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AIRCRAFT MODELLING AND PERFORMANCE PREDICTION SOFTWARE
Key Aspects
INTRODUCTION
PA-31-325 C/R Navajo Accident
Objectives
MODEL BUILDING
Aircraft Description
Geometry
Assembling the Model
TESTING AND QUALIFYING THE MODEL
Sanity Checks
Reference Checks
PRELIMINARY ACCIDENT MODELLING
Updating the Model
Initialising the Scenario
Preliminary Dynamic Investigation
Adding the Detachment Force
AAIB INSPECTOR WORKSHOP
Additional Information
Matching the AAIB Findings
CONCLUSIONS
Supporting and Supplementing the Evidence
Pilot Evaluations
.
The J2 Universal Tool-Kit –
Supporting Accident Investigation
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INTRODUCTION
The AAIB asked the question whether the j2 Universal Tool-Kit, a full flight sciences design,
modelling and analysis tool, could be used to investigate the characteristics of crash scenarios. When
this was confirmed, j2 were requested to investigate the scenario of a particular accident that had
occurred. The results were compared to the findings from the report which is available in the public
domain.
PA-31-325 C/R Navajo Accident
The AAIB investigator provided the following case information and a series of questions/objectives.
The problem – calculate the response of the aircraft due to a section of the outer left wing becoming
detached due to tree strikes. The outer 2.4m of wing detached due to impacts with trees, which
coincidentally ties in with the outboard end of the left flap i.e. no left aileron remaining.
Representation of Accident - Note that the scan shows a PA-31-350 Chieftain with a
longer fuselage than the Navajo, but the wing is the same on both aircraft.
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Aircraft details
Weight at time of accident = 5,800lbs
Fuel on board at time of accident = 300ltr
CG position = 27.5” aft of datum
Landing gear was DOWN
Flaps set 10 degrees
Engines set to low power, props fully fine
Airspeed at point of impact with trees = 115 KIAS
Objectives
1. Assuming that the outer 2.3m of the left wing is instantaneously detached, and the right aileron
floats at zero deflection (aileron cable circuit now no longer continuous due to damage to left
wing), demonstrate the roll due to the imbalance
2. Apply a Point Load at the severed end of the left wing, due to the detachment forces of the
severed portion of left wing. Can J2 calculate the aircraft response in pitch, roll and yaw to the
left yawing moment?
3. Combine the above two effects to create a simulation of the flight path of the aircraft following
detachment of the left wing section?
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MODEL BUILDING
As the information provided was limited, the first stage of any project is to try to gather as much
information as possible. Whilst the investigator would be able to access some data from the
manufacturers directly, this was not available to j2 at this point, so all information had to be sourced
from data that was available in the public domain.
Aircraft Description
On 30 September 1964 Piper flew the prototype of a new twin-engine executive aircraft which was
then the largest built by the company. Identified at first as the Piper PA-31 Inca, the aircraft had
been re-designated as the PA-31 Navajo when deliveries began on 17 April 1967. A six/eight-seat
corporate/ commuter transport of cantilever low-wing monoplane configuration with retractable
tricycle landing gear, it was powered by two 224kW Avco Lycoming IO-540-K flat-six engines, and
was available in optional Standard, Commuter and Executive versions with differing interior layouts.
