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SAFE HOMEBUILT DESIGN AND FAA CERTIFICATION REQUIREMENTS
BY
DAVID B. THURSTON EAA Sport Aviation Oct.
1982
Introduction
Seven years after the initial
announcement of the Model TA 16 TROJAN homebuilt amphibian in SPORT
AVIATION (August 1975, page 62 and September 1977. Page 29), this
four-place, 250 horsepower design of Figures 1 and 2 is being constructed
by 45 builders in the United States and Canada; with the first homebuilt
TROJAN expected to fly early in 1983. As design work progressed it became
evident that a large international market existed for an airplane of this
type and size.
As a result, International Aeromarine Corporation was
formed to develop the TA 16 as an FAA Type Certified production amphibian.
In order to distinguish between homebuilt and certified production
aircraft, the name SEAFIRE was adopted for the production model. Our
prototype SEAFIRE has been structurally designed and static tested to FAR
Part 23, Amendment 24 requirements, with final assembly, high speed taxi
tests, and ground cooling runs successfully completed.
Since detail
design is now finished, this airplane provides an excellent basis for
comparison between safe, conservative homebuilt design criteria and FAA
Type Certification requirements some of which can be overly burdensome,
if not completely unnecessary for small aircraft flown primarily for
personal pleasure, travel, and sport activity.
This is not to
suggest that all Part 23 requirements are unfavorable, far from it, but
rather that a fairly small percentage add considerably to Type
Certification costs without providing any appreciable improvement in
personal aircraft flight safety or ease of operation. In view of this, we
should be able to develop and offer homebuilt designs complying with the
basic structural and flight safety features of Part 23 at a considerable
saving in design and construction cost when compared to factory produced
certified aircraft. In other words, designers of homebuilt aircraft can
and should offer new models which will be as structurally and
operationally safe as FAA certified design, but which can be built for
considerably less cost than a comparable production airplane.
This
presentation highlights those areas in which FAR Part 23 design criteria
result in unnecessary design and manufacturing cost increases while
providing slight if any contribution to safety, or ease or operation, and,
possibly more importantly, those Part 23 requirements which deserve full
consideration for adequate structural and flight safety of any design
intended for homebuilt construction.
The Negative Aspects of Part
23
Bearing in mind that these comments apply to small personal aircraft
and do not necessarily include designs intended to carry passengers for
hire, the following regulation areas add little or nothing but cost to
aircraft development:
Tab control and drive system criteria
Drop
test requirements
Warning horn in amphibian throttle and flap control
systems to indicate landing gear position
Lightning strike analysis
report
Structural life fatigue analysis report
Time permits only a
brief review of these items, so let us look at each one in order of
presentation.
For many years, a bungee spring loaded elevator trim
system was considered acceptable for providing the mount of elevator
surface deflection necessary to trim aircraft in flight. The use of trim
springs as shown in my book "Design for Safety", pages 37 and 38, were
both simple and effective for small, light aircraft. Not only did this
system eliminate control surface tabs which complicate surface
construction and are very subject to flutter, but it also eliminated long
runs of small diameter trim cable which could act as an oscillating spring
under high load conditions. In addition, the bungee trim system could be
designed to increase stick load with elevator surface deflection
frequently a desirable feature on light aircraft having low inertia in
pitch (small longitudinal moment if inertia).
As time went by, the
FAA determined that springs in the elevator system might cause surface
flutter. They also developed a criteria requiring flight demonstration to
show that in the event of failure of the primary pitch control system (the
elevator control), the pilot must be able to land the airplane by use of
the pitch control. This necessitated provision of an elevator trim control
and surface tab separate from the primary elevator control
system.
In 42 years of aircraft design I have never heard of the
structural failure of any elevator control system unless it had been shot
away by enemy gunfire.
The recent problems associated with elevator
tab failures and fatal crashes of a certified twin engine commuter
airplane have now resulted in regulation changes requiring dual elevator
trim tab controls for each tab surface running from the irreversible
control point to the tab itself. This eliminated the simple push-pull
cable running from the cockpit to the elevator tab, as used successfully
on my TEAL Amphibian certified in 1969, unless the irreversible control
(screwjack or worm gear) system is located in the cockpit and dual
push-pull controls or twin pushrods are run to each elevator tab surface.
The added weight, cost, and installation problems resulting from this
regulation are obvious as well as completely unnecessary for any small,
comparatively low-powered, low-speed homebuilt or certified
airplane.
