing rotorcraft
and for checking numerical and analytical schemes for advanced rotors. This
paper has described an oscillating mechanism for wind tunnel studies which
should provide the motions necessary for generating this aerodynamic data.
180
ORIGINAL PAGE IS
OF. POOR QUALITY
TWO-DIMENSIONAL OSCILLATING AIRFOIL TEST APPARATUS
WIND TUNNEL TEST
SIDE STRUT
MODEL WING
AIRFLOW_
SPLITTER PLATES
HYDRAULIC ACTUATORS
SECTION OUTLINE
I _, SPLITTER FLAP
Z
DRIVE RODS
__-_ SUPPORT COLUMNS
.--.x
Figure 1 WIND TUNNEL INSTALLATION OF TWO DIMENSIONAL
OSCI LLATING WING TEST APPARATUS
181
TWO-DIMENSIONAL OSCILLATING AIRFOIL TEST APPARATUS
CYLINDER ROD
LOAD BIASING CYLINDER
ACCUMULATORS •
DISPLACEMENT SERVO VALVE
ACCUM U LATO RS
LOADING BIASING
SERVO VALVE
Figure 2 HYDRAULIC ACTUATOR
182
TWO-DIMENSIONAL OSCILLATING AIRFOIL TEST APPARATUS
4'"
CLAMP BOLTS
Oo
¢0
ANGLE OF ATTACK WEDGE
BLOCKS FORE AND AFT
FLEXURE
PRESSURE TRANSDUCER
Figure 3 MODEL WING
GO
LATERAL STABLIZER
MODEL WING
DRAG LINK
TWO-DIMENSIONAL OSCILLATING AIRFOIL TEST APPARATUS
DRAG LOAD SIMULATOR
FLEXURE
AFT DRIVE RODS
I
SUPPORT BASE
• TEST STAND
HYDRAULIC ACTUATORS
FORE AND AFT
Figure 4 DYNAMIC TEST STAND


Two-dimensional oscillating airfoil test apparatus

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