Made available at the same time was the optional PA-31T Turbo Navajo, which differed only by
having two 231kW TIO-540-A turbocharged engines, and the range was extended in 1970 by
introduction of the PA-PA-31P Pressurized Navajo with a fail-safe fuselage structure in the
pressurised section and two 317kW Avco Lycoming TIGO-541-E1A engines. Production of the PA-31
Navajo ended during 1972 and at the same time the company introduced for 1973 the PA-31-350
Navajo Chieftain which, by comparison with its predecessor, had the fuselage lengthened by 0.61m
and was powered by two 261kW TIO-540-J2BD turbocharged engines driving counter-rotating
propellers. A significant advance in the Navajo family came on 22 October 1973 when Piper flew the
first production example of the PA-31T Cheyenne, which combined an airframe generally similar to
that of the Pressurized Navajo with two 462kW Pratt & Whitney Aircraft of Canada PT6A-28
turboprop engines. In the following year an additional model of the Turbo Navajo was made
available, the PA-31-325 Turbo Navajo C/R, which introduced a 242kW version of the counter-
rotating engines installed in the Chieftain. Production of the PA-31P pressurized Navajo ended
during 1977, at which time a total of 248 had been built, but at the same time the company
introduced a new version of the Cheyenne, the PA-31T-1 Cheyenne I, the original Cheyenne then
becoming re-designated PA-31T Cheyenne II. Deliveries of the new Cheyenne I, which differed
primarily from its predecessor by having 373kW Pratt & Whitney Aircraft of Canada PT6A-11
turboprop engines, began towards the end of April 1978. The Cheyenne range was extended for
1981 by introduction of the PA-31T-Cheyenne IIXL, with the fuselage lengthened by 0.61m and
559kW Pratt & Whitney Aircraft of Canada PT6A-135 engines flat-rated to 462kW. In 1982
production of the PA-31 Navajo terminated after 1,317 had been built. Later production versions of
the Navajo family include the PA-31-325 Navajo C/R, PA-31-350 Chieftain and the PA-31T-1
Cheyenne I, PA-31T Cheyenne II and PA-31T-2 Cheyenne IIXL. However, the loss of the Navajo was
compensated for in 1983 by introduction of the PA-31P-350 Mojave, which basically combined the
airframe of the Cheyenne II with the powerplant of the PA-315-350 Chieftain.
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Summary
Aircraft Dimensions
Wing span 40 ft 8 in
Length 34 ft 7 in
Height 11ft 3in
Cabin Dimensions
Height: 4.25 ft
Width: 4.15 ft
Length: 8.6 ft
Volume: 151 ft3
Door Height: 3.5 ft
Door Width: 2.25 ft
Baggage External: 63 ft3
Occupancy
Crew: 1 or 2
Passengers: 5 - 7
Weight and Balance
Empty weight 4,250 lb
Maximum Takeoff Weight (MTOW) 8,130 lb
Max Landing Weight: 6,500 lb
Fuel Capacity: 1,122 lb
Payload with Full Fuel: 1,284 lb
Payload Max: 2,070 lb
Propulsion
Two Lycoming engines
Left TIO-540-J2BD
Right LTIO-540-J2BD
Horsepower 350 @ 2,575 RPM ea.
Propeller
Manufacturer Hartzell
Hub Model Left HC-E3YR-2A
Hub Model Right HC-EY3YR-2AL
Number of blades 3
Performance
Altitude Limits
Maximum Operating Altitude 27,000 ft
Service Ceiling 28,300 ft
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Speeds
Normal Cruise Speed: 178 kts
Max Cruise Speed: 226 kts
Climb Rate All Engines: 1,395 fpm
Climb Rate One Engine Inop: 115 fpm
Vs – Stall Speed 63.5 kts (flaps down)
VA-Manoeuvring Speed 162 KIAS
VMo-Maximum operating Speed 187 KIAS
VFE-Flaps Extend Speed
Mid: 152 KIAS
Full: 130 KIAS
VLE-Maximum Landing Gear Operating Speed 130 KIAS
Range
Normal Range: 640 NM
Max Range: 810 NM
Full Range (vfr): 757 NM
Max Range (vfr): 927 NM
Landing Distance
Landing Distance: 2,902 ft
Balanced Field Length: 1,750 ft
Geometry
The geometry is obtained from various sources (including Jane’s AWA) and aircraft 3-Views. Where
data is not available, it is estimated from aircraft of a similar class.
Wing
Span 12.4m
Area 21.3m2
AR 7.37
Chord
Root - Jane’s, at fuselage intersection 2.610m
Inboard Section Projected to ℄ 3.135m
Outboard Section Projected to ℄ 2.465m
Tip 0.97m
mac 1.826
Airfoil
Root NACA 63A415
Tip NACA 63A212
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Horizontal Tail
Span 5.52m
Area 5.603m2
AR 5.438
Chord
Root 1.380m
Tip 0.650m
mac 1.059m
Airfoil
Root NACA 0011
Tip NACA 0011
Dihedral 0°
Twist
Root 0°
Tip 0°
Sweep
LE 9.53°
25% 5.81°
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Horizontal Tail Geometry for use in Model Build
Vertical Tail
Span 2.1m
Area 2.898 m2
AR (for symmetric wing) 3.043
Chord
Root 2.070m
Tip 0.690m
mac 1.495m
Airfoil
Root NACA 0011
Tip NACA 0011
Dihedral 0°
Twist
Root 0°
Tip 0°
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Sweep
LE 43.46°
25% 38.073°
Vertical Tail Geometry for use in Model Build
Assembling the Model
Once the research had been performed and the geometry identified the next stage was to establish
the aerodynamic characteristics for the lifting surfaces, fuselage and undercarriage.
