A DLM-BASED MSC Nastran AERODYNAMIC FLUTTER SIMULATOR FOR AIRCRAFT …pages.mscsoftware.com/rs/mscsoftware/images/AM-CO… · · 2018-04-19A DLM-BASED MSC Nastran AERODYNAMIC FLUTTER

MSC Software Confidential MSC Software Confidential

A DLM-BASED MSC Nastran AERODYNAMIC

FLUTTER SIMULATOR FOR AIRCRAFT LIFTING

SURFACES 2013 Regional User Conference

Paper No. AM-CONF13-34

Presented By: Emil Suciu

L-3 Communications Platform Integration Division, Waco, Texas

May 7, 2013

THE T-TAIL FLUTTER MECHANISM

REVISITED

Paper No. IFASD-2011-121

IFASD-2011, Paris, France, June 26-30, 2011

by Emil Suciu1, Nicholas Stathopoulos2,

Martin Dickinson2 and John Glaser3

1L-3 Communications, Platform Integration Division, Waco, Texas 2Bombardier Aerospace 400 Cote-Vertu Road West Dorval, Quebec, Canada H4S 1Y9 3Bombardier Aerospace (Retired)

mailto:[email protected]

MSC Software Confidential

SUMMARY

A DLM-based aerodynamic simulator for flutter is used to

identify some of the most important aerodynamic drivers for

the T-Tail flutter mechanism of a complete aircraft. The

simulator is using the Modal Descrambling Factoring

Method, which permits individual variations of each direct

and each interference aerodynamic force, moment and

hinge moment independently of any other force or moment.

The sensitivity of the flutter solution to individual variations

of very large numbers of direct and interference

aerodynamic derivatives can be studied with ease.


LINEAR FLUTTER ANALYSIS vs.

REALITY; SOME POSSIBLE OUTCOMES* 1. Analysis predicts flutter and it is there and at the right

speed; desired outcome

2. Analysis predicts flutter but it is not there; can cause headaches and unnecessary work

3. Analysis predicts flutter and it is there but at the wrong speed; can cause headaches and unnecessary work

4. ANALYSIS DOES NOT PREDICT FLUTTER AND IT IS THERE; A MOST UNDESIRABLE OUTCOME

*This assumes that we have some experimental or Navier-Stokes guidance


WHY USE AERODYNAMIC FACTORING IN

FLUTTER ANALYSES?

• Theodorsen & Garrick, as far back as 1940 recognized that 2-D strip theory, especially when applied to full span wings, needs to be corrected for finite span effects, viscosity and compressibility

• The Doublet Lattice Method (DLM) is a 3-D linear theory (SUPERPOSITION APPLIES) which does not account for viscosity, shocks, vortices and lifting surfaces correct relative positions (interference), etc.

• FAA Advisory Circular AC No. 25.629.1A recognizes that intersecting surfaces pose special aerodynamic problems and recommends that intersecting lifting surfaces in-plane and interference effects be included in flutter analyses


WHY USE AERODYNAMIC FACTORING?

continued

THERE ARE NUMEROUS OPPORTUNITIES TO

MESS UP FLUTTER ANALYSES; LACK OF

AERODYNAMIC FACTORING CAN BE ONE SUCH

OPPORTUNITY


The T-46A Airplane Wing-Aileron-Tab Flutter Incident and

Post-Incident Analyses with Factoring (Ref. J. Aircraft Paper, May 1988,

by French, R.M., Noll, T.,Cooley, D.E., Moore, R. and Zapata, F.)

A Case When Analysis Did Not Predict Flutter And It Occurred in

Flight; Aerodynamic Factoring Was Not Used on Aileron and Tab

Bill Rodden used to hand out this paper at conferences

and recommend factoring to every flutter analyst


Aerodynamic Factoring, continued

Direct quotation from the Giesing, Kalman and Rodden 1976 Report on Correction Factor Techniques for the DLM. On pressure factoring:

“One set of correction factors, determined from one mode, to other modes has not met with much success. Specifically, correction factors obtained using a pitch mode cannot be applied to pressures due to control surface deflections. The converse is also true.”*

*This conclusion is as valid today as it was in 1976.


Aerodynamic Factoring, continued

HOWEVER, if we can apply the correction factors obtained from a k=0.0 pitch mode to a k≠0.0 pitch (or torsion) mode ONLY, from a k=0.0 control surface deflection mode to a k≠0.0 control surface deflection mode ONLY, from a tab k=0.0 deflection mode to a tab k≠0.0 deflection mode ONLY, etc., it would appear that we could meet with more success, as anticipated by Rodden in 1976.

It turns out that through MODAL DESCRAMBLING it is possible to replace any general mode of vibration with 5 simple and well ordered modes to which the approppriate factors can be applied.


