Structural and dynamic analyses are presented for a shroudless, hollow titanium fan blade proposed for future use in aircraft turbine engines. The blade was modeled and analyzed using the composite blade structural analysis computer program (COBSTRAN); an integrated program consisting of mesh generators, composite mechanics codes, NASTRAN, and pre- and post-processors. Vibration and impact analyses are presented. The vibration analysis was conducted with COBSTRAN. Results show the effect of the centrifugal force field on frequencies, twist, and blade camber. Bird impact analysis was performed with the multi-mode blade impact computer program. This program uses the geometric model and modal analysis from the COBSTRAN vibration analysis to determine the gross impact response of the fan blades to bird strikes. The structural performance of this blade is also compared to a blade of similar design but with composite in-lays on the outer surface. Results show that the composite in-lays can be selected (designed) to substantially modify the mechanical performance of the shroudless, hollow fan blade

Aiello, R. A.

Chamis, C. C.

Hirschbein, M. S.

English

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1A Tachnical Memorandum 82816
(N ASA-TM- 8281 6 )	 ST iUCTUitAL LYNAMICS CF
at1R000LESS, t1ULLUW FAN ELALL-" WITH CUMFUSTIE
IN-LAYS (NA6A)	 12 p LiC A0.2/ME A01 CSCL 21E
N82-22266
Uucaas
63/07 09636
Structural Dynamics of Shroudless, Hollow
Fan Blades With Composite In-Lays
R. A. Aiello, M. S. Hirschbein and C. C. Charms
Lewis Research Center
Cleveland, Ohio
Prepared for the
Twenty-seventh Annual International Gas Turbine Conference
sponsored by the American Society of Mechanical Engineers
London, England, April 18-22, 1982
A
APR
Y
STRUCTURAL DYNAMICS OF SHROUDLESS, HOLLOW FAN BLADES
WITH COMPOSITE IN-LAYS
by
R. A. Aiello, M. S. Hirschbein, and C. C. Chamis
National Aeronautics and Space Administration
Lewis Research Center
Cleveland, Ohio, U.S.A.
ABSTRACT
M
Structural and dynamic analyses are presented for a shroudless, hol-
low titanium fan blade proposed for future use in aircraft turbine en-
gines. The blade was modeled and analyzed using the Lewis Research Cen-
ter's Composite Blade Structural Analysis Computer Program (COBSTRAN).
COBSTRAN is an integrated program consisting of mesh generators, compo-
site mechanics codes, NASTRAN, and pre- and post-processors. Vibration
and impact analyses are presented. The vibration analysis was conducted
with COBSTRAN. Results show the effect of the centrifugal force field on
frequencies, twist, and blade camber. Bird impact analysis was performed
with the Multi-Mode Blade Impact Computer Program. This program uses the
geometric model and modal analysis from the COBSTRAN vibration analysis
to determine the gross impact response of the fan blades to bird strikes.
The structural performance of this blade is also compared to a blade of
similar design but with composite in-lays on the outer surface. Results
show that the composite in-lays can be selected (designed) to substan-
tially modify the mechanical performance of the shroudless, hollow fan
blade.
INTRODUCTION
The shroudless, hollow titanium fan blade design incorporates ad-
vanced fabrication, aerodynamic and structural concepts. The improved
fan efficiency resulting from this advanced design represents up to a
2 percent efficiency gain which is a significant portion of the total
performance improvement goal (about 12 percent) for an energy efficient
engine. Therefore, this analysis was made to evaluate the present design
and to aid in determining the risks involved using this type of blade as
well as to demonstra'.e improvements in structural performance by using
composite in-lays.
ANALYSIS METHODS
NASA Lewis Research Center's Composite Blade Structural Analysis com-
puter code (COBSTRAN) was used for the structural and dynamic analysis of
the fan blades. This code models a blade composed of various layers of
materials as a solid blade having the same structural properties as the
layered blade. This form of modeling simplifies the analysis of blades
with composite in-lays on the surface or internal voids. For the blade
with compositie in-lays the fiber direction is parallel to the span. The
hollow blade was modeled such that the solid sections were layers of ti-
tanium and the hollow sections were simulated by layers of a soft, low
density artificial material. Compared to modeling the hollow part as a
shell, this modeling approach: (1) predicts the same structural response;
(2) has about 25 percent fewer elements; (3) uses plate bending elements
throughout; (4) simplifies rib modeling; and (5) uses the mean camber
line for element connections. Equivalent homogeneous section properties
were calculated at each node from the nodal thickness and material prop-
erties. A finite element mesh of triangular elements and their proper-
ties was generated in a format compatible with NASTRAN input data re-
quirements.
A modified form of the Interactive Multi-Mode Blade Impact (MMBI)
computer code, developed under contract for NASA (ref. 2) was used to
model a bird strike. This code models the bird as a composite of 2-D and
3-D fluid jets. The blade is modeled by a set of normal modes. During
the impact event the program calculates the pressure loading on the face
of the blade 3ccountiog for instantaneous blade position, including blade
motion, and curvature of the blade chord. The blade response is then
calculated incrementally using modal synthesis over each time step.
A computer code (ZULU) was written specifically to convert the blade