To comply with current certification regulations we have
reverted to a separately controlled spring loaded elevator trim control
system on the SEAFIRE in addition to the primary elevator control system.
However, we must flight demonstrate that no flutter is present in this
system up to design dive speed with the springs in or out of the trim
system.
So the homebuilt TROJAN has a trim tab on the elevator
controlled by a simple push-pull cable as used on the TEAL, while the
SEAFIRE has a separate spring trim system deflecting the entire elevator
and requiring considerable flight development time and cost to satisfy
current TC criteria.
Drop test requirements may be necessary and
desirable to assure structural integrity of the landing gear and related
support structure for FAA certified production aircraft, if for no other
reason than to protect the manufacturer from product liability claims.
However, homebuilt aircraft can realize about the same degree of
structural safety by designing each main gear to independently support the
airplane gross weight times a 1.50 factor of safety. This means that each
main wheel, strut, and attachment structure for a 1500 pound airplane
should be designed for a 1500 pound vertical load times a 1.50 factor of
safety = 2250 pounds per side.
The nose or tail gear should be
similarly designed for at least twice the largest static loading
calculated times 1.50 factor of safety. In fact, 3 times the largest
static load times a 1.50 factor of safety would be more desirable. It is
also necessary to consider the aft and side load conditions of Part 23.471
and subsequent landing gear load criteria.
Since the desired level
of ride softness will influence the final shock absorber design as based
upon taxi tests over relatively rough terrain, the above design loads can
be safely used for homebuilt aircraft to eliminate the costly and time
consuming procedure of drop testing a landing gear to prove it will stay
together.
c) All landplanes, whether FAA certified or homebuilt,
should incorporate a throttle actuated
warning horn to indicate
retractable landing gear up at low power setting. Part 23 now requires
that this horn also be activated by flap deflection. While an
excellent feature for landplanes, a
blowing horn is not welcome on
any amphibian since the pilot might instinctively lower the
gear if
the horn blows during a water landing approach. While check-off lists
should preclude
gear down water landings. Human nature vs. checklists
being what it is, we can expect to con-
tinue having gear-down water
landing amphibian accidents under the best of conditions even
without the assistance of a blowing horn.
The FAA has now
required that a horn warning system must be used on the SEAFIRE
Amphibian if we desire certification. Although we have developed an
interconnected set of
Push button lights which will eliminate the horn
for water work if they are operated in proper
Sequence, we are
nevertheless protesting this requirement since we know from long
experience
That it will result in one or more water landing accidents
before many amphibious flight hours have been realized.
So the
TROJAN Amphibian has gear position lights clearly marked for water or land
operation with no horn, while the SEAFIRE has a rather complex set of
landing gear position lights, a throttle and flap position interconnected
warning horn, a horn cut-out button system for water landings and quite
likely a built-in accident potential. The added complexity and cost of
the certified system are again obvious, to say nothing of the safety
aspects. Of course, new owners
can disconnect the horn if desired
once they leave the factory and probably will do so.
The recent Part
23 requirement for a lightning strike analysis study is particularly
annoying since it is difficult, if not impossible, to find any two people
in the FAA regional engineering offices who either know what is desired or
agree on what would be acceptable.
The only constructive
suggestions I have is that a metal airplane is apparently safer than a
plastic one since metal can conduct a strike charge on the skin surface;
and that fuel contained near the wing root is safer than that stored in a
tip tank because a wing tip is frequently the initial strike point. Also
note that due to propeller strike fuel stored in the fuselage is not as
safe as that located near the wing root.
On a semi-humorous or
semi-serious note, I suggest that we placard personal aircraft to prohibit
operating in thunderstorm conditions the same as we do for icing, an so
save the cost of the entire frustrating lightning analysis report. After
all, no responsible pilot wants to fly into icing or thunderstorm areas,
so why cant a placard statement approved for one condition also apply to
the other? This comment was not appreciated by the FAA.
Probably
the most useless requirement of all for small aircraft is the structural
life fatigue analysis requirement introduced during the 1970s. Over the
past 40 years many personal aircraft have provided continuous operation
without first having been subjected to a thorough and time consuming
analysis of their probable safe operating life. What these survivors have
received, however, is fairly excellent maintenance - which does not enter
into the fatigue life analysis report at all.