Lifting Surfaces
The lifting surfaces were calculated by putting the airfoil section information into JavaFoil and
running analyses over a range of angle of attack values at the cruise Reynolds’ Number.
The analysis for the wing was performed for Root and Tip Airfoil Sections with a range of flap
deployments, and a range of aileron deployments. Although the Flaps do not extend to the tips, and
the ailerons are not up to the fuselage section, the airfoil is only known at these sections. By putting
the values of the root and tips in, and specifying the limits for the flaps and ailerons, the JavaFoil
software will interpolate the information only over the region specified.
The Ailerons were calculated as a plain flap at 25% chord. For the Flaps, these were identified as
single slotted at 20% chord. As no geometry was available for these, a set of coordinates and pivot
point were calculated for the root and tip airfoil sections.
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Detail of Slot Geometry for Flaps
Flow Field with Flaps Deployed
Notes
To partially account for the spanwise lift distribution, the
values for the coefficients calculated at the root were
calculated using double the aspect ratio of the wing whilst
those for the tips were calculated using half the aspect ratio.
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From here we were able to establish the clean airfoil section details, and the increments due to the
flaps and the ailerons.
Lift Coefficient for Root Airfoil Section
Drag Coefficient for Root Airfoil Section
-1.5
-1
-0.5
0
0.5
1
1.5
2
2.5
3
3.5
-20 -15 -10 -5 0 5 10 15 20
α [°]
Flaps 0°
Flaps 5°
Flaps 10°
Flaps 15°
Flaps 25°
Flaps 35°
Flaps 40°
0
0.05
0.1
0.15
0.2
0.25
0.3
0.35
-20 -15 -10 -5 0 5 10 15 20
α [°]
Flaps 0°
Flaps 5°
Flaps 10°
Flaps 15°
Flaps 25°
Flaps 35°
Flaps 40°
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Pitching Moment Coefficient for Root Airfoil Section
A similar approach was used for the NACA0011 Airfoil with a plan flap for Elevator (37%) and Rudder
(40%).
-0.6
-0.5
-0.4
-0.3
-0.2
-0.1
0
-20 -15 -10 -5 0 5 10 15 20
α [°]
Flaps 0°
Flaps 5°
Flaps 10°
Flaps 15°
Flaps 25°
Flaps 35°
Flaps 40°
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Fuselage
To calculate the fuselage characteristics, the fuselage profile along with a simple lifting surface
geometry was run through DatCom. The resulting data is then collated into tables and the fuselage
contribution extracted.
Fuselage and Lifting Surface Geometry used with DatCom
Undercarriage
The undercarriage data was calculated from methods found in Roskam.
Propulsion
The propulsion system was broken into 2 sections.
Engines
The engines were simple powerplants that provided a direct link between the pilot throttle and
the output power of the engine. The engines were limited in their maximum power.
Propeller
The propeller takes the power from the engine and converts it into a thrust using the true
airspeed and the propeller efficiency. The efficiency was fixed at 70%. Whilst this is not the
most accurate approach, the variation in thrust that will occur over the crash profile will be very
minimal and as such an almost constant thrust can be assumed calculated from the thrust
requirements to maintain steady flight prior to the crash event.
Mass & Inertia
Whilst the Mass and CG position were specified in the initial request, this did not include the inertia
for the aircraft. The inertias were calculated using Radius of Gyration and the approach outlined in
Roskam.
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Assembly
With all the data and geometry defined, the final
stage is to put it all together. The Aircraft model is
assembled in j2 software using the structural
components and layout of the real aircraft.
