structural and mechanism design of an active trailing-edge flap blade

Journal of Mechanical Science and Technology 27 (9) (2013) 2605~2617

www.springerlink.com/content/1738-494x

DOI 10.1007/s12206-013-0704-8

Structural and mechanism design of an active trailing-edge flap blade†

Jae Hwan Lee1, Balakumaran Natarajan2, Won Jong Eun2, Viswamurthy S. R.3, Jae-Sang Park4, Taesong Kim5 and Sang Joon Shin2,*

1Samsung Techwin R&D Center, Bundang-gu, Seongnam-si, Gyeonggi-do, 463-400, Korea 2School of Mechanical and Aerospace Engineering, Institute of Advanced Aerospace Technology, Seoul National University, Seoul, 151-742, Korea

3Advanced Composites Division, National Aerospace Laboratories (CSIR), Bangalore, India 44th R&D Institute, Warhead/Material, Agency for Defense Development, Yuseong-gu, Deajeon, Korea

5Department of Wind Energy, Technical University of Denmark, Risoe Campus, Roskilde, Denmark

(Manuscript Received August 15, 2012; Revised January 10, 2013; Accepted March 29, 2013)

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Abstract

A conventional rotor control system restricted at 1/rev frequency component is unable to vary the hub vibratory loads and the

aeroacoustic noise, which exist in high frequency components. Various active rotor control methodologies have been examined in the

literature to alleviate the problem of excessive hub vibratory loads and noise. The active control device manipulates the blade pitch angle

with arbitrary higher harmonic frequencies individually. In this paper, an active trailing-edge flap blade, which is one of the active control

methods, is developed to reduce vibratory loads and noise of the rotor through modification of unsteady aerodynamic loads. Piezoelectric

actuators installed inside the blade manipulate the motion of the trailing edge flap. The proposed blade rotates at higher speed and addi-

tional structures are included to support the actuators and the flap. This improves the design, as the blade is able to withstand increased

centrifugal force. The cross-section of the active blade is designed first. A stress/strain recovery analysis is then conducted to verify its

structural integrity. A one-dimensional beam analysis is also carried out to assist with the construction of the fan diagram. To select the

actuator and design the flap actuation region, the flap hinge moment is estimated via a CFD analysis. To obtain the desired flap deflection

of ±4°, three actuators are required. The design of the flap actuation region is validated using a test bed with a skin hinge. However, be-

cause the skin hinge induces additional flap hinge moment, it does not provide sufficient deflection angle. Therefore, the flap hinge is

replaced by a pin-type hinge, and the results are evaluated.

Keywords: Active trailing-edge flap; Piezoelectric actuator; Active rotor blade design; Structural integrity

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1. Introduction

Helicopters generally have a wide range of mission capa-

bilities such as vertical flight and hovering. However, due to

their rotor system, helicopters typically operate in fairly com-

plex aerodynamic environments compared to a fixed wing

aircraft, since the flow speed is different between the retreat-

ing and advancing sides of the rotor disk during forward flight

(Fig. 1). In addition, wake vortices from the preceding blade

impinge upon the following blade in a process widely known

as blade-vortex interaction (BVI) in low speed flight.

The unsteady aerodynamic environment induces high vibra-

tory loads and critical aeroacoustic noise. Hence, helicopters

exhibit many shortcomings, such as a poor ride quality, a re-

stricted flight envelope, and increased maintenance costs, due

to frequent replacement of components. The frequencies of

the primary hub vibratory loads in the rotor with N blades are

N times the rotor blade revolution frequency (N/rev). Conven-

tional rotor control systems composed of swashplates and

pitch links are incapable of adjusting such a vibratory load

because the rotor control is restricted to 1/rev in those systems.

Both passive and active approaches have attempted to re-

duce the vibratory load and noise. Passive control approaches

use simple additional components on the rotor, for example a

bifilar vibration absorber or a rotor head absorber. However,

these methods lead to an increase in the gross weight of the

rotorcraft. They are also associated with a lack of adaptability

with respect to changes in the dynamic characteristics caused

by cargo, fuel and passenger loads [1]. Therefore, several ac-

tive rotor control approaches have been investigated. Active

control methods manipulate the blade pitch with a higher

harmonic displacement based on either conventional or ad-

vanced materials. In particular, a considerable amount of re-

search using piezoelectric materials for vibration and noise

reduction has been conducted [2-5].

