Investigation of Crack in Gyroplane Main Rotor Blade

Transcription

Investigation of Crack in
Gyroplane Main Rotor Blade
P E Irving
N Smyth
September 2010
Rept GS1
School of Applied Sciences
28th September 2010
Page 1 of 20
BACKGROUND
On the 9th of September 2010 Mr. G. Speich, of RotorSport UK visited Cranfield
University to deliver part of the main rotor blade assembly of a RotorSport UK
Gyroplane. During a pre flight inspection, a crack had been found on the bottom
surface of the main rotor blade 5106A. Cranfield University were requested to
investigate the failure and determine:
1. The extent of the crack,
2. If the crack originated from the inside or outside surface,
3. Whether an internal rectangular support bar, the termination of which was
located close to the centre of the crack, played a role in crack initiation,
4. The length of service that the crack had been present in the structure.
DESCRIPTION OF ROTOR ASSEMBLY
The main rotor consists of two blades each manufactured of extruded aluminium alloy
6005. The blades presented to Cranfield University have serial numbers 5106A
(visible crack) and 5106B (still assembled). The dimensions of the blades are
approximately 0.2m chord length, 0.024m maximum depth, and 4m long. A cross
section through a blade is shown in Figure 1. Each blade is connected to the rotor hub
assembly via a bolted joint with 9 bolts equidistant apart along a line adjacent to rib B.
In the joint the blade extrusion is sandwiched between an aluminium doubler (0.3m
long and of varying thickness) and two plates of aluminium (0.01m thick and 0.07m
wide) on the top and bottom surfaces, blade 5106B is shown assembled in Figure 2.
An aluminium bar is inserted and bonded into the internal space between ribs B and C
shown in Figure 1. This extends the entire distance of the bolted connection,
terminating at a point outboard of the final hole and approximately coincident with the
end of the upper and lower connection plates. Figure 2 shows the disassembled
connection area; the dark outline indicated marking the boundary of the clamped area
between the lower connection plate and the blade extrusion. The mark outboard of the
outermost connection hole is the approximate crack origin.
BLADE MATERIAL
The blade was stated to be manufactured of 6005 aluminium. This is a low strength
aluminium alloy with a minimum 0.2% proof strength of 240MPa, minimum ultimate
tensile strength of 260MPa, and 8% ductility. Based on knowledge that the high cycle
fatigue strength of aluminium alloys are approximately 0.3 of the UTS, the plane
specimen fatigue strength of the alloy will be approximately ±80-90MPa, at a mean
stress of zero. The lower surface of the rotor will experience fatigue loadings on
Ground Air Ground (GAG) cycles with a minimum stress of approximately zero.
Under these circumstances the fatigue strength for zero to tension loading applying a
Goodman correction can be estimated as ±65MPa, with a maximum stress of 130MPa
and a minimum of zero.
Rib A
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Rib B
Rib C Rib D
Figure 1: Labelled Cross Section through Rotor Blade
Plates
Doubler
Crack
End of Doubler
Figure 2: Main Rotor Blades 5106A (Bottom) and 5106B (Top)
CRACK LOCATIONS
The location of the crack on blade 5106A is shown in Figure 3 and was measured at
27mm from the centre of the outermost bolt hole and to be 73mm in length. The blade
5106B was disassembled from the supporting structure and a small crack was
revealed, dimensions shown in Figure 4. The crack was located on the underside of
the blade in a similar location to the crack on blade 5106A. Both cracks on 5106A and
5106B showed evidence of fretting corrosion on the lower surface, as seen in Figure 3
and Figure 4. The origin process can be seen more clearly on 5106B (Figure 4). There
are a number of dark patches caused by mating sites rubbing (local fretting). The
crack has initiated from one of these outboard of the outer bolt hole. There are similar
marks on 5106A (Figure 3) though the details of the origin are obscured by fretting
product. The blades were visibly observed to have a slight bend upwards.
The location of the crack on both blades relative to the doubler/plate end and internal
reinforcing bar is shown in Figure 5. The cracks on both rotors were located inboard
of the end of the doubler/plate and reinforcing bar.