Each item was also given its reference location
relative to their parent. The Centre of Gravity (C of
G) CG, Location of the Lift (NP) and in the case of
Propulsion Items the Centre of Thrust (C of T) was
defined.
The Z-Location of the undercarriage NP is variable
dependant up the deployment of the gear.
The aerodynamics for the fuselage can be added at
this stage because they are independent upon
anything else.
X Locations of Structural Items Relative to the Centre of the Nose (-ve Aft)
Structural Items on the Aircraft
Model
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Z Locations of Structural Items Relative to the Centre of the Nose (-ve Up)
With the structure laid out, the next stage is to add the
lifting surfaces. This is done through adding in the stripped
items. Each stripped item contains the geometry of the
lifting surface, including dihedral and twist. With each
horizontal item a maximum and minimum Y value is
included to account for the start and stop point. In this
way we can use the same geometry for all the wing
components. The j2 software will only calculate the
values for the appropriate spanwise positions.
Different Spanwise Max & Min
for Each Surface
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The final stage is to add the
aerodynamics. This is taken from
the values calculated earlier. For the
wing, the coefficients are functions
of the local angle of attack and the
spanwise locations. This takes into
account the change in airfoil section.
With the Ailerons and Flaps, there is
an additional parameter in the look-
up tables for the relative surface
deflection.
For the clean wing the values with
no surface deflection are added. For
the Ailerons and Flaps the values are
the increments in the coefficients
due to the surface deflection only.
When considering the Horizontal and
Vertical Tail, the Elevator and Rudder
were assumed to be full span. In this
case the total values for the clean
surface and the contributions due to
the surface deflection are included in
a single coefficient.
One aspect that needs to be
considered is the downwash from
the wing to the Horizontal Tail. This
uses the methods outlined in
Roskam, using the geometry and
relative locations of the Wing and
Horizontal Tail. To account for the
fact that there is downwash at zero
angle of attack, an alpha equivalent
value is used. This is calculated from
the lift curve slope of the clean wing,
and the total lift for the wing,
including flaps. This alpha
equivalent value is used in the
downwash gradient calculations.
The value of the downwash is then
applied to the horizontal tail through
applying a constant deflection (twist)
to the whole tail.
With the strip aerodynamics complete, we add in the propulsion characteristics, the undercarriage
drag contributions and the mass and inertia.
The complete aircraft model can now be qualified.
Adding in the Aerodynamics
Calculating the Downwash and Applying a Twist to the
Horizontal Tail
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TESTING AND QUALIFYING THE MODEL
With a completed model, the first stage is to run a series of tests. These tests serve two purposes.
The first is to ensure that the model has been assembled correctly, and that there are no unusual
characteristics present. The second purpose of the tests is to check that the values the model is
outputting are in line with expected results. Two types of tests can be run:
Sanity Checks
This is where the j2 Universal Tool-Kit is used as a virtual wind tunnel. A series of alpha and
beta sweeps will be run along with angular velocities to identify the total static characteristics
for the aircraft, and the dynamic derivatives. These numbers can be compared to values from
an aircraft of a similar type and class to ensure that they are not wildly different.
Reference Checks
The reference checks are when there are real data for the aircraft under test. This could be
wind tunnel data, or flight test data or whatever scenarios/data are available. In these
situations, the model can be set up to mimic the same condition and the resultant values
calculated and compared.
Sanity Checks
When running the sanity checks on the piper Navajo, the results for the static/dynamic derivatives
and surface contributions were found by running a range of sweeps for alpha and beta, and then for
each of the control surfaces. Additional roll, pitch and yaw rate cases were also run over a range of
angle of attack values. The derivatives were calculated and results were then compared to those for
Airplane B in Roskam’s Airplane Design: Part VI. These analyses can be set-up and re-used for
different aircraft with the results plotted on a template chart. This provides consistency and speed
in the analytical process.
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From the Sanity checks, the majority of characteristics were within the expected ranges for this class
of aircraft. Where there was a discrepancy, in the Roll Damping (𝐶𝑙 𝑝̂ ), where this was slightly higher
than expected, this was easily corrected by adding a correction factor to the overall model.