Several active rotor control methods have been widely stud-

*Corresponding author. Tel.: +82 2 880 1642, Fax.: +82 2 887 2662

E-mail address: [email protected] † Recommended by Editor Yeon June Kang

© KSME & Springer 2013

2606 J. H. Lee et al. / Journal of Mechanical Science and Technology 27 (9) (2013) 2605~2617

ied, such as the active twist rotor (ATR) blade [6, 7] and the

active trailing-edge flap (ATF) blade [8, 9]. There has also

been a hybrid type that uses two more methods simultane-

ously [10-12]. The ATF method uses a trailing edge flap to

change the effective camber of the airfoil. Because the ATF

method adopts smaller hardware, in this case a trailing-edge

flap, compared to other active rotor control methods, relatively

low electric voltage input is required to generate the desired

flap deflection. The active fiber composites embedded in an

ATR blade generally require much higher input voltages (ap-

proximately 1,000 volts) compared to that required by a pie-

zostack actuator. According to DYMORE [13], the ATF

method results in good performance in vibratory loads reduc-

tion with low input voltage [14]. When a 3/rev flap control

input was used for a four-blade rotor, approximately 90%

reduction in hub vibratory loads was obtained. SMART active

flap rotor was tested in the 40- by 80-foot anechoic wind tun-

nel at NASA Ames Research Center [15]. The results showed

reductions up to 6dB in BVI noise. In addition, its hub vibra-

tory loads were alleviated by approximately 80%. Flap deflec-

tion of ±4° was required to reduce the vibratory loads effec-

tively according to a simulation. The ATF method also has

other advantages, such as high control authority and less intru-

siveness regarding the structural integrity of the blade. On the

other hand, the proposed active rotor blade is capable of exe-

cuting primary helicopter control as a substitute for the

swashplate system [16]. Thus, the ATF method is used in this

paper. The present active blade is termed the Seoul National

University Flap (SNUF) blade. Several other active rotor

blades using the ATF method have been developed, such as

the smart material actuated rotor technology (SMART) active

flap rotor [8, 9], the active blade concept (ABC) rotor blade

[17], and the smart hybrid active rotor control system

(SHARCS) rotor blade [10-12]. The SMART active flap rotor

is a full-scale blade that uses X-frame actuators [18]. The

ABC and SHARCS rotor blades are small-scale rotor blades

that use the amplified piezoelectric actuator (APA) supplied

by Cedrat Ltd. Also, there is a piezoelectric bender [19-21]. A

flap actuation mechanism using a piezoelectric bender is rela-

tively simple compared to the use of a piezostack actuator, as

a piezostack actuator requires a specific mechanism that am-

plifies its resulting displacement. However, it is weak against

fatigue, as its displacement is generated by repetitive bending,

and also its block force is insufficient to actuate the flap.

Therefore, most of ATF blades under recent development use

the piezo-stack type actuator with an amplification mechanism.

We selected a piezostack actuator, specifically the APA ac-

tuator, for the SNUF blade to drive the flap. Since the APA

actuator has been used for other ATF rotor blades, its per-

formance is considered to be verified. Even for the preferred

piezo-stack type actuation, rather complicated designs resulted

in order to avoid the possible actuation performance degrada-

tion due to the undesirable contact and friction taking place

among the inner components. Therefore, a heedful effort is

required to simplify the design of the active blade and the

inner components. At the same time, a more robust design is

required that is capable of minimizing the possible perform-

ance degradation in the flap deflection. It is also required to

guarantee the structural integrity of the active blade and the

components. SNUF blade is to be developed to satisfy those

specific requirements.

This paper presents the detailed design of the SNUF blade

to alleviate primarily the N/rev hub vibratory loads in the hub

vertical direction, by using available open- and closed-loop

control algorithms. However, a practical active rotor may en-

counter various factors which will deteriorate the vibration re-

duction performance. Thus, by going through prototype SNUF

blade design, fabrication, and testing procedure, those practical

factors will be identified and improvement will be devised.

Even though such deteriorating factors are expected to exist in

the practical hardware, the aim is that at least 50% of vibratory

loads reduction will be alleviated in the foregoing prototype

SNUF rotor system in a typical forward flight condition. One-

dimensional beam analysis and a two-dimensional cross-

sectional analysis will be conducted separately to design and

analyze SNUF blade. A stress/strain recovery analysis will also

be conducted to investigate the structural integrity of the blade.

The local stress/strain predicted for each layer in the blade skin

will be verified as to whether failure will occur or not.

2. Design requirement

The SNUF blade is a small-scale rotor blade based on the

design of the existing SHARCS and NASA/Army/MIT ATR

blades. Table 1 summarizes the general properties of the ATR

and SHARCS rotor blades.

The ATR blade showed satisfactory performance in vibra-

tion reduction during the test at the NASA Langley wind tun-

nel [6] in a heavy gas environment. Therefore, its structural

Fig. 1. Aerodynamic environment of the rotor disk.

J. H. Lee et al. / Journal of Mechanical Science and Technology 27 (9) (2013) 2605~2617 2607

properties were selected as the target properties in the design of

the present SNUF blade. However, these properties need to be

modified to be suitable for the present testing condition. Be-

cause the SNUF blade will rotate under normal atmospheric

condition, its rotating speed must be higher than that of the ATR

blade in a heavy and dense medium. To maintain the tip Mach

number (0.6) of the ATR blade, the nominal rotation speed of

the SNUF blade is determined as 1,528 RPM. Due to such dif-

ference, the resulting design for the present SNUF blade will

vary from that of ATR. The following similarities are consid-

ered to obtain the design requirements of the SNUF blade.

(1) Dynamic similarity.

(2) Mach number similarity.

(3) Lock number similarity.

The design requirements of the SNUF blade are described

in Table 2. For simplicity, a uniform cross-section along the

blade span and a partially linear built-in twist distribution were

considered for the present SNUF blade. Flap deflection of ±4°

as an angle requirement was established according to previous

vibratory load reduction prediction results based on the ATF

method [14]. All feathering, flapping, and lead-lag hinges are

located at 4% radius position.

3. Preliminary design

3.1 Blade design

The cross-sectional design of the SNUF blade is based on

that of the previous ATR blade. Therefore, its cross-section is

designed to be fabricated with a similar spar configuration

with the same composite materials used in the ATR blade.

The present blade design uses a two-dimensional cross-section

analysis and a one-dimensional beam analysis. The present

cross-sectional design uses a two-cell thin-walled beam sec-

tion analysis [22] based on the variational asymptotic method.