Page 3 of 20
The fretting corrosion residue (dark region on Figure 3) was only found on the
outboard side of the crack. This may be because of the centrifugal force from the
rotating rotor conveying the residue in the outboard direction only. A similar
observation can be made on the internal side of the crack (Figure 6).
INBOARD
27mm
73mm
Figure 3: Crack Dimensions on Blade 5106A
INBOARD
6mm
23mm
Figure 4: Crack Dimensions on Blade 5106B
Page 4 of 20
Upper and Lower Skin on Doubler
Internal Reinforcing Bar
Position of Crack
Upper & Lower Attachment Plate
Upper and Lower
Skin of Airfoil
Figure 5: Cross Section through Blade Assembly
Figure 6: Inside Surface of Blade 5106A after Cutting
The blade was cut further and pulled apart to reveal the fatigue fracture face, a
photograph of which is shown in Figure 7. There was further evidence of fretting
corrosion on the face. During the crack separation procedure new fractures were
created, the ends of the original fatigue fracture region are highlighted in Figure 7 by
means of white dotted lines.
A representative SEM image is shown in Figure 8 clearly showing fatigue striations.
The striations appear to be grouped into bands and it is conceived these bands may
represent landing/take-off cycles with the intermediate striations representing in flight
manoeuvre loads. The varying distance between the bands represent the varying
durations of the flights.
Page 5 of 20
Initiation Region
Figure 7: Fatigue Fracture Face of Blade 5106A
Manoeuvre Loads
Landing/Take-Off
Figure 8: Fatigue Striations on Fracture Face of Blade 5106A
Measured striation spacings for the GAG (indicating growth increments per flight) at
three locations on the fatigue fracture face are given in Figure 9, Figure 10, and Figure
11. Figure 9 is measured 25mm from the initiation point, Figure 10 is 8mm, and
Figure 11 is 4mm. The average GAG spacing at the three locations is 46µm, 13µm,
and 31µm respectively.
Figure 9: GAG Spacings 25 mm from Initiation Region
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Figure 10: GAG Spacings 8 mm from Initiation Region
Figure 11: GAG Spacing’s 4 mm from Initiation Region
Page 7 of 20
Page 8 of 20
DISCUSSION OF OBSERVATIONS
Initiation Site and Conditions
The crack in both the A and B blades has initiated at locations of fretting damage at
almost identical positions, 20-30mm outboard of the outermost hole centre,
approximately along an extension of the line of the holes and 4-5mm inboard of the
termination of the inner aluminium rectangular bar. The crack origin is within the area
clamped by the outer aluminium doubler and is therefore subject to compressive
forces at the location point. From the origin the crack spread down through the outer
skin and outwards in both directions towards the leading edge and the trailing edges
and down into the ribs within the extrusion. The initiation process can be seen at an
earlier stage in blade B (Figure 4) where there are a number of local fretting sites one
of which has initiated a fatigue crack. Fretting conditions will reduce the material high
cycle fatigue strength by a factor of 3, making initiation at these locations extremely
probable.
Role of Reinforcing Bar and Bending Overload in Crack Initiation
The observation that the blades were slightly bent upwards, suggests that they may
have been overloaded during service creating local plasticity. As 6005 is a very
ductile alloy (minimum of 10% elongation to failure), it is unlikely that bending
overload producing 2-3% plasticity would directly cause cracking. Strains in excess of
10% would promote local cracking, but there was no sign of deformations of this
extent at the crack locations.
The internal reinforcing bar may have influenced the load transfer from doubler to
blade. A redesigned blade without a bar is reported to have failed via a crack in the
outermost hole. This would be the expected site in the absence of an internal
reinforcing bar as stresses will be at their greatest at the outermost hole edge. In the
blade under investigation the internal bar may have acted to transfer local loads away
from the hole and into the blade skin outboard of the bolt hole; leading to enhanced
stresses at the observed crack location. During the crack extraction process it was
observed that the bar/blade bonding was locally delaminated at the crack site. This
will have reduced the effectiveness of the bar/blade load transfer making the region
between the outermost hole and the bar end a region of enhanced stress, rather than
the bar end itself.