Reference Checks
Having got the model to a point that it is showing the appropriate characteristics for an aircraft of
this class, the next stage was to try some Reference Checks. For these, information was found from
the University of Tennessee Space Institute who have their own Piper Navajo that is used for
research and experimentation. From this aircraft various tests have been run than in turn have
produced some aerodynamic characteristics. By replicating the tests and comparing the results from
j2 Universal Tool-Kit to those from the real aircraft, further corrections can be found.
In this scenario, we have already built an aircraft model
with the mass and cg characteristics relating to the crash
to be investigated. As expected, these are not the same
conditions as for the Experiment. Therefore, a new Delta
Model was created where the mass and cg information
matched the UTSI Test.
The Aircraft was then initialised to the same flight
conditions and a further set of tests performed to enable
the data to be compared (classic flight matching).
Once the initial tests had been performed and the results
compared, it was found that whilst the Lift Curve Slope
(𝐶𝐿𝑖𝑓𝑡 𝛼
) was comparable, there was an initial lift offset
(𝐶𝐿𝑖𝑓𝑡0
) error. Small Drag Coefficient errors were also
found. Static corrections to the Lift and Drag for the
Clean Wing were added to the Clean Wing strips local
coefficients and the test cases re-run. The corrections
identified were shown to produce acceptable results.
Corrections to Mass & CG to Match UTSI
Test Point
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Comparison of Lift Coefficient (Before and After Correction) to the UTSI Experimental
Values
Comparison of Drag Coefficient (Before and After Correction) to the UTSI Experimental
Values
Once the corrections were shown to produce acceptable results, these were added back to the
baseline model. At the same time a small correction to the roll damping was also added as a global
coefficient on the Total Airframe.
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Corrections to 𝐂 𝐋𝐢𝐟𝐭 𝛂
,𝐂 𝐋𝐢𝐟𝐭 𝟎
,𝐂 𝐃𝐫𝐚𝐠 𝐈𝐧𝐝𝐮𝐜𝐞𝐝
, 𝐂 𝐃𝐫𝐚𝐠 𝟎
Applied to the Baseline Model
Correction to 𝐂𝐥 𝐩̂
Applied to the Complete Airframe
From here the sanity checks were re-run to establish the impact of the changes to Wing Lift and Roll
Damping to the Overall Dynamic Derivatives. These now fit into the acceptable ranges.
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PRELIMINARY ACCIDENT MODELLING
Updating the Model
When modelling the crash, we did not want to impact directly on the baseline model that we have
tested and qualified. The best approach was to create a delta model where the adjustments could
be made to account for the tree strike and the resulting loss of the outboard section of the wing.
This was done instantly within the j2 software
On the delta model, 2 new inputs were added.
Wing Lost (m)
This enabled us to define how much of the
wing, in metres, was severed as the analysis is
taking place.
Detachment Force (N)
This enabled us to add a force, in Newtons, to
the model to represent the impulse when
hitting the tree.
These were then connected into the model. The value from the Wing Lost (m) input was added to
the left outboard value (Min.y) of the clean wing geometry thus reducing the span of the left wing.
As this is driven from an input, the value of the Wing Lost (m) can be modified at any time during the
analyses
New Inputs on Crash Delta Model
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Changing the span of the Left Wing
When considering adding a force to account for the
tree strike, a Propulsion Item was used as this could
have a force added directly. The engine was given a
tilt value of 180°- so that it was always horizontal
to the ground regardless as to orientation of the
aircraft. The force to be applied would be at the
new wing tip position, and would be equal to the
Detachment Force (N) value that can be entered as
an input.
Initialising the Scenario
The next stage was to initialise the aircraft. As there was no information given regarding the flight
path angle, the aircraft was initialised in straight and level flight over a range of approach angles.
Using the trim rules built into the j2 Universal Tool-Kit, it was very easy to set the aircraft up in any
configuration.