The problem is that the flutter analyst does not encounter the modes

of vibration of an aircraft in a predictable sequence (pure bending,

then pure torsion, then pure control surface rotation and torsion, then

pure tab rotation and torsion and finally pure elastic streamwise

camber deformation); the average mode of vibration is scrambled

and it looks forbidding, as below:


USING MODAL DESCRAMBLING, FOR ANY LIFTING SURFACE

ARRANGEMENT, AT ANY AERODYNAMIC STRIP,

EVERY GENERAL MODE OF

VIBRATION LOOKS THE SAME! (1) pure bending +

(2) pure torsion +

(3) pure control surface rotation and torsion +

(4) pure tab rotation and torsion +

(5) pure elastic streamwise camber deformation


WE CAN THEREFORE FACTOR EXACTLY

EACH OF THE DESCRAMBLED PURE MODES

EXCEPT FOR THE ELASTIC STREAMWISE

CAMBER DEFORMATION MODE, FOR WHICH

CORRECTION DATA IS GENERALLY NOT

AVAILABLE


Schematic of Descrambling Process of General

Wing-Control Surface-Tab Mode Shape at the

Aerodynamic Surface


There Is More to Aerodynamic Factoring Than

Only Modal Descrambling at the Aerodynamic

Surface; There Is Also Aerodynamic

Interference From Surface to Surface to

Consider


Aerodynamic Forces on Lifting Surfaces: Direct,

Interference and Stacked Forces


Aerodynamic Forces: Direct, Interference

and Stacked Forces (continued) For an arrangement of n lifting surfaces in the same

interference group, each surface or component

experiences 1 (one) direct set of forces and moments and

n-1 sets of interference forces and moments;

For n>2,there are more interference forces

than direct forces on any lifting surface!

Most (if not all) factoring schemes apply direct factors

ONLY TO THE STACKED FORCES


α >0°

α >0° DLM model; only slopes for deflection

H.S. incidence can be controlled separately

Wing – Flap - Horizontal Stabilizer Interference Problem

at α >0°; Reality vs. DLM model (Capabilities & Limitations)

Reality: relative location of wing-flap-tail and their vortices

δF >0°

δF > 0°


Wing – Flap –Horizontal Stabilizer Interference Problem

Calculated with Program ILSA (a MORINO METHOD)


The Original Aerodynamic Derivatives (now

Modal Descrambling) Factoring Method Developed at De Havilland Aircraft of Canada (now

Bombardier) in 1987, then a division of the Boeing Company; method in use at DHC (Q300, Q200, Q400) and Boeing (on 777 aircraft)

This method accounts for surface-on-itself interference but does not treat lifting surface-to-lifting surface interference correctly, same as other factoring methods; no camber

Q300 Q400


The Modal Descrambling Factoring Method

The modal descrambling factoring scheme descrambles all modal

motion and separates interference forces from direct forces and

permits separate factoring

Let’s say we divide the aircraft lifting surfaces into 4 components: (1)

wing, (2) H.S., (3) V. Fin, (4) engines + pylons + ventral fins

THEN: surface 1 moves only in bending; calculate all direct and

interference forces on all the components due to surface 1 bending;

THEN: surface 1 moves only in torsion; calculate all direct and

interference forces on all the components due to surface 1 torsion;

THEN: only the control surfaces of surface 1 rotate; etc.


Factoring Direct and Interference

Aerodynamic Forces at any Aerodynamic

Strip of Any Component

Each aerodynamic strip has 64 factors if ncomp=4;

for 209 strips, for one Mach Number, we need a total

of 13376 factors


Airplane Models Analyzed

Twin engine T-Tail airplane with H.S. anhedral

Manual control surfaces

3 stiffness levels of vertical fin are analyzed

Flutter analyses at one transonic Mach Number

The Modal Descrambling Aerodynamic Factoring Method is Used with the DLM to identify the important aerodynamic drivers of the T-Tail flutter mechanism


Figure 1. Scrambled Vertical Fin Bending/Torsion/HS Roll/Elevator

Rotation/Rudder Rotation Mode at the DLM Aerodynamic Surface

with a Few DESCRAMBLED AND SEPARATED Direct and

Interference Aerodynamic Forces and Moments on the H.S.

ChEδE

ChEh

ChE

ChEβ

ChEδR

Clh

Cl

ClhVF

Clβ

ClδR

CYVFβ

CYVFδR

ChRβ

ChRδR

Blue= DIRECT FORCE

Red = INTERFERENCE FORCE


Horizontal Stabilizer DLM Symmetric Derivatives

Calculation

Three Elementary Unit Rigid Mode Shapes for Calculating

Horizontal Stabilizer - Elevator Symmetric Aerodynamic Derivatives with the DLM

(1) Rigid Heave (2) Rigid Pitch

(3) Rigid Elevator Rotation

The Direct Clα Factor Is Derived from Here; Most

Factoring Schemes Only Factor the Direct Clα on the

H.S.