The two-cell thin-walled beam section analysis is capable of

estimating the properties of the cross-section of the beam. A

stress/strain recovery analysis was also feasible due to the use

of the two-cell thin-walled beam section analysis. The result-

ing cross-sectional configuration is described in Fig. 2.

The cross-section property of the present design is described

in Table 3. A few ballast weights were added to the present

cross-sectional design to meet the target weight requirement

and to locate the center of gravity as close to 25% chord as

possible. Detailed illustration for these ballast weights will be

included in section 4.2, where it will be denoted as the ballast

Table 1. Properties of the ATR and SHARCS rotor blades.

ATR blade SHARCS Blade

Control method Active twist Hybrid

Rotor radius (cm) 139.7 109.6

Rotation speed (rpm) 687.5 1,555

Blade chord (mm) 108 80

Airfoil type NACA 0012 NACA 0015

Tip Mach number 0.603 0.52

Lock number 4.55 5

Table 2. Design requirements of the SNUF blade.

Property Value

Rotor type Articulated

Rotor radius, R (cm) 128

Rotation speed (rpm) 1,528 (160 rad/s)

Blade chord, c (cm) 10.24

Hinge offset (cm) 5.12

Root cutout (% span) 20

Airfoil type NACA0012

Tip Mach number 0.60

Lock number 5.0

Mass per unit length (kg/m) 0.55

Pretwist (deg) -10

GJ (N-m2) 68

EIflap (N-m2) 57

EIlag (N-m2) 1900

EA (N) 4.6×106

Iyy (kg-m) 2.7×10-5

Izz (kg-m) 2.25×10-4

Ipolar (kg-m) 2.52×10-4

Flap displacement ±4˚

Table 3. Comparison of the cross-section properties.

Target

properties

Present

results

Difference

(%)

EA (N) 4.600×106 4.421×106 -7.0

GJ (N-m2) 6.800×101 6.792×101 -0.1

EIflap (N-m2) 5.700×101 5.828×101 2.5

EIlag (N-m2) 1.900×103 2.736×103 29.0

Mass (kg/m) 5.500×10-1 5.502×10-1 0.0

Iyy (lagwise moment of

inertia, kg-m) 2.700×10-5 1.638×10-5 -39.3

Izz (flapwise moment of

inertia, kg-m) 2.250×10-4 1.596×10-4 -29.0

Fig. 2. Preliminary cross-section configuration of the SNUF blade.


weight distribution for the non-active region. Although the

present design attempted to match the target sectional proper-

ties via a manually iterative process as closely as possible,

there are moderate differences between the present result and

the target properties.

The detailed mass distribution will be determined after the

stress/strain recovery analysis is completed. Since there are

several components inside the blade, a concentrated load will

be induced by centrifugal force in the flap actuation region.

The leading edge and nose are further reinforced to withstand

the centrifugal force. Because the dynamic responses in the

torsion and flap bending modes must be satisfactory, the de-

sign effort needs to be conducted via a manually iterative

process to minimize the differences.

The SNUF blade uses a constant chord and partially linear

built-in twist distribution along the blade span. The baseline

built-in twist angle of the blade is -10°. Since it was found to

be difficult to fabricate a blade with a twist in the flap actua-

tion region, a built-in twist was not applied in that region as

seen in Fig. 3.

The actuators and other relevant components are located in-

side the blade. Therefore, a specially designed housing and

hatch is used in the region, as shown in Fig. 4.

It will be more beneficial for the aerodynamic force control

when the flap is located at a distance further outboard of the

blade. However, this will also be dangerous for the flap struc-

tures, as the centrifugal force acting on it will be increased as

well. Thus, the location of the flap is presently determined to

be at 75% of the radius of the blade according to the result of

the SHARCS blade. In addition, the span length (20% of the

blade radius) and chord length (15% of the blade total chord

length) of the present flap dimension are selected based on

those of SHARCS rotor blade. The present flap actuation re-

gion was designed to have a similar configuration with that in

SHARCS. Its detailed design will be modified to accommo-

date inner components such as the actuators, housing, and

push rod. Structural weakness may occur, and thus, a metallic

housing is used to remedy this problem. Because it is not pos-

sible for two-cell thin-walled beam section analysis to analyze

the flap actuation region in a detailed level, a three-

dimensional NASTRAN structural analysis will be conducted.

In that effort, shell, rigid bodies, and three-dimensional solid

elements will be used to represent the blade structural compo-

nents, such as skin, spar, actuator mechanism, and ballast

weight. Thus, in that case, a separate cross-sectional analysis

will not be required. In addition, the sectional properties in the

blade root region are assumed to have the similar properties of

the ATR blade root.

3.2 Actuator selection

The actuator needs to provide sufficient stroke and block

force to the flap under a specified input voltage. Furthermore,

it must exhibit high-frequency response characteristics to ma-

nipulate the flap at a higher harmonic frequency under a high

rotating speed condition. Therefore, piezostack was consid-

ered as the actuator. Specifically, the APA actuator was se-

lected. There are several other types of APA actuators avail-

able for the present purpose. Given that the SNUF blade is a

small-scale blade, its inner space is not sufficient. Therefore,

the size of actuators was the first consideration in the selection

of an actuator. Among the family of APA actuators, APA 200

M [29] was adopted (Fig. 5). The properties of the APA 200

M actuator are summarized in Table 4.