The aircraft was initialised with:
Flaps at 10°
Landing Gear Down
Elevator Trimming out Any Pitch Acceleration
Wings Level
Constant Airspeed of 115 KCAS
Flight Path Angle from -7° to 0°
Zero Forward Acceleration
Altitude aligned to flight path angle so after 5s the altitude is 60ft
Adding the Tree Strike Force to the
Model
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Completed Analyses Initialising Model
Preliminary Dynamic Investigation
The first objective was to identify the rolling characteristics of the aircraft with the wing section
removed. To do this a response model was set up that allowed the aircraft to travel unaffected for
5s and then at 5s a step input of +2.4 was added to the Wing Lost (m) input. As shown before, this
will result in the Left Wing having a semi-span of 3.8m. The aerodynamics within j2 will now ignore
any contributions outboard of 3.8m on the left wing, but will include all the contributions from the
right wing.
Defining the Instantaneous Loss of 2.4m of the Left Wing
Starting from the Level Flight Initial Condition (gamma=0°), the response of the aircraft to this event
was analysed. The analysis automatically terminates when the aircraft hits the ground.
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Lateral Response to Instantaneous Loss of 2.4m Left Wing
As we can see from the above chart, when the left wing section is lost, there is an imbalance in the
lift with increased lift on the right wing and reduced lift on the left wing. This results in a negative
rolling moment (to the left). In this scenario, the aircraft reaches a maximum roll rate of 24°/s, to
the left, and a bank angle of almost 90°, left wing down, before impact with the ground. What can
also be seen here is the time from tree-strike to impact is approximately 4.5s.
Whilst charts are very useful at showing us the behaviour of the aircraft, it is often still difficult to
fully interpret the complete behaviour. It is especially difficult in a situation such as this where the
aircraft is moving in all 3 axes. In this situation, engineers will typically resort to using small models
to follow the attitude values of the aircraft to try to fully understand what is happening. This is not
necessary with j2, as it is possible to instantly take the time history generated and display it in a 3-D
Virtual Environment Playback. Within this playback, it is possible to move around the aircraft,
viewing it from different angles, zoom in/out, and speed up/slow down time to get a more detailed
and highly visual understanding of what actually happens.
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3-D Playback of Preliminary Study
Adding the Detachment Force
Having satisfied Objective 1, the next stage was to
see what would happen should a detachment force
be applied. A method for applying a force to the
model has already been created, so the next stage
was to add in the detachment force through the
response model. It was estimated that the time
taken to sever the wing at the flight speed would
have been approximately 0.03 seconds. So to
account for this, an initial force was applied that then
dissipated over 0.03s and then ceased. This force was
combined with the wing loss to build the scenario.
No value for the detachment force was provided, so the mechanism used allowed for different
forces to be applied to help to identify what was probable. An initial estimate for the force was
calculated by j2.
Combining the Detachment Force and the
Severed Wing in a Response Model
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The Detachment Force Applied to the Wing at the Severed Location
We can now re-run the response analysis from the same start point as before and chart the new
values, without writing any code.
Lateral Response to Instantaneous Loss of 2.4m Left Wing and Detachment Force
We can see from the new chart that the reaction is much more violent. There is an instantaneous
yawing moment due to the impact with the tree that causes the aircraft to yaw up to -50°/s. This
Yaw rate increases the lift over the right wing which gives an even greater rolling moment such that
the aircraft now reaches a maximum roll rate of 140°/s, to the left, and a maximum bank angle of
130°, left wing down, before impact with the ground.
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3-D Playback of Wing Loss plus Detachment Force
At this point the primary 3 Objectives had been satisfied, and the next stage was to discuss our
findings with the AAIB inspector.
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AAIB INSPECTOR WORKSHOP
Once the primary objectives had been satisfied, a short workshop was arranged, where we could
present our results and discuss the findings with the AAIB Inspector.
The Initial presentation of method and results was understood, and various key points of the
modelling and scenario we agreed upon. However, there was viewed to be some discrepancy
between the evidence found by the AAIB inspector and the orientation of the aircraft, from the
modelling, as it hit the ground. The evidence trail from the crash indicated that the aircraft hit the
ground in the following sequence, Right Wing Tip, Right Engine, Nose. Whereas the modelling
indicated, that the left wing would strike the ground first.
Additional Information
When discussing the discrepancies, further information was provided that was not presented in the
original data supplied to j2. In the first instance, the tree strike took place on a hill, and the impact
with the ground was 120ft below that location. Secondly, the aircraft was on a gradual descent,
approx. 3°, as shown from the radar track. Below is an extract of the AAIB Report, available in the
Public Domain.