profile data to a form compatible with the input requirements of COBSTRAN
(ref. 1)
FINITE ELEMENT MODEL
The external profile and internal cavity details of the blade were
obtained from a preliminary design. Blade stations selected for the ana-
lysis are from a radius of 15.3 in. (0.389 m) at the root of the blade
to a radius of 43.2 in. (1.10 m) at the tip or a span of about 30 in.
(0.762 m). The blade had an aspect ratio of 2.5, a hub-tip ratio of
0.34, a weighted average pressure ratio of 1.7 and a tip speed of about
1500 ft/;ec (457 m/s) at design speed. The weight of the blade modeled
by NASTRAN is 18.2 lb (8.27 kg).
The finite element model was generated by establishing node points
along the mean camber line of each of 34 blade stations for a total of
433 nodes. These nodes were connected by 777 triangular plate bending
elements. The thickness of each ?lement was obtained by averaging the
blade thickness at each of its nodes. A computer plot of the finite ele-
ment model used for the NASTRAN analysis is shown in figure 1.
All constraint to blade motion was at the nine nodes of station 15.3
(blade root). Eight of the nodes were constrained in five degrees of
freedom (DOF) but allowed to translate in the direction of the axis of
rotation (chord-wise at the root). The middle node was constrained in
all six degrees of freedom. The blad. , as modeled contains 2552 degrees
of freedom which is considered to be more than adequate for this struc-
tural analysis.
ANALYSIS AND RESULTS
All-Titanium Blade
Vibration modes of the blade were calculated under a no-load condi-
tion and also at various rotational speeds to include the effect of cen-
trifugal stiffening. The frequencies determined for the all-titanium
.._i^ow blade are summarized in table I. These results were obtained by
using constraints along the blade root as described previously.
The frequencies in parentheses in table I were obtained by con-
straining the b'iade alon g the platform streamline in the same manner as
for the blade root. Constraint along this streamline implies a rigid
locking o` the blade in the rotor. Therefore these results can be con-
sidered c.nly as an upper bound of the frequency range.
The effects of centrifugal force stiffening on the frequencies are
customarily presented as a plot of frequency vs. rotor speed (Campbell
diagram) as shown in figure 2 in which are also shown the integral order
resonances at the intersection with the engine order lines. The corre-
sponding mode shapes are shown in figure 3.
Tie twist of the blade is reduced as the centrifugal force is in-
creased. The reduction in blade twist along the span at 4000 rpm is
shown in figure 4. The maximum change in twist is about 5 degrees at the
tip. This reduction in twist (untwist) is substantial and, .herefore,
its effect on blade performance should be carefully evaluated. This is
especially critical since the pressure ratio is very sensitive to small
changes in the incidence angle.
The change in blade camber along the span at 4000 rpm is shown in
figure 5. Camber angle decreases in the solid sections of the blade and
increases in the hollow sections. The camber also decreases at the tip.
The maximum change in camber is about a 0.6 degree increase midway
through the hollow span of the blade. This change in camber is signifi-
cant compared to acceptable standards for a shrouded blade.
An indicator of possible blade torsional flutter at minimum cruise,
used in preliminary design, is the "reduced-velocity parameter" defineo
as:
Vr = w
b ft
where
b = 112 chord at 5/6 span (ft)
w = average air velocity relative to the blade o^,er
the outer 1/3 of the span (ft/sec)
ft
 = first torsional frequency at design rpm
(rad/sec)
Experience has shown that, for this case, a value of V greater
than about 1.4 indicates possible blade torsional flutter w6 n operated
at minimum cruise speed. The values of the three terms determined for
this blade at aircraft cruise conditions are:
b = 0.56 ft, w = 1400 ft/sec, f t = 1670 rad/sec
(0.171 m)	 (427 m/s)
The calculated value of the "reduced-velocity parameter" is 1.5.
Based on this parameter a possible flutter problem exists in the 3rd
(first torsional) mode.
A comparison of the bird strike tolerance was modeled by considering
the impact from a 1 lb (0.455 kg) fluid sphere with an impact speed of
900 ft/sec (274 m/s). The diameter of the simulated bird was 3.75 in.
(9.73 cm). The impact point was near the 70 percent span position, and
3
I
i
the missile first contacted the blade about 2 in. (5.08 cm) back from the
leading edge. The impact angle was 25 degrees relative to the chord
line. The first 6 modes at 4000 rpm, obtained using COBSTRAN, were used
to synthesize the blade response to impact using the MMBI code.
The radial stress in the all-titanium blade due to impact is shown
in figure 6. The response is for a point near the stacking axis on the
impacted side of the blade about 1 in. (2.54 cm) above the base of the
blade. This region of the blade carries most of the steady and impact
loading. The impact event lasted about 0.7 millisec. The peak stress is
about 110,000 psi (758 MPa). The stress response at other points near
the base show the same basic time history with the peak stress being
slightly higher near the leading edge. The maximum steady root stress at
the reference point is 41,000 psi (282 MPa). Based on a yield strength
of 125,000 psi (861 MPa) for titanium, the current blade design appar-
ently will yield under the combined stress of the bird strike and the
steady state stress (151,000 vs. 175,000 psi) at the reference point.