Its size is small enough to be installed inside the present

SNUF blade, but its width is wider than the available space,

which is a maximum of 7 mm. Therefore, it was decreased

specifically to 5 mm by the manufacturer according to the

requirements of the SNUF blade. However, the block force of

a single actuator may not be sufficient to deflect the flap under

the rotating condition. Detailed discussion on this incapability

of a single actuator will be given in section 3.4 along with the

selection of moment arm length in the flap actuation mecha-

nism. Therefore, multiple actuators are considered to enable the

required flap deflection. Again, detailed explanation regarding

the number of actuators will be included in section 3.4.

Fig. 3. Built-in twist angle distribution of the SNUF blade.

Fig. 4. External configuration of the SNUF blade.


3.3 Estimation of the hinge moment

To design a linkage mechanism between the actuator and

flap, the hinge moment should be predicted. The block force

and translation motion of the piezostack actuator are trans-

ferred as the torque and rotational motion in the flap actuation

region. Therefore, the flap hinge moment must be accurately

predicted to determine the length of the required moment arm.

The hinge moment, Mtot, is obtained by adding the aerodynamic,

Maero, inertial, Minertial, and centrifugal moments, Mcentrifugal, as

follows:

Mtot=Maero+Minertial+Mcentrifugal.. (1)

Generally, the hinge moment generated by the aerodynamic

force, Maero, is much larger than the other terms [10]. There-

fore, only the aerodynamic hinge moment is estimated here

using the following two methodologies. First, it is predicted

by using an analytical methodology for a hover condition, and

then by using FLUENT, a commercial CFD (computational

fluid dynamics) analysis program for the forward flight condi-

tion. The model used in the present estimation contains only

two-dimensional airfoil and flap.

Between the two methodologies, an analytical estimation is

first carried out for hover condition only. It is done by adopt-

ing the formulas suggested by Lee [23], under the assumptions

of uniform inflow and untwisted blade using blade ele-

ment/momentum theory. The deflection of the trailing edge

flap, δ, is considered through a change in the local effective

angle of attack over the portion of the blade with flap. For a

plain flap, the local lift for a blade with a deflected flap can be

written as

21( )

2ll r cCρ= Ω (2)

where

00 .l l

RC C

rα

α λθ δ

δ λ∆ Ω

= + − ∆ (3)

The value of inflow ratio λ can be obtained by substituting

the thrust coefficient in terms of λ in the above equation and

solving the quadratic equation. The hinge moment co-efficient

and hinge moment per unit span can be determined by using

h hh l

l

dC dCC C

dC dδ

δ

= +

(4)

2 21( ) .

2h fh C r cρ= Ω (5)

Here, the values of dCh/dCl, dCh/dδ and ∆α0/∆s can be found

as approximated functions of only geometric parameter cf /c.

By substituting Cl in Eq. (4) from Eq. (3) and finally into Ch in

Eq. (5), the hinge moment per unit span can be obtained. The

final hinge moment can be obtained by integrating Eq. (5)

over the finite flap span. In this estimation, the collective pitch

angles of SNUF rotor are varied to be 0o, 5

o, and 10

o for a

constant flap deflection angle of 4o. And the resulting hinge

moments are summarized as shown in Table 5.

As a second attempt of prediction, a CFD analysis is con-

ducted using FLUENT for forward flight, under the condition

parameters described in Table 6. NACA 0012 airfoil, which

contains flap, is used to represent the 75% spanwise section of

SNUF blade. Then, a static two-dimensional CFD analysis is

conducted under a condition that the blade section momentar-

ily passes through the advancing side of the rotor in forward

flight at advance ratio µ = 0.3. It is thus anticipated that the

Table 4. APA 200M actuator properties.

Properties Unit APA200M

Displacement (µm) 230

Blocked force (N) 73

Stiffness (N/µm) 0.32

Resonance frequency (free-free) (Hz) 4600

Resonance time (free-free) (ms) 0.11

Resonance frequency (blocked-free) (Hz) 900

Resonance time (blocked-free) (ms) 0.56

Voltage range (V) -20 ~ 150

Capacitance (µF) 3.2

Resolution (nm) 2.3

Thermo-mechanical behavior (µm /˚K) 2.72

Height H (in the actuation direction) (mm) 17.0

Width (incl. edges, wires) (mm) 9.0

Length (mm) 55.00

Mass (g) 15.7

Fig. 5. APA 200 M actuator.

Table 5. Hinge moment predicted for the SNUF blade.

Angle of attack

(°)

CFD

(forward flight)

Collective

pitch (°)

Analytical

(hover)

0 0.0896 0 0.0538

5 0.1074 5 0.0652

10 0.1268 10 0.0843


maximum values of the aerodynamic hinge moment acting

upon the flap will be obtained.

The hinge moments are obtained with regard to three differ-

ent static airfoil angles of attack, i.e., 0o, 5

o and 10

o, with a

fixed flap deflection angle, 4o. Results of both the analytical

prediction and the CFD analysis are summarized in Table 5.

The critical hinge moment is estimated to be 0.1268 N·m

amongst the several predictions obtained from both analyses,

although the maximum target flap deflection angle is selected

to be an alternating value, ±4o, in the foregoing design of

SNUF blade.

3.4 Flap actuation region design

It is quite important to design the flap actuation region care-

fully so as to obtain sufficient deflection of the flap. If the

weight of the flap actuation components becomes possibly

smaller, it will be beneficial to the structural integrity of the

whole blade, since lower centrifugal loads will act on it. One

of the best ways to do this is to fabricate this region by using a

composite material. However, the difficulties related to this

type of fabrication due to the small space still need to be con-

sidered. Therefore, aluminum is selected to be a material for

the housing and fairing block. However, it is determined that

the push rod is to be made of steel to prevent buckling. One of

the most important components is the flap hinge, since friction

force is expected to occur due to the deformation of the flap

during the rotation. A skin hinge is considered initially as a

candidate to minimize the friction force. However, since the

rotational speed is quite high and only a relatively short area

exists where the flap is attached to the blade, this type of hinge

is incapable of withstanding the loads acting on the hinge.