“The aircraft’s left wing had struck two 80 ft pine trees approximately 20 ft below treetop height,
causing the outer 2.2 m of the left wing, outboard of the left flap, to fragment and detach. Witness
marks on the severed tree trunks indicated that the roll attitude of the aircraft at impact with the
trees was wings level.
The wreckage trail, from the point of
the initial tree strikes to the aircraft’s
final resting position, was 230 m in
length and was orientated on a
heading of 298°M. Parts of all major
sections of the structure and flying
controls were identified at the
accident site. The first ground impact
scar was 185 m beyond the initial tree
strike and had been made by the right
wingtip; it was 20 cm wide and the
narrowness of this mark indicated
that the roll angle at ground impact
was approximately 90° right wing low.
The ground impact marks and
distribution of the wreckage indicated
that following the right wingtip strike,
the right engine hit the ground and
detached, shortly after which the
aircraft impacted heavily on its nose.
The absence of any significant ground
scars between the nose impact crater
and the main wreckage indicated that
the aircraft then bounced a distance
of 34 m, before finally coming to rest
facing uphill.”
G-BWHF Radar Track (Map terrain elevations are shown in
metres)
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Whilst this information was
different to the original
information presented, it
was easy to accommodate
these changes instantly. To
account for the change in
height, the response was
simply offset by 120ft and as
we had already initialised
the aircraft over several
flight path angles the
analysis was run from the 3°
descent initial conditions.
Response Results to Updated Altitude and Approach
What can be seen when we change the initial altitude and flight path angle is that there is very little
difference between the Initial Conditions that matched the AAIB Report and the baseline case. As
expected, the approach angle has very little effect on the maximum Roll Rate (P) and despite the
flight lasting a little longer due to the increased altitude, the aircraft does not achieve any larger a
bank angle as the roll rate dropped considerable by the time the aircraft is approaching the ground
such that there is very little increase in bank angle.
Adding an Offset (Disturbance) to the Initial Altitude
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Comparison of Bank Angle for the Original Case and Using the Initial Conditions from
the AAIB Report
Matching the AAIB Findings
Part of the Workshop with the AAIB Inspector was to look at evaluating different scenarios in real
time based on additional information as it is presented. Due to the flexibility and power of the j2
Universal Tool-Kit we were able to study numerous and varied what-if scenarios to identify what
conditions may have occurred to present an impact similar to that found through the debris trail.
We had already established that the physics model and aerospace knowledge dictates that the
aircraft will have a negative (left wing down) roll rate for an incident of this type. Thus we knew that
the only way for the right wing to hit the ground first was for the aircraft to rotate beyond the
inverted position (180°) before impact. Therefore the first objective was to identify what areas of
the analysis may need adjusting to achieve this scenario within reasonable limits. Three areas were
investigated:
Aircraft Inertia
The inertia estimated for the model may be slightly high. In addition, losing part of the wing
would reduce the roll inertia further. A trade study could be performed to identify the effect of
reducing the inertia.
Roll Damping
The roll damping had already been found to be slightly high for the class of aircraft, and a
correction had been applied. A study of the effect of a further reduction in the roll damping
coefficient could be performed.
Detachment Force
This had been estimated from simple calculations. A trade study could be performed to assess
the bank angle when changing the values of the impact force.
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Control Surface Deflection
The original briefing explained that the aileron cables were severed but the aileron remained at
0° deflection. However, previous Accident Studies performed by j2 have found that there is a
major impact when a control surface is able to “float” during a dynamic manoeuvre. If the
cables had been severed, then it was highly likely that the right aileron would move freely as a
result of the aircraft dynamics and inertia.
All the studies were quickly and easily performed and the resulting flights charted and visualised
during the 2 hour workshop with the AAIB Inspector watching and commenting. This meant we
were able to perform live analysis with a qualified inspector adding knowledge and experience into
the process concurrently.
Roll Characteristics
The model was first updated by reducing its roll inertia (IXX) and the scenario ran again. It was
quickly found that to get the aircraft to roll beyond 180° the roll inertia would need to be reduced to
a level that it was deemed to be unrealistic.