If an airplane is
designed to the usual 3.8g normal category limit load factor (5.7g
ultimate)
the entire structure will be at quite a low stress level
during normal operation. It can be shown that if the most versal of +2g
once per minute let us say along the beam can angles, 2014T6 extruded
material would have a life expectancy in excess of 15,000 hours of
airplane operation. Critical 4130 steel components would experience a
considerable greater life expectancy.
Considering 300 hours annual use,
an abnormally high rate for a small personal airplane, such an airframe
would have a life expectancy of 50 years; if not properly maintained, the
structure would corrode away long before any critical fatigue could
occur.
Since all small personal use aircraft are designed to at least
3.8g limit load, why go through a lengthy and detailed analysis which has
a foregone conclusion of extreme life expectancy? The engineering time and
cost incurred in this exercise might be better employed in refining the
detail design to permit more economical production.
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Although none of the above required procedures is
prohibitive by itself, they will collectively add about 2500 engineering
and shop hours plus $65,000 to the development and certification costs of
our SEAFIRE Amphibian without making it one bit safer to fly, cheaper to
produce, or less costly to operate all of which factors are desperately
needed if we are to maintain or expand the current personal airplane
market. Now let us consider the more important design aspects of Part
23.
The Favorable Effects of Using Part 23
The more positive
aspects of Part 23 can be applied to homebuilt aircraft designs with
considerable benefit as regards both structural integrity and improved
flight handling characteristics. And in my opinion, at least, these
requirements should be closely followed in the following
areas.
Development of applied loads
Relieving load
considerations
Structural analysis and static tests
Ground Flutter
tests
Fuel tank and delivery line considerations
Flammability
criteria for interior fabrics
Basic flight stability
tests.
While other Part 23 criteria are important, particularly
those covering powerplant installation and instrument systems, the above
seven items deserve some detail consideration.
The Applied Loads report
is the basis for the entire airplane design. This includes determination
of preliminary aerodynamic performance to establish surface load
distributions which in turn provide wing, tail, and fuselage bending loads
as well as those for the landing gear, engine mount, and control
systems.
When applied loads were developed for the TEAL Amphibian
in 1967 the complete report required 40 pages. The SEAFIRE Amphibian
Applied Loads Report of 1980 consists of 127 pages for an essentially
similar although larger single engine airplane. This says something about
the growth in Part 23 criteria complexity during those 13
years.
The important point for homebuilt design being that at least
a basic applied loads analysis should be completed for every new airplane
design based upon the highest design airspeed (largest powerplant) and
gross weight expected for that model. While this approach will tend to
penalize aircraft of the same model with smaller powerplants, the
alternative is to prepare a separate load analysis for each powerplant
gross weight configuration.
Once run through, the format for
determining applied loads is fairly routine although somewhat time
consuming. Regardless, I do not understand how any airplane can be
properly designed unless applied loads have been developed and used for
detail structural analysis. The criteria of FAR Part 23 are an excellent
reference source for determining required basic loads.
While not
noted in Part 23 in so many words, current FAA engineering procedure
requires that airplane wing analysis be conducted at gross weight with no
fuel in the wings, even if fuel tanks are located within the wing envelope
(as I believe they should be). IT seems that a few years ago someone
loaded an airplane to gross weight by storing heavy machinery within the
fuselage. Each wing tank contained about 10 gallons of fuel in tanks
designed to hold 60 gallons per side. The airplane was destroyed in
turbulent air, with the structural failure determined to have occurred
because the full 60 gallon load per side had been considered as wing
bending relieving load at gross weight. In other words, at gross weight
loading the download of 60 gallons of fuel plus the wing weight both at
3,8g limit load had been used to reduce the total wing lift bending load
at gross weight. As a result, when considerable weight was added at the
fuselage centerline instead of in wing fuel tanks, the wing bending load
in turbulent air exceeded design conditions causing each wing panel to
fail in the root area. The design of a wing structure relieved only by
wing and main landing gear weight, assuming the main gear is wing mounted,
is both conservative and fairly new. (Relieving weight is discussed on
pages 184 and 185 of "Design for Flying".) This approach seems overly
conservative, but has been followed in designing the SEAFIRE Amphibian and
is worth doing if the resulting structural weight penalty is not too
great. I personally believe at least one quarter of the wing fuel weight
distribution could be considered as relieving load when developing wing
bending loads at gross weight. However, if this procedure is followed for
wing analysis the pilots manual should indicate that all wing tanks must
be at lease one-quarter full at airplane gross weight
. d) Detail
Structural analysis should be completed for all major portions of the
airframe such as: wing
main beam or wing beams; major wing ribs
particularly those taking flap, aileron, or landing gear loads; tail
surface beams and their fuselage attachments; fuselage structure as based
upon engine mount, wing connection, and tail surface loads; seat, belt,
and harness installations; engine mount; and the control system.