(3) Rigid Elevator Rotation

MSC Software Confidential BOMBARDIER PRIVILEDGE AND CONFIDENTIAL

Empennage DLM Antisymmetric Derivatives Calculation

1. Rigid Side-to-Side Motion 2. Rigid Yaw

Four Elementary Unit Rigid Mode Shapes for Calculating Fin Aerodynamic Derivatives with the

DLM; Interference to Horizontal Stabilizer is Automatic

3. Rigid Rudder Rotation 4. Rigid Tab Rotation


Figure 5. Typical DLM vs. CFD Horizontal

Stabilizer Spanwise Clβ Distribution. k=0.000.

Figure 6. Typical CLα and CLβ Total Factors vs.

Mach Number for the Horizontal Stabilizer;

k=0.000

Rigid Unit Vertical Fin Yaw

Typical Horizontal Stabilizer CLα and CLβ Factors vs.

Mach Number; k=0.000

Mach Number M

Aer

od

ynam

ic F

acto

rs

CLα Factor vs. M

CLβ Factor vs. M

Analysis CLβ Factor

Analysis CLα Factor

1.000

CLα Factor<1.000

CLβ Factor>1.000

CLβ Factor<1.000

DLM vs. CFD Clβ Distribution on Horizontal Stabilizer; k=0.000

-1.0 -0.8 -0.6 -0.4 -0.2 0.0 0.2 0.4 0.6 0.8 1.0

semispan, η

Cn*c

/cav

e

CFD

DLM


Figure 3. Real Parts of Unfactored Descrambled

Direct and Interference Cn on Horizontal Stabilizer

of the T-Tail for General Mode of Vibration Shown

in Figure 1; k=0.700.

Figure 4. Imaginary Parts of Unfactored

Descrambled Direct and Interference Cn on

Horizontal Stabilizer of the T-Tail for General Mode

of Vibration Shown in Figure 1; k=0.700.

DESCRAMBLED UNFACTORED DIRECT AND INTERFERENCE REAL Cn ON

HORIZONTAL STABILIZER OF T-TAIL FOR GENERAL (SCRAMBLED) MODE OF

VIBRATION; k=0.700

-0.010

-0.008

-0.006

-0.004

-0.002

0.000

0.002

0.004

0.006

0.008

0 0.2 0.4 0.6 0.8 1

semispan η

Rea

l Cn

CN DUE TO HS ROLL/BENDING

CN DUE TO HS TORSION

CN DUE TO ELEVATOR ROTATION

CN DUE TO HS CAMBER

CN DUE TO V. FIN LAT BENDING

CN DUE TO V. FIN TORSION

CN DUE TO RUDDER ROTATION

CN DUE TO RUDDER TAB

ROTATION

CN DUE TO V. FIN CAMBER

DESCRAMBLED UNFACTORED DIRECT AND INTERFERENCE IMAGINARY Cn ON

HORIZONTAL STABILIZER OF T-TAIL FOR GENERAL (SCRAMBLED) MODE OF

VIBRATION; k=0.700

-0.005

0.000

0.005

0.010

0.015

0.020

0.025

0.030

0 0.2 0.4 0.6 0.8 1

semispan η

Imag

inar

y C

n

CN DUE TO HS ROLL/BENDING

CN DUE TO HS TORSION

CN DUE TO ELEVATOR ROTATION

CN DUE TO HS CAMBER

CN DUE TO V. FIN LAT BENDING

CN DUE TO V. FIN TORSION

CN DUE TO RUDDER ROTATION

CN DUE TO RUDDER TAB

ROTATION

CN DUE TO V. FIN CAMBER


Effect of Aerodynamic Factoring on T-Tail Flutter Speed.

%(Vf-VfRef)/VfRef vs. Vertical Fin Stiffnesses; All Speeds in KEAS

0

20

40

60

80

100

120

140

160

180

1.000 1.125 1.250 1.375 1.500

Vertical Fin EI & GJ Stiffness Ratios

%(V

f-VfR

ef)/

VfR

ef

No Factoring

Nominal Factoring

Clβ factor=Clα factor

ChEh'*0.500

ChE(h+α)'*0.500

ChEδE'*1.2

ClδR Theoretical

ChEδR'*0.500

ChEβ'*1.2

Figure 9. Effect of Vertical Fin Stiffness Variations and of Aerodynamic Factoring

Variations, One Derivative at a Time from Nominal on Flutter Speed of T-Tailed

Aircraft. Reference Speed Is for Nominal Factoring, Model 1.