The next stage was to look at the roll damping. As with the roll inertia, the level of change required
to achieve the desired result was viewed to be too severe to be realistic.
A final study was run to see if a “realistic” reduction in roll inertia and roll damping would cause the
desired effect. This was shown not to be the case.
Detachment Force
When looking at the detachment force, it was increased by 10% to see what sort of impact this may
have on the overall bank angle of the aircraft. It was found that this increase would be sufficient to
roll the aircraft over to the right wing.
Lateral Response to 10% Increase in Detachment Force
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Right Wing Hits the Ground first with 10% increase in Detachment Force
Control Surfaces
From previous accident
studies performed by j2, it
has been identified that
when left to float (i.e.
Control Cables Snapped)
control surface are unlikely
to remain at zero deflection.
This is especially true when
there are highly dynamic
manoeuvres. As such, it was
decided that the right
aileron may move following
the crash. The motion of the
aileron was dependent upon
the roll acceleration and
damped out by the roll rate.
Driving the Ailerons using the Roll Dynamics
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Lateral Response to “Floating” Aileron
Left Aileron Deflection when “Floating”
It can be seen that the aircraft rolls beyond the inverted position and the aileron deflects up to 8°.
However, when looking at the visualisation we can see that the rotation is insufficient for the right
wing to hit first.
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Right Engine First with Aileron “Floating”
Combinations
With a small increase in the Detachment Force or a small deflection of floating aileron both increase
the bank angle beyond the inverted. Whilst the detachment force increase can achieve the desired
situation alone, a 10% increase was viewed to be too much of a correction. Similarly, any further
aileron deflection was assumed to be too large. As such a combination situation was put together.
The Detachment Force was increased by 5% and the impact of the roll acceleration (P’) on the
aileron was reduced slightly. The aileron characteristics were adjusted and re-flown several times
and each time the flight was visualised instantly to see what the aileron deflection was, and how the
aircraft impacted with the ground.
After only a few attempts, a response was found that appeared to fit the evidence very closely. The
Detachment Force was within limits, and the resultant “Floating” Right Aileron Deflection was within
limits. The aircraft orientation upon termination of the flight was 95° Right Wing Down. Thus the
right wing tip hits the ground first, followed by right engine and finally nose.
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Ground Impact with Combination Scenario
One final sanity check was to look at the predicted distance the aircraft travelled from impact with
the tree to the first impact with the ground.
Flight Path Profile
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What was found from the analysis was that the analyses terminated when the aircraft was 191m
from the tree. If we look at the first impact with the ground, this is calculated at 184.5m from the
impact with the tree. This compares very well to the 185m presented as evidence from the crash
site.
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CONCLUSIONS
Supporting and Supplementing the Evidence
The model build, analysis performed was able to satisfy all the objectives presented by the AAIB
inspector and provide dynamic “video” footage of the scenario.
The subsequent workshop with the AAIB inspector was able to fill in the additional information
provided and use this to identify how the aircraft was able to rotate almost 270° from impact with
the trees. These results were consistent with the AAIB findings including the location of the first
impact with the ground.
AAIB Photo of Crash Site and Corresponding j2 Flight Path
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From being presented with the project to the final workshop and identifying a scenario to match the
AAIB findings was 10man days. This shows that the j2 Universal Tool-Kit is not only a very powerful
and capable tool for investigating air accidents, but the results can be obtained in a fraction of the
time of more traditional approaches.
Automatic Generated Playback of Flight from j2 Virtual
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Pilot Evaluations
One further capability available when working with the j2 Universal Tool-Kit is the ability to fly the
aircraft in a manned simulator without the need to write and develop any code. This means that as
the aircraft model is being developed, it can be tested and evaluated by a pilot concurrently. Finally,
when the delta model was constructed to add the tree strike force and loss of wing, it is possible to
have these automatically added into the Instructor/Operator Station and can be injected at any
time. This means the controllability of the aircraft after the tree strike can be fully evaluated.
Configurable Simulator provided by j2 Pilot
Providing a Fully Integrated System with Engineering Workstation Directly Integrated
with Desktop or High Fidelity Simulator
For further information please go to www.j2aircraft.com or contact info@j2aircraft.com