In addition, these assemblies should be static tested at least up
to limit design load of 3.8g for the airframe components, and up to pilot
applied loads required by FAR Parts 23.391 - .405 for the control systems.
While control sticks and rudder pedals may be loaded outside the airplane
in test fixtures, the complete control system must be tested with the
surfaces loaded in accordance with applied loads calculations and as
installed in the airframe ready for flight.
This last test is
particularly critical and important; when operated under load, the control
surface should move through an arc of a least one-half their design
deflection. This means that an elevator surface designed to operate at a
maximum of 20 degrees upward movement, must be capable of moving up at
least 10 degrees under test load. If this displacement is not realized,
either the control cable size should be increase or some critical brackets
must be reinforced to reduce their deflection under load.
d)
Ground flutter vibration test may be simply performed by rigidly mounting
an eccentric counter-weight driven by a small variable speed motor onto
the wing main beam in the vicinity of either wing tip. The object of this
investigation is to determine structural vibration frequencies that bear a
relationship to powerplant rpm in flight or at ground idle speeds. This
set of tests should be conducted prior to flight and need not be
elaborate; however, it is necessary to have a correct reading of
counterweight rpm at the various test speeds.
For example, if the
tail surface or wing panel start to shake violently at 600 rpm of the
counterweight, this indicates the possibility of structural vibration at
600 engine rpm on the ground or at 1200, 1800, or 2400 engine rpm in
flight, etc. The rotating weight need not exceed 6-8 ounces for small
aircraft and should be mounted on an 8-10 inch radius arm in such manner
that it may be moved in and out along the arm to vary the vibratory input
load at a given rotation speed. This test should be conducted at a number
of different counterweight rotational speeds and radius positions to
thoroughly scan the structural vibration spectrum. Air pressure in all
tires should be reduced to 10 pounds or less to isolate the airplane from
stabilizing effects of ground surface friction.
During the course of
this test you will be surprised to see how metal panels will suddenly
"take off" and vibrate with considerable noise and vigor. If necessary,
local stiffeners may be added to dampen this vibration which can cause
considerable noise in flight. Any major structural assembly vibration such
as wing panel or surface flutter, tail surface oscillation, or control
system parts banging about should be further investigated and corrected by
reinforcement or rebalancing prior to flight.
The most recent Part
23 regulations really require that all fuel tanks and lines be eliminated
from the passenger compartment. Certainly large volumes of fuel should be
located where it cannot splash over anyone during a forced landing or
crash. This precludes locating fuel tanks ahead, behind, or alongside the
pilot or passengers in the cabin or cockpit area.
I have recently
been acting as a technical consultant for some aircraft accident cases, in
one of which there probably would have been no serious injuries following
a crash from low altitude if a fuel tank had not been located forward of
the front passenger in a tandem cockpit airplane. This crash ruptured the
fuselage fuel tank which covered the front passenger with fuel and also
splashed the pilot in the rear cockpit. The fuel ignited from exhaust
sparks, killing the occupant of the forward cockpit and badly burning the
pilot who survived a painful recovery.
Keeping fuel lines out of
the cockpit, in the strictest interpretation of current regulations,
requires installation of a remotely located selector valve controlled
either by extension shaft drive, cable and chain control, or electrically
operated solenoid or gate valves. If the selector valve cans be located
under the cabin floor with the selector/indicator handle mounted above the
floorboards in a position readily accessible to the pilot, the fuel lines
leading to and from the selector valve can be considered as being out of
the cabin area provided they are led below floor level for the entire
cabin/cockpit area.
Such an installation is satisfactory for
landplane design, but is unacceptable for flying boat type amphibian since
the fuel system cannot be preflight drained when on the water. To overcome
this, both the TROJAN and SEAFIRE amphibians have remote controlled
electrically operated gate type selector valves located behind the wing
main beam and aft of the passenger compartment. Fuel drains are located on
the side of the hull under the wing in a position accessible from the
cabin when afloat.