Concentrated loads may act upon the hinge, making it difficult

for the skin hinge to maintain the desired flap deflection. This

incapability arises because the skin hinge has too small an area

to support the flap. Thus, a conventional pin hinge is consid-

ered in the second phase, and it is fabricated out of steel. A

schematic of the linkage mechanism is illustrated in Fig. 6.

An appropriate length of the moment arm (L2) needs to be

chosen to deflect the flap with the desired angle against the

externally applied flap hinge moment. The block force of a

single APA 200M actuator itself is 73N. However, its

free/maximum stroke is 0.23 mm, according to the product

specification (Fig. 7(a)). This signifies incapability of the sin-

Table 6. Condition parameters used in the CFD simulation.

Flap location 75%R

Flap length 20%R

Flap chord 15%c

Advance ratio (µ) 0.30

V (advancing side) 215 m/s

Mach number 0.63 (sea level)

Flap deflection (δ) 4 ˚

Fluid density (ρ) 1.225 kg/m3 (sea level)

Temperature 300 ˚K

Flow condition Incompressible, inviscid

Fig. 6. Schematic of the linkage mechanism.

(a) Force displacement relationship property of the APA 200M

actuator

(b) Relation curve between flap deflection angle and length of flap

hinge moment arm when dual actuators are used

(c) Relation curve when three actuators are used

Fig. 7. Block force diagram of the APA 200M actuator and relation

curve between flap deflection angle and moment arm.


gle APA 200M actuator to produce the required flap deflec-

tion. Thus, multiple actuators are considered to overcome such

incapability. Then, the relationship between L2 and the num-

ber of the actuators is not usually linear. The appropriate value

of L2 and an appropriate number of actuators are determined

by examining such nonlinear relationship between the flap

deflection and the moment arm length. This prediction is per-

formed via the following manual iterative process. As a first

attempt, the relationship curve corresponding to the dual ac-

tuators is illustrated in Fig. 7(b). According to this curve, the

maximum flap deflection angle will be ±2.7° when dual actua-

tors are used. The optimum length of the hinge moment arm is

about 1.2 mm. However, this is insufficient compared to the

target value of ±4°. Therefore, three actuators are considered

to provide the required flap hinge moment. The resulting rela-

tionship curve for three actuators is shown in Fig. 7(c).

According to Fig. 7, the maximum flap deflection will be

about ±3.4° and the optimum length of moment arm is about

0.9 mm. Although the aerodynamic effect is not considered,

this flap deflection is still insufficient compared to the target

value. However, with more than three actuators, the space

inside the blade to accommodate these actuators will be insuf-

ficient. Furthermore additional structural problems could en-

sue due to the increased weight. Therefore, three actuators

configuration is selected.

To locate the actuator and linkage mechanism inside the

blade and facilitate the access to these inner components, a

hatch will be included on the flap actuation region. The hatch,

however, may cause degradation in structural integrity, espe-

cially near the flap actuation region. Thus the present housing,

which was originally designed to protect the inner actuator

components, also plays a role of preventing such degradation

in structural integrity. The flap actuation region design with

three actuators is shown in Fig. 8.

4. Final design

4.1 Blade redesign

The cross-sectional design of the SNUF blade is intended to

satisfy the design requirements established through the use of

materials similar to those in the ATR blade. However, those

specific materials are not available domestically. Different raw

materials are required to fabricate the present blade. Therefore,

the cross-section of SNUF blade is redesigned using materials

that are actually available. The properties and strength charac-

teristics of the new materials are summarized in Tables 7(a)

and 7(b), respectively.

In the original design of the ATR blade, the front spar has

two different regions of the ply lay-up to accommodate the

piezoelectric fiber actuators. However, such a complicated

design of the front spar is not required in the present blade

design. Under the presently simplified design of the front spar,

both the flapwise bending and axial stiffness are increased.

Axial stiffness is required to overcome the increased centrifu-

gal force, and this was gained by a redesign. However, the

Table 7. (a) Properties of the new materials in SNUF blade design; (b)

Strength characteristics of the new materials in SNUF blade design.

Carbon UD Carbon fabric Glass fabric

Thickness (µm) 140 230 120

Density (kg/m3) 1531 1466 1684

E11 (GPa) 159.6 58.4 20.3

E22 (GPa) 8.1 60.39 21.2

ν12 0.34 0.04 0.13

G12 (GPa) 4.6 4.07 3.03

(a)

Allowable strain

(micro strain) Carbon UD

Carbon

fabric

Glass

fabric

Tension 16,540 13,298 13,592 Fiber

Compression 16,540 13,298 13,592

Tension 5,253 13,443 12,143 Transverse

Compression 5,253 13,443 12,143

Shear 13,110 16,609 16,435

(b)

(a) Design of the flap linkage mechanism

(b) Design of the flap actuation region

(c) Assembly design of the blade and flap actuation region.

Fig. 8. Preliminary design of the flap actuation region.


flapwise bending stiffness is also increased in this case. The

revised design result is shown in Fig. 9 and Table 8.

Ballast weights are applied in the nose region in the cross-

section to place the center of gravity at a 25% chord location.