Typical Horizontal Stabilizer CLα and CLβ Factors vs.

Mach Number; k=0.000

Mach Number M

Aer

od

ynam

ic F

acto

rs

CLα Factor vs. M

CLβ Factor vs. M

Analysis CLβ Factor

Analysis CLα Factor

1.000

CLα Factor<1.000

CLβ Factor>1.000

CLβ Factor<1.000

MOST ANALYSES ARE HERE

ALL SHOULD BE HERE

SOME ARE HERE


Figure 10. V-g and V-f Plots of Model No. 1 T-Tail Flutter Solution for

Nominal Aerodynamic Factoring; Clβ factors > Clα factors

Figure 11. V-g and V-f Plots of Model No. 1 T-Tail Flutter Solution for

Nominal Aerodynamic Factoring; Clβ factors = Clα factors.

Modal Damping; T-Tail Aircraft Model #1; Nominal Aerodynamics Factoring;

Clβ factors > Clα factors

-0.30

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

V (KEAS)/Vmax

da

mp

ing

, gVertical Fin Bending

Vertical Fin Torsion

Rudder Rotation

Modal Frequency; T-Tail Aircraft Model #1; Nominal Aerodynamics Factoring;

Clβ factors > Clα factors

0

5

10

15

20

25

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

V (KEAS)/Vmax

Fre

qu

en

cy

(H

z)

Vertical Fin Bending Vertical Fin Torsion

Rudder Rotation

Modal Damping; T-Tail Aircraft Model #1; Nominal Aerodynamics Factoring;

Clβ factors = Clα factors

-0.30

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

V (KEAS)/Vmax

da

mp

ing

, g

Vertical Fin Bending

Vertical Fin TorsionRudder Rotation

Modal Frequency; T-Tail Aircraft Model #1; Nominal Aerodynamics Factoring;

Clβ factors = Clα factors

0

5

10

15

20

25

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

V (KEAS)/Vmax

Fre

qu

en

cy

(H

z)

Vertical Fin Bending Vertical Fin Torsion

Rudder Rotation


A Summary of the Aerodynamic Forces and their Phases on the Horizontal

Stabilizer Affecting T-Tail Flutter; No Control Surfaces; Positive Yaw Is Assumed.


For an airplane having 2600 aerodynamic boxes, for

109 modes, the MSC Nastran DLM complete flutter

solution with 8 k-values takes approximately 30 minutes

on a desktop PC with 8GB RAM.

For the same configuration divided into 4 groups, the

general aerodynamic factoring program labors for

approximately 9 minutes; it uses the stored AICs

calculated by the DLM, but it processes 20 downwash

vectors/mode (4X5) vs. 1/mode for the DLM; LSP3G

mode processing is much more extensive than DLM; no

attempt so far to optimize LSP3G.


CONCLUSIONS

The unsteady aerodynamic forces on the horizontal

stabilizer drive the T-Tail flutter mechanism

The interference Clβ factor is the most important

aerodynamic driver for the antisymmetric T-Tail flutter

mechanism

The direct ChEδE, ChEα, ChEh and interference ChEδR and

ClδR also are important derivatives drivers


CONCLUSIONS, continued

Variations of the symmetric direct Clα factor have no

measurable effect on the antisymmetric T-Tail flutter

mechanism;

The antisymmetric T-Tail flutter mechanism exhibits a lot of

sensitivity with the EI & GJ values of the vertical fin (well

known)


CONCLUSIONS, continued The use of the modal descrambling factoring method will

also benefit the flutter and gust analyses of:

Wing-low mounted horizontal stabilizer flutter with

differently factored direct and interference

aerodynamics

Wing-engine nacelle flutter with differently factored

direct and interference aerodynamics

Gust Loads analyses with differently factored direct

and interference aerodynamics


• The Modal Descrambling Factoring Method allows the

user unprecedented access to and individual control

of all direct and all interference aerodynamic forces,

moments, control surface and tab hinge moments at

any lifting surface of an aircraft; in effect a flutter

simulator

CONCLUSIONS, the end


Part of 1995 effort to calculate T-Tail flutter speed

reduction with horizontal stabilizer upload; first

proposal to separate and factor differently Clα and

Clβ


G-II T-Tail Wind Tunnel-

Measured Flutter Boundary

MD=0.90 (20% Margin)

GV T-Tail DLM-Calculated Flutter Boundary

MD=0.97 (15% Margin); factors on Clβ = Clα



Documents

A DLM-BASED MSC Nastran AERODYNAMIC FLUTTER SIMULATOR FOR AIRCRAFT …pages.mscsoftware.com/rs/mscsoftware/images/AM-CO… · · 2018-04-19A DLM-BASED MSC Nastran AERODYNAMIC FLUTTER