The role of the bend overload may have been to change the contact conditions making
fretting more likely. When fretting fatigue conditions exist, the fatigue strength could
be reduced from the ± 65MPa suggested earlier for normal fatigue, to the order of
±20MPa in aluminium alloys in tension. The stresses associated with the overloads
will promote additional fatigue damage but they will not cause cracking directly.
Fractographic Observations
Observations on the fracture surface did not suggest any material defect which could
have promoted the crack initiation. Indeed the observation of cracks in both blades
and in other rotors suggests that the failure origin is not in material anomalies, but is
to do with either design or operation of the rotor.
The probable identification of GAG growth increments on the fracture surface allows
very rough estimates to be made of how long the fatigue crack has existed. At the
Page 9 of 20
largest increments observed of 50µm/GAG, growth will be at 1mm every 20 flights.
Taking a half crack length (crack is growing both ways and has 2 fronts) of 36mm,
this suggests that the crack has existed for at least 700 flights.
Service load measurements
The initial service load measurements presented in Appendix A and interpreted in
Appendix B are very limited and brief but they do allow initial calculations to be
made of the likely fatigue and fatigue crack growth behaviour of the blade under the
measured service loads. These are described in Appendix C. It is unlikely that the
stresses measured fully reflect the extremes of loading placed on the failed blades, as
a maximum stress of 179MPa has been recorded. This is insufficient to cause
plasticity and permanently bend the blade. The 6005 alloy has a specified minimum
proof strength of 240MPa, possibly the blade strength level is greater still, and
stresses larger than this will be required to cause plastic deformation.
Nevertheless, the recorded fatigue cyclic stresses are sufficient to cause very early
crack initiation in fretting. Data developed in Cranfield fatigue laboratories suggest
that only a few hundred cycles of stresses of this range would be necessary to create a
fatigue crack in aluminium alloys under fretting conditions.
Use of the service load measurements to calculate crack growth life using AFGROW
Appendix C describes fatigue crack growth life calculations made using a
conservative flight service loading spectrum constructed using the supplied service
loading information as input to the software fatigue crack growth calculation package
AFGROW. The constructed spectrum assumed that each flight consisted of 60 repeats
of the worst case manoeuvre loading recorded in the data supplied. Material crack
growth data for 6061 aluminium sourced from the AFGROW database was used in
the analysis. The results show that under the conditions assumed, the number of
flights to grow the crack from 1mm starting half crack size up to a crack length of
72mm total length is approximately 297 flights. The analysis was repeated using
different sets of assumptions about the material data, this resulted in a life of 700
flights.
There are both conservative and non conservative aspects to this calculation. On the
one hand, the constructed spectrum represents a severe worst case load history derived
from the data supplied. On the other hand, it is unlikely that this limited data set fully
represents the extremes of the possible loads which as noted above can be so severe as
to cause blade bending.
The best course of action would be to derive a spectrum fully representing rotor
stresses on a comprehensive service loads measurement data set. This would enable
an accurate calculation which could be given with full confidence. The present
analysis can only be regarded as preliminary.
Page 10 of 20
CONCLUSIONS
1. Visual and scanning microscope observations of fatigue cracks propagating in
both blades of a gyroplane rotor suggest that they originated in fretting fatigue at
locations in the external surface within the clamped area of the doubler connecting
the blade to the hub. The blade was bent suggesting it had been overloaded; this
may have promoted fretting fatigue conditions.
2. The crack subsequently propagated transversely towards both the leading and
trailing edges and down through the ribs within the extrusion.
3. Fractographic observations suggest that GAG growth increments can be
identified. These have a maximum spacing of approximately 50µm and suggest
therefore that 20 flights are required to grow the crack 1mm. The blade with the
longest 73mm crack had therefore approximately been cracked for at least 700
flights.
4. Preliminary service load information used as input into a fracture mechanics
analysis of fatigue crack growth behaviour shows that service life to grow the
crack from 1 mm half crack length to 36mm final half crack length is between 300
and 700 flights.