Thus, the center of gravity was modified to be at the 25%

chordwise location. However, less ballast weight, 0.13 kg/m,

than needed was applied onto the upper and lower surface skin.

Therefore, the sectional lagwise moment of inertia, Iyy, be-

comes smaller than the target value. Those masses were lo-

cated on 43.1% chordwise position and inside ±0.6% of the

maximum thickness in the airfoil (Fig. 10). More ballast

weight is needed at the trailing edge to match the required

value of Izz. However, it was found to be difficult to place the

center of gravity at the 25% chord location while matching the

target blade section mass. Therefore, the final value for the

sectional flapwise moment of inertia, Izz, value became

smaller than the target value. To locate the center of gravity at

the 25%, a ballast weight of 0.197 kg/m was applied in the

3.6% chordwise location (Fig. 10). Since both axial and tor-

sional stiffness influence the structural integrity and torsional

response of the rotor blade, the present design will be kept and

used in spite of the significant discrepancies between the de-

sign results and the target properties. The flap actuation region

design was also revised so that it would be compatible with

the present blade design.

4.2 Stress/strain recovery analysis

The aerodynamic configuration of the SNUF blade, such as

airfoil shape, built-in twist, is similar to that of the ATR blade.

However, for the present SNUF blade with a higher rotational

speed under the normal atmosphere, a precise prediction of

structural loads for the SNUF blade will be required. A com-

prehensive rotorcraft analysis program, CAMRAD II [24], is

used to predict the magnitude of the loads acting on the inter-

nal blade structure. The rotor is trimmed by using a wind tun-

nel trim condition for a specified forward flight condition: the

advance ratio µ = 0.3, the thrust CT/σ = 0.08, the rotor shaft

tilting angle αs = -2o. The rolled-up single tip free wake model

is used. The plots of the flap bending, lead-lag bending, and

torsional moments against the azimuth angle for the various

radial locations are shown in Figs. 11(a)-(c), respectively.

Due to a higher rotational speed and relatively heavy com-

ponents inside the flap actuation region, considerable concen-

trated loads may act on the flap actuation region. Therefore, a

careful stress/strain recovery analysis is required to evaluate

the integrity of the structure. The centrifugal force is obtained

by a one-dimensional geometrically exact beam analysis [25].

The largest magnitudes of the structural loads predicted from

CAMRAD II analysis as indicated in Fig. 11 are extracted and

combined with the centrifugal loads from 1-D beam analysis,

additionally with a safety factor of 1.5. After these structural

loads are obtained, the cross-section design described in Table

9 is used for the present stress/strain analysis using the maxi-

mum strain criterion. It is assumed in Table 9 that the stiffness

properties of the flap actuation region are the same as those

corresponding to the section without the actuators, although

few variations exist in the structural configuration. Detailed

distinction in the stiffness properties between two sections,

with and without flap actuators, will be possible when the

three-dimensional analysis is conducted. Only variation of the

inertial properties will be considered due to addition of the

actuation mechanism.

This specific property of the cross-sectional design reflects

the detailed weight information of the various components

included in the flap actuation region. The weight of inner

components, involving the actuator, linkage mechanism, and

housing, is assumed as 0.8kg/m based on the drawing shown

in Fig. 8. At the current stage, mass of the whole components

was regarded as two concentrated point masses for simplicity.

Table 8. Comparison of the cross-section properties of the revised

design.

Target

properties

Design

result

Difference

(%)

EA (N) 4.600×106 4.569×106 -0.7

GJ (N-m2) 6.800×101 6.810×101 0.1

EIflap (N-m2) 5.700×101 9.892×101 73.5

EIlag (N-m2) 1.900×103 2.237×103 17.7

Mass (kg/m) 5.500×10-1 5.500×10-1 0

Iyy (lagwise moment of

inertia, kg-m) 2.700×10-5 6.252×10-6 -76.8

Izz (flapwise moment of

inertia, kg-m) 2.250×10-4 2.939×10-4 -30.6

Fig. 9. Revised configuration of the cross-section.

Fig. 10. Ballast weight distribution for the non-active region.


To locate the center of gravity at 25% chordwise position, a

ballast weight of 0.76 kg/m was applied. Finally, revised bal-

last weights corresponding to the active region consist of the

following two parts: 0.76 kg/m at 3.6% chordwise and two

masses of each 0.4 kg/m at 43.1% chordwise locations (Fig.

12).

In contrast to it, ballast weight distribution designed for the

remaining non-active region is already explained in section

3.1 and illustrated in Fig. 10. Thus, the present inertial proper-

ties became slightly heavier than the target properties. A

stress/strain recovery analysis was conducted at ten locations

on the spanwise reference line via a two-cell thin-walled

analysis. Details of the ten spanwise locations considered for

the analysis are described in Fig. 13(a).

Among those ten locations, the locations corresponding to

10% and 100% of the radius are not considered in the analysis.

The centrifugal and structural loads used in the present analy-

sis are summarized in Tables 10(a) and 10(b), and the result-

ing stress/strain recovery analysis is summarized in Table 11.

Both the maximum stress and the maximum strain occur at the

30% radius location, as illustrated in Fig. 13(b). Numerical

Table 9. Cross-section properties of the flap actuation region.

Design results

EA (N) 4.569×106

GJ (N-m2) 6.810×101

EIflap (N-m2) 9.892×101

EIlag (N-m2) 2.237×103

Mass (kg/m) 17.81×10-1

Iyy (lagwise moment of inertia, kg-m) 15.30×10-6

Izz (flapwise moment of inertia, kg-m) 7.945×10-4

(a) Flapwise bending moment vs. azimuth angle

(b) Lagwise bending moment vs. azimuth angle

(c) Torsional moment vs. azimuth angle

Fig. 11. Structural moments vs. azimuth angle for various radial locations.

