Transcript Rubrikbild - MCI Stockholm

Structures Concepts for a Mach 6
Supersonic Transport Aircraft
Rolf Jarlås
Flygteknik 2010
Outline
• Performance estimates
• Sizing
Baseline
(Lockheed HYCAT-1A)
Length 105 m
Gross weight 278 tonnes
Pax: 200
Range: 9000 km
Speed, Mach 6 at 28000 m
Design work is separated into three phases
(Aircraft Design: A conceptual approach [Raymer-99])
Conceptual design takes a look at the requirements:
 which requirements drives the design
 which technologies are needed
 how will the aircraft look like (estimate weight, fuel-fraction, wing-area, cost)
 will these requirements result in a viable aircraft at a saleable cost
Preliminary design will result in:




surface definition (lofting)
configuration freeze
major items design
test and analysis data base
Detail design: At the end
of the Detail design, when all drawings has been
produced and major items such as the structure, landing-gears, fuel-system etc. have
been optimized and tested, then
performance calculations.
it is possible to finalize weight and
Baseline
SI-units
Wing ref Area,Sref
Fraction
662,4
Fuselage length
(ref-length)
105,1
Capture area
8,11
0,01225
Wing span
29,7
Chord Centerline
40,1
Wing
”Centerline”
1,40
0,035
Wing thickness ”Tip”
0,14
0,035
Fuselage width
Trust/weight SLS
0,195898
17564
0,0632
Air inlets
7339
0,0264
Fuel&oil.syst.
2831
0,0102
LH2-containment
(fuel-tanks)
24253
0,0872
8,93
Cruise engines
2096
0,0075
10,50
Engine controls
402
0,0014
1363744
Altitude cruise
54485
Turbojets
thickness
Pass cabin length
PROPULSION
WEIGHT
0,5
28651
Range (km)
9260
GROSS WEIGHT
278128
Empty weight
155736
0,5599
Fuel
97014
0,3488
Payload
19051
0,0685
Oper.&Stand. it.
6328
0,0228
STRUCTURE WEIGHT
84666
0,304415
FURN.&EQUIPMENT 16585
FUEL
97014
0,059629
0,3488
Wing weight
18198
0,0654
Fuel Fwd tank
42201
0,435
Tail weight
5661
0,0204
Fuel rear tank
54813
0,565
Body
33865
0,1218
Landing gear
10836
0,0390
Surface controls
2495
0,0090
Thermal protection
12206
0,0439
Nacelle & thrust
structure
1405
0,0051
Aerodyn-surface coordinates
(288000 points)
Identify intersection points
50
42
24
Frame-stations
(in percentage)
36
26
90%
20
100%
70%
x
1
50%
25%
40
38
18
8
10
28
Determine cross-section shape
f given as input
41
35
Connect points (Meshing)
FS1
FS5
7
FS10
D3
51
25
1
7
Sizing D1
frames,
stringers,
skin.
(Create a few
groups)
D6
D4
D5
D2
Stiffness
Constraints
Aeroelasticity
Manoeuvrability
Performance
Strength
Manoeuvre-loads
Landing-loads
Actuator loads
Thermal loads
Minimum weight
(max. range, min cost)
Multidisciplinary trade-off
(Available volume etc.)
Initial baseline mass estimate is used
M1+LH2/2+THERM/3
PAY+OTH+THERM/3
ENG+
GEAR
M2+LH2/2+THERM/3
Structural weight is distributed by the FE-program
based on material(density)and sizing
Sizing is iterated to obtain baseline weight budget
Eigenmodes and Eigenfrequences
1.7 Hz
3.3Hz
Results after about 20+20 iterations to get
structural mass right and to increase lowest frequencies
Geometry update from DLR based on CFD-results
2.4 times as many surface-points, wing & cross-section redesign etc.
Steel was used as a “generic material”
Steel Stainless:
E/=200GPa/7640kg/m3=26 M(m/s)2 , 480 oC
Titanium:
E/=109GPa/4500kg/m3=24 M(m/s)2 , 450 oC
Aluminium:
E/= 70GPa/2700kg/m3=26 M(m/s)2 , 230 oC
Lockalloy/AlBeMet: E/=200GPa/2076kg/m3=96 M(m/s)2 , 370 oC
New lowest frequency with 3.7 tonnes of steel replaced by AlBeMet:
2.5 Hz (44% increase)
According to experience from test flights by NASA of the YF-12/SR-71 it is
possible to handle this fairly large aircraft with a bending-frequency of
about 2.5 Hz. The NASA-study was initiated after pilot induced oscillation
(PIO) problems occurring during in-flight refuelling. The YF-12/SR-71 had a
sophisticated stability augmentation system for its time to counteract pilot
induced oscillations. It is also likely that the test-pilots very skilled.
Aeroelasticity
No flutter or divergence problems identified
Strength/Stress
Stainless Steel (17-7PH):
 y /=900MPa/7600kg/m3=120 k(m/s)2 , at 480 oC
Titanium:
 y /=470MPa/4500kg/m3=105 k(m/s)2 , at 450 oC
Aluminium(Dural):
 y /=160MPa/2700kg/m3= 60 k(m/s)2 , at 230 oC
Lockalloy/AlBeMet:
 y /= 73MPa/2076kg/m3= 50 k(m/s)2 , at 390 oC
The temperature due to air-friction is higher, so some thermal insulation
and use of LH2 as heat-sink will be necessary
Horizontal stabilizer
Symmetric pull-up
150 MPa
Tail & rear fuselage
150 MPa at nz=1 correspond to
405 MPa at nz=2.7
Thermal load
Obviously DT=500 oC is to much!
, E=210 GPa
Concluding remarks
•General interface between CFD-surface-grid and structures model was created
•Flexible. A new model is based on the new surface-grid and
the sizing done for previous model(s)
•Weight estimates based on history should always be used, if available
•Stiffness rather than strength determine material, size and weight for baseline
•Thermal stress is a big issue (design, materials, thermal management)
•Material choice depend on temperature (thermal management)
•Detail design and analysis for a number of major parts needed to get more
accurate structural mass estimates
Aknowledgements
The ATLLAS project was co-funded by the
European Commission within the Sixth
Framework Programme, Thematic Priority-1.4,
AERONAUTICS and SPACE
Participation in ATLLAS by FOI was co-funded
by The Swedish Armed Forces