Another fuel line isolation solution considered
acceptable by FAA engineering is that of encasing any cockpit or cabin
area fuel lines inside a larger tube which is sealed at either end and has
an overboard vent to drain any leaking fuel that might occur. This seems
to be a rather heavy and costly solution, and one less safe than a
remotely operated selector valve eliminating all fuel lines within the
cabin/cockpit area.
We increasingly hear about deaths from hotel
fires associated with toxic gasses produced by burning drapes, carpets,
and similar synthetic fabrics and foams. The same situation can arise from
inflight fires unless all plastics used in the cabin area are certified as
being flash and flame resistant. The FAA has established fire protection
criteria in FAR Part 23.853 and subsequent paragraphs. These guidelines
can be readily applied to homebuilt aircraft with little additional cost
simply by ordering materials that meet the requirements of FAA Flight
Standards Service Release No. 453, Attachment A. There are many sources
capable of supplying and certifying materials to this specification, so
elaborate test procedures are not necessary for each
application.
The FAA recommendation for fire resistant sleeves on
all flexible fuel and oil lines located in the engine compartment ahead of
the firewall should also be observed for additional safety (inflight time)
in the event of engine fire does occur.
Basic flight stability test
specified in Part 23 not only provide a more uniform set of handling
characteristics for all aircraft, but also establish safe flight limits
for center of gravity travel at gross weight.
Assuming that a new
design has acceptable lateral/directional and pitch stability as
determined by a competent test pilot following Part 23 recommended flight
test investigation procedures, the aft c.g.limit should be carefully
determined by loading the airplane to normal gross weight c.g. position
and then gradually moving the weight further aft in small increments of
one percent mean aerodynamic chord (see "Design for Flying", pages 92-95,
116-117, and 140). This is done at a constant cruising airspeed with
increasingly further aft gross weight c.g. locations until a position is
reached where very slight stick force is required to pitch the airplane
nose-up from a trimmed attitude. This is at or near the point of neutral
pitch stability c.g. location.
To provide a greater margin of
safety for flight operation the permissible aft c.g. limit should be
established at least one percent MAC forward of this neutral position; low
or not stick force is an open invitation to approach stalls, low altitude
spins on final, high speed stalls shedding wings, etc. (See "Design for
Safety", page 31.)
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I trust these
comments have been of interest; if followed, I know they will contribute
toward improved safety of future bomebuilt aircraft. In that regard, I
have recently been investigating two landplane configurations that satisfy
these design criteria while also providing simplified flight
characteristics combined with efficient operation and short field
capability. These models are briefly discussed in the following section
and are shown in profile Figures 3 and 4.
Future Homebuilt
Designs
I personally believe the time has come for homebuilders to be
offered a choice of aircraft which have been designed essentially in
accordance with FAR Part 23 requirements. This approach encompasses not
only a well designed structure but also safe flight characteristics. To me
this means a rugged, easily handled airplane capable of operating from
small airports but one incapable of stalling or spinning. Proposed
landplane Models TA20 and TA22 will satisfy these combined criteria for
the first time in an airplane available for homebuilt
construction.
Accompanying Figures 3 and 4 show these designs in
preliminary side views. Model TA20 is a two-place, 110 hp Lycoming engine
powered airplane capable of cruising 125 mph at 2/3 power (20
miles/gallon), while the larger TA 22 is a four-place model powered by a
160 hp Lycoming engine, cruising at 135 mph at 2/3 power (approx. 16
miles/gallon).
Each model will incorporate the following
features:
Complete kit for all-metal construction with skins
predrilled for ease of assembly.
Designed essentially in accordance
with FAR Part 23 requirements
Structure will be both analyzed and
static tested per Part 23
Full span flap/aileron (flaperon)
system
Coordinated two-control system with restricted elevator to
prevent wing stall
Speed brakes for precise descent control operated
by a foot pedal which also applies main
Wheel brakes when fully
depressed
Stick control
Tricycle landing gear with steerable
nosewheel
Excellent rate of climb (approx. 900 feet/minute)
The
two-place Model TA20 will be offered first if sufficient builder interest
warrants construction and development, followed closely by the four-place
TA22. Let me know your interest in these new aircraft by writing to me
at:
Thurston Aeromarine Corporation
16 Jericho Drive
Old Lyme, CT
06371
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