Table 10. (a) Stress-strain recovery analysis loads (upto 50%R); (b)

Stress-strain recovery analysis loads (60%R ~ 90%R).

20%R 30%R 40%R 50%R

Centrifugal force (N) 2.1020 1.8265 1.7457 1.6431

-2.8 -2.5 -2.0 -1.7

-1.4 -1.2 -1.1 -0.98 Torsional moment

(N-m)

-3.2 -2.9 -2.5 -2.1

-4.7 0.97 -1.1 -1.2

-3.1 3.5 1.4 1.6 Flapwise bending moment

(N-m)

-4.7 1.3 -0.85 -0.65

-78.0 -58.4 -47.7 -43.6

-89.6 -67.1 -55.0 -48.9 Lagwise bending moment

(N-m)

-84.1 -62.7 -50.5 -45.3

(a)

60%R 70%R 80%R 90%R

Centrifugal force (N) 1.5164 1.1632 0.6016 0.2191

-1.4 -1.0 -0.67 -0.40

-0.92 -0.82 -0.60 -0.33 Torsional moment

(N-m)

-1.8 -1.4 -1.0 -0.59

-1.2 -0.70 0.33 0.87

2.0 2.1 1.7 0.82 Flapwise bending moment

(N-m)

-0.12 0.48 1.1 1.1

-42.4 -40.1 -35.2 -22.5

-45.7 -42.2 -35.4 -22.3 Lagwise bending moment

(N-m)

-43.1 -40.7 -34.9 -22.3

(b)

Fig. 12. Ballast weight distribution for the active region.


results indicate that the structures is unlikely to fail, as the

maximum strain levels in all of the components were found to

be lower than the allowable strain levels.

4.3 Fan diagram

A fan plot analysis was conducted to investigate the dy-

namic characteristics of the SNUF blade during rotation, as

shown in Fig. 14.

The fan diagram is constructed after a one-dimensional

geometrically exact beam analysis. A comparison plot of natu-

ral frequencies of SNUF and SHARCS rotor blades at the

nominal rotation speed is shown in Fig. 15. Although the pre-

sent SNUF blade is designed based on the NASA/Army/MIT

ATR blade properties, the rotation speed of the SNUF blade

rather becomes similar to that of the SHARCS rotor blade.

Thus the present frequencies are compared with those of the

SHARCS rotor blade. The natural frequency corresponding to

the first flapwise bending mode is higher than that of the

SHARCS rotor blade because the present blade is designed to

have greater flapwise bending stiffness than the target stiffness.

The natural frequency corresponding to the first torsional

mode is lower than that of the SHARCS blade, the value of

which is 4.3 /rev. It is anticipated that such a decrease in the

torsional mode frequency can induce aeroelastic instability.

All of the other natural frequencies were found to be satisfac-

tory when compared to the natural frequencies of the

SHARCS blade, except for the second flapwise mode. Al-

though the second flapwise bending mode intersects with ap-

proximately 5/rev excitation frequency at 100% RPM, it will

not become a serious problem since the present rotor system

will be under the primary 4/rev excitation frequency. A de-

tailed aeroelastic stability analysis will be conducted in future.

4.4 Flap actuation region design improvement

In the preliminary design, a pin-type hinge is selected to

hold and move the flap, as this design is simple and is capable

of withstanding severe loads, as already shown in Fig. 8. Fur-

ther improvement of the design based on the original concept

in Fig. 8 was done, as illustrated in Fig. 16.

The push rod, which transfers the force from the actuators

to the flap, was reinforced to prevent possible buckling due to

compressive force. The flap hinge is located at a 10.2%

chordwise position from the trailing edge. A test bed based on

the improved design is fabricated as shown in Fig. 17. The test

Table 11. Stress/strain analysis results.

Component Max. strain

(micro strain)

Ratio between the existing maximum

strain and the allowable

Longitudinal 11,116.8 0.54

Transverse 3,314.9 0.40

Shear -2,418.4 0.18

(a) Spanwise location where the maximum strain occurs

(b) Location of the maximum strain within the cross section

Fig. 13. Spanwise locations and test condition for the stress/strain

recovery analysis and the location of the resulting maximum stress.

Fig. 14. Fan diagram of the SNUF blade.

Fig. 15. Comparison plot of SNUF and SHARCS blade natural fre-

quencies.

Fig. 16. Revised design of the flap actuation component.


bed accommodates three APA 200M actuators, housing, full-

size flap, push rods and rear block.

Flap deflection is being measured using a laser-

displacement sensor and data acquisition hardware by Na-

tional Instruments (Fig. 18). The APA actuator is operated

under the voltage between 0 ~ 150 Volts. The flap actuation

test is now being conducted to identify any undesirable factors

to degrade the flap deflection. At the same time, more im-

provements in order to increase the flap deflection will be

suggested and added in the final design. Such further im-

proved design will be implemented in the prototype blade.

The current characterization test upon the test bed will be

completed in the next few months.

5. Conclusions

A further simplified and improved design was attempted for

an active rotor blade with a trailing edge flap. For its design,

cross-sectional design and structural analysis were performed.