It should be noted that t t... mnact analysis considered in this paper
is inherently limited to small .,rformation and is not intended for a de-
tailed analysis of local or severe impact damage. However, it can be
used to assess the gross loading near the blade root. The tolerance at
the impact region, however, needs to be assessed by a detailed stress
analysis which accounts for nonlinear geometry and material effects.
Titanium Blade with Composite In-Lays
A modified hollow titanium blade with an outer surface of inlaid
boron/aluminum composite was considered as a possible improvement to the
basic all-titanium blade. The modified blade has the same dimensions as
the basic blade except that three layers of 0.007 in. (0.0178 cm) thick
unidirectional boron/aluminum composite have replaced the outer 0.021 in.
(0.033 cm) of titanium. The primary effect of the composite in-lays is
to stiffen and lighten the blade. A comparison of the general structural
performance of both blades is shown in table II.
The most significant change is to the 3rd mode frequency (first
torsion) which increased by 20 percent at 4000 rpm for the blade with
composite in-lays. This correspondingly increased the reduced-velocity
flutter parameter from a negative margin of 0.1 to a positive margin of
0.26. The amount of untwist showed a comparable decrease of 24 percent
at 4000 rpm. The absolute value of the camber change at 4000 rpm was
reduced by about one-half compared to the all-titanium blade. Also
noteworthy are the increases in the frequency of the first mode of the
blade with the in-lays. These increases will shift the first mode
intersection with the second engine order (2E) line (figure 2) to about
2500 rpm which is well below the 80 percent operating speed (minimum
cruise speed).
The impact response of the blade with composite in-lays is shown in
figure 7. The impact conditions and reference point are the same as for
the all-titanium blade. The stress shown in figure 7 is for the titanium
part of the blade. It is seen that the maximum radial impact stress in
the titanium is about 120,000 psi (827 MPa). This is slightly higher
than for the all-titanium blade. The stress distribution in the boron/
aluminum composite has the same shape but almost twice the magnitude. As
such, the modified blade is a well balanced hybrid since the failure
stre s' s of the composite is about twice the yie l d stress of the titanium.
_._ __jj
The maximum steady radial stress showed the same behavior. This stress
in the titanium is 40,000 psi (276 MPa) at the reference point, slightly
less than for the all-titanium blade, and the maximum radial steady
stress in the composite is 82,000 psi (551 MPa). It is noted that the
composite in-lays are diffusion.-bonded to the titanium substrate. The
diffusion bond provides substantial interlaminar shear resistance at the
composite in-lay/titanium interface.
SUMMARY
The results of the structural dynamics analysis of shroudless, hol-
low, all-titanium fan blades and hollow titanium blades with composite
in-lays are as follows:
1. Th3 effect of engine speed on the untwist of both hollow tita-
nium blades is substantial and should be carefully considered in the
evaluation of blade performance since the aerodynamic load may couple and
magnify this effect as wel'i as influence the blade flutter stability.
2. The effect of engine speed on the camber of both hollow titanium
blades is relatively significant compared to acceptable changes permitted
for this parameter.
3. A possible blade flutter problem is indicated in the 3rd (first
torsional) mode for the hollow all-titanium blade.
4. The impact tolerance to bird strikes more than likely will yield
both hollow titanium blades.
5. The blade with the composite in-lays has a positive torsional
flutter margin, substantially less twist and uncamber and about the same
impact tolerance at the root as the all-titanium blade.
6. The design concept of shroudless, hollow titanium blades with
composite in-lays appears to be an effective approach to design energy
efficient fan blades for high bypass ratio turbofan engines.
REFERENCES
1. Chamis, C. C., "Integrated Analysis of Engine Structures", NASA
TM 82713, 1981.
2. Alexander, A., Cornell, R., "Interactive Multi-Mode Blade Impact
Analysis" NASA CR 159462, Aug. 1979.
TABLE 1. - HOLLOW ALL-TITANIUM BLADE
FREQUENCY - CPS
rpm Mode 1 Mode 2 Mode 3
0 48.2 142.1 252.1
1000 54.8 148.3 253.5
2000 70.5 165.0 257.3
3000 90.3 188.9 263.2
4000 111.8 216.4 270.8
(115.2)* (229.2) (280.8)
*The frequencies in parentheses were ob-
tained by constraining the blade along
the platform streamline.
TABLE II. - COMPARISON OF BLADE PERFORMANCE CHARACTERISTICS
Hollow
all-titanium
Hollow
titanium with
boron/aluminum
in-lays
Percent
change
Frequency (HZ)
0 rpm mode 1 43.2 56.3 +17
2 142.1 167.8 +18
3 252.1 309.9 +23
4000 rpm mode 1 111.8 115.9 +4
2 216.4 234.4 +8
3 270.8 326.3 +20
Weight	 (lb) 18.2 17.1 -5
Untwist	 (4000 rpm)
100 percent span 5.37 degrees 4.06 degrees -24
70 percent span 3.53 degrees 2.88 degrees -18
Camber change (4000 rpm)
100 percent span -.27 degrees -.13 degrees -52
70 percent span +.62 degrees +.34 degrees -40
Flutter margin -.1 +.26
(relative	 to	 1.4)
TIP
LEADING EDGE—'"
ROOT
6E 5E	 4E
ZD MODE
OND MODE
iT MODE
Figure L - Nastran model of the hollow fan blade (777 elements).
320
290
240
i' 200
z 160W
a
120
80
40
0	 2000	 4000	 6000
ROTOR SPEED, rpm
Figure 2. - Campbell diagram for the all-titanium hollor,
fan blade.
UNDEFORMED
LEADING EDGE,
(a) 1ST MODE SHAPE.
0 rum FREQUENCY 4B. 2 Hz
4000 rpm FREQUENCY 111.8 Hz
UN ')I,' +tMEDUNDEFORMED
LEADING EDGE —	LEADING EDGE
	
