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BY DAVID B. THURSTON EAA Sport Aviation Oct. 1982

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 can’t 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 pilot’s 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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