Two-cell thin-walled beam section analysis was used for the

present cross-section design. Maximum strain criterion was

used to verify the structural integrity under the action of the

centrifugal force and structural loads obtained from

CAMRAD II analysis. According to the stress/strain failure

analysis, the present design showed satisfactory result in terms

of the structural integrity. However, a much more detailed

three-dimensional finite element analysis will be conducted to

confirm the inner components included within the flap actua-

tion region. The dynamic characteristics were examined

through a fan plot analysis. Resonance was not expected to

arise during the rotation of the blade at a nominal speed ac-

cording to the proposed fan diagram. The preliminary blade

cross-sectional design was modified to enable fabrication us-

ing available materials in domestic. An aerodynamic analysis

was conducted to estimate the hinge moment via a CFD pre-

diction as well as a simple analytical solution. According to

this analysis, the hinge moment generated by the aerodynamic

force can be predicted as 0.1269 N-m. The APA200M actua-

tor was selected to deflect the trailing-edge flap and the de-

tailed design of the flap actuation region was conducted. A

test bed was fabricated that contained the actuators, housing,

push rods, full size flap, and rear block. Flap deflection meas-

urement is being conducted, and will be used to identify the

resulting deflection in the near future. This preliminary test

will help to identify any possible factors which degrade the

actuation performance, and also help to improve the design.

Such improvement will be reflected in the prototype blade

fabrication in the near future.

Acknowledgment

This work was supported by the Korea Science and Engi-

neering Foundation (KOSEF), through a grant funded by the

Korean government (MOST) (No.2009-0075614). This work

has been supported in part by Defense Acquisition Program

Administration (DAPA) and Agency for Defense Develop-

ment (ADD).

Nomenclature------------------------------------------------------------------------

α0 : Zero-lift angle of attack

αs : Rotor shaft tilting angle

c : Chord length of airfoil

cf : Chord length of flap

Ch : Hinge moment co-efficient

Cl : Lift co-efficient of airfoil

Clα : Lift curve slope of airfoil

CT : Thrust co-efficient

δ : Flap deflection angle

E : Young’s modulus

EA : Axial stiffness

EIflap : Flapwise (transverse bending) stiffness

EIlag : Lagwise (in-plane bending) stiffness

G : Torsion modulus

GJ : Torsion stiffness

H : Hinge moment

Ipolar : Polar moment of inertia

Iyy, Izz : Principal moments of inertia of blade cross-section

l : Lift per unit span

L,L1,L2 : Lengths

λ : Inflow ratio

M : Moment

µ : Advance ratio

ν : Poisson’s ratio

Fig. 17. Test bed fabricated, based on the improved design.

Fig. 18. Flap deflection measurement upon the test bed using the laser

sensor.


Ω : Rotational velocity

ρ : Fluid density

r : Radius of blade section

R : Rotor radius

σ : Solidity ratio

V : Velocity

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Jae Hwan Lee received his B.S. in

Aeronautical Engineering from Sejong

University, Korea in 2008. He received

the M.S. from Seoul National University

in 2011. He is a researcher in CAE En-

gineering Division, Samsung Techwin

Corp. in the Advanced Technology Cen-

ter. His research interests include multi

body dynamics and finite element analysis.


Balakumaran Natarajan received his

B.S. in Aeronautical Engineering from

MIT, Anna University, India in 2007.

He worked with Helicopter Division,

HAL in the Design Liaison Engineering

Dept. during 2007-2011. He received

his M.S. from Seoul National University

in 2013. His research interests include

finite element analysis, wind turbine and helicopter rotor blade

design.

Won Jong Eun received his B.S. in

Mechanical and Aerospace Engineering,

from Seoul National University, Seoul,

Korea in 2012. Mr. Eun is currently in

an M.S. course at the Active Aeroelas-

ticity and Rotorcraft Laboratory, Seoul

National University. His research inter-

ests include multi body dynamics, rotor-

craft dynamics, and structural dynamics.

S. R. Viswamurthy received his B.S. in

Aerospace Engineering from Indian

Institute of Technology Madras, India in

2001. He completed his doctoral thesis

in the area of Helicopter Aeroelasticity

& Vibration control from the Indian

Institute of Science, Bangalore, India in

2007. Presently, he is a Scientist at the

National Aerospace Laboratories (CSIR), Bangalore, India.

His research interests include composite materials, finite ele-

ment analysis, helicopter aeroelasticity and vibration control.

Jae-Sang Park received the Ph.D. in

Mechanical and Aerospace Engineering,

from Seoul National University, Seoul,

Korea in 2006. Dr. Park is currently a

researcher in Agency for Defense De-

velopment, Daejeon, Korea. His re-

search interests include aeroelasticity,

rotorcraft aeromechanics using multi-

body dynamics, and smart structures.

Taeseong Kim received his B.S. in

Aerospace Engineering from Korea

Aviation University, Korea, in 2004. He

then received his M.S. and Ph.D. de-

grees from Seoul National University in

2006 and 2009, respectively. Dr. Kim is

currently a Research Scientist at Wind

Energy Division at Risø National Labo-

ratory for Sustainable Energy in Roskilde, Denmark. His re-

search interests include aeroelasticity, wind turbine dynamics,

rotorcraft dynamics, and structural dynamics.

Sang Joon Shin received S.M. and

Ph.D. degrees in Aeronautics and As-

tronautics from Massachusetts Institute

of Technology in 1999 and 2001, re-

spectively. Since 2003, he has been a

professor at the School of Mechanical

and Aerospace Engineering in Seoul

National University. His research inter-

ests include aeroelasticty, rotorcraft dynamics, and smart

structures.

structural and mechanism design of an active trailing-edge flap blade

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