Page 11 of 20
APPENDIX A – RotorSport UK Service Strain Measurement Report
Page 12 of 20
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Page 18 of 20
APPENDIX B – RotorSport UK Service Strain Measurement Interpretation
The interpretation of the service strain measurement report is detailed here. It is stated
in the report (see Appendix A) that “the strain gauges are calibrated in microns per m”
which was assumed to be micro-strain.
It is also stated that the blades were “fitted with GS3BT strain gauges, one on the
lower surface of the blade adjacent to the blade at the end of the hub bar...and one
outboard approx 1m from the end of the hub bar to measure the centripetal force in
the blade”. However it should be noted that using only one strain gauge on the lower
surface it is not possible to differentiate between bending and tension induced strain.
However for the purposes of the current analysis a combined tension/bending strain
measurement is sufficient.
For each of the graphs of the RotorSport UK service strain measurement report the
maximum and minimum strains were noted for both the blue and green lines. The
report mentioned that “the scaling factor within the system are such that the blue line
MUST be read double scale”, therefore strains were doubled for the blue line. To
convert strain to stress, the measured micro-strain was multiplied by the Young’s
Modulus (69GPa). The noted strains and resulting stresses are shown in Table 1.
Table 1: Calculated Stresses from Strain Data
Graph Page No.
2
3, 5, 7
4
6
Line
Blue
Green
Blue
Green
Blue
Green
Blue
Green
Strain (µε)
Min Max
1200-1600
650-750
1200-2150
625-925
1150-2600
600-1075
950-2150
575-900
Stress (MPa)
Min Max
83-110
45-52
83-148
43-64
79-179
41-74
65-148
40-62
The greatest minimum and maximum stresses at the location of the crack (blue line)
were 65 and 179MPa respectively.
Page 19 of 20
APPENDIX C – AVGROW ANALYSIS
AFGROW is a software package which simulates fatigue crack development in
metallic materials using fracture mechanics approaches to fatigue crack growth. It was
developed by the US Airforce over a 10 year period to solve fatigue crack growth
problems on military aircraft and is publicly available.
AFGROW contains a database of material fatigue crack growth properties necessary
for performing the fracture mechanics calculations. The alloy 6005 used for the rotor
blade is not available in the database, but the closely related 6061 alloy is in the data
base. It is unlikely that there will be significant differences in the crack growth
characteristics of 6061 and 6005 as they are of similar strength and similar aluminium
alloy series.
Apart from the material fatigue crack growth properties and the static fracture
toughness of the material, also in the data base, the other data required for the crack
growth prediction are the stress spectrum experienced by the blade in service. These
data were extracted from the preliminary data contained in Appendix A, and a simple
spectrum constructed containing two components, based on the worst case stresses at
the maxima and minima for the Ground Air Ground cycles and the worst case stresses
associated with the manoeuvres. These should give a conservative calculated number
of flights to static fracture.
The spectrum derived from the worst case data in Appendix B is shown in Figure 12
below. It was assumed that each flight consisted after takeoff of 60 worst case
manoeuvres in a 1 hour flight, followed by landing. In the analysis, this spectrum was
repeated until failure of the rotor was predicted.
The final item of data required for the analysis is a starting crack size. This is the
crack size at the end of the initiation stage of the failure. In fretting fatigue we believe
that the initiation stage took a very small number of cycles at the GAG stress level of
zero – 179 MPa.- perhaps only a few hundred cycles. This would create a crack of
around 1 mm size. To perform the fracture mechanics calculation of life it was
assumed that the blade lower surface could be approximated to a plate 200 mm wide,
1.4 mm thick and of length in excess of 400 mm.
The analysis was repeated using different assumptions about material properties and
the form of the crack growth data.
The result of the analysis showing a plot of cycles Vs Crack length describing the
crack growth process is shown in Figure 13 below.
Figure 12: Stress Spectrum used in AVGROW Analysis
Figure 13: Life Vs Crack Length
Page 20 of 20

Investigation of Crack in Gyroplane Main Rotor Blade

Transcription

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