(b) 2ND MOUE SHANE.	 (c) 3RD MODE ; HA PE
	
0 rpm FREQUENCY 142.0 Hz 	 0 rpm FREQUENCY 252.1 Hz
	
4000 rpm FREQUENCY 216.4 Hz	 4000 rpm FREQUENCY 270.8 Hz
Figure 3. - Vibration mode shapes for the all-titanium hollow fan
blade (blade constrained at the root).
START OF HOLLOW SECTION
	 CROSS RIB
m
0	 t---- ^;
Figure 4 - Change in camber for the all-titanium fan
blade at 4000 rpm.
6y 4
3
zd
'_^ 2
START OF HOLLOW SECTION
	 CROSS RIB
L_
1 ksi - 6890 Pa
0
	
20	 40	 60	 80	 100
LOCATION, percent OF SPAN
Figure 5. - Reduction in twist for the all-titanium fan
blade at 4000 rpm.
150
120
90
60
.N 30
Y
l/1	 0
W
Ln
N
-30
-660
-M
-
000	 ON	 .00B	 .012	 .016	 .020
TIME, sec.
Figure 6. - Radial stress in the all-titanium hollow fan blade.
150
120
90
60
30H
3^ 0W
H-
N -30
-60
-90
-120
000	 .004	 .006	 .012	 .016	 .020
TIME, sec.
Figure 7. - Radial stress in the titanium /composite hollow fan
blade.


Structural dynamics of shroudless, hollow fan blades with composite in-lays

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