SECTION 15 FINAL INSPECTION AND FLIGHT TEST
RV AIRCRAFT
15-21
SEC 15r8 12/23/10
Similarly, assuming that the atmosphere were perfectly stable, when flying at 200 mph, a pilot error of 1/2 degree
pitch attitude will cause a 150 fpm climb or descent rate and several mph speed variation. Thus, great care must be
taken to find smooth air and fly precisely in order that truly accurate speeds be recorded. It is a good idea to fly the
speed box more than once to check consistency and obtain averages if speed variation occur.
LIMIT TESTING
Limit testing of a homebuilt, particularly a high performance one such as the RV, is an endeavor to be approached
with caution and preparation. What the pilot is doing is challenging the airframe to withstand the limit loads he is im-
posing on it, or in a sense, daring it to fail. Most homebuilder/pilots are not daredevils and would just as soon not do
limit testing. However, as it is the best available means of verifying design limits, it must be done if all future flights
are to be made with confidence. With proper preparation, limit testing need not be as frightening and dangerous as it
might appear. Particularly during this phase of testing, the pilot should wear a parachute and familiarize himself with
its operation. Also, emergency egress of the airplane should be reviewed and memorized. Limit testing should be
done at altitudes of at least 5000 ft. above ground, preferably around 8,000 ft. Along with careful planning, altitude
can be a lifesaver. While the thought of structural failure or loss of control is not at all appealing, it is far better that it
be encountered during controlled testing than under conditions where no options exist (low altitude, no parachute,
etc.). By assuming and preparing for the worst, limit testing can be done with reasonable confidence. Flight testing
of the RV prototypes proved to be routine and uneventful. With thoughtful construction and preparation, testing of
homebuilt RVs should be the same. Limit testing categories include FLUTTER TESTING, G-LOAD TESTING, and
SPIN TESTING. (Spin testing is also classified under Stability testing, so has been included in that section of this
chapter)
FLUTTER TESTING
Flutter in an aircraft structure results from the interaction of aerodynamic inputs, the elastic properties of the structure,
the mass or weight distribution of the various elements, and airspeed. The word “flutter” suggests to most people a
flag's movement as the wind blows across it. In a light breeze the flag waves gently but, as the wind speed in-
creases, the flag's motion becomes more and more excited. It is easy to see that if something similar happened to
an aircraft's structure the effects would be catastrophic. In fact, the parallel to a flag is quite close.
Think of a primary surface with a control hinged to it (e.g., aileron). Imagine that the aircraft hits a thermal. The initial
response of the wing is to bend upwards relative to the fuselage. If the center of mass of the aileron is not exactly on
the hinge line, it will tend to lag behind the wing as it bends upwards.
In a simple, unbalanced, flap-type hinged aileron, the center of mass will be downward. This will result in the wing
momentarily generating more lift, which will increase its upward bending moment and its velocity relative to the fuse-
lage. The inertia of the wing will carry it upwards beyond its equilibrium position to a point where more energy is
stored in the deformed structure than can be opposed by the aerodynamic forces acting on it.
The wing “bounces back” and starts to move downward but, as before, the aileron lags behind and is deflected up-
wards this time. This adds to the aerodynamic down force on the wing, once more driving it beyond its equilibrium
position and the cycle repeats.
At low airspeeds, structural and aerodynamic damping quickly suppresses the motion but, as the airspeed increases,
so do the aerodynamic driving forces generated by the aileron. When they are large enough to cancel the damping,
the motion becomes continuous. Further small increases in airspeed will produce a divergent, or increasing, oscilla-
tion, which can quickly exceed the structural limits of the airframe. Even when flutter is on the verge of becoming
catastrophic, it can still be very hard to detect. What makes this so is the high frequency of the oscillation which is
typically between 5 and 20 HZ (cycles per second). It will take only a very small increase in speed to remove what
little damping remains and the motion will become divergent rapidly.
Flutter testing of factory prototypes has resulted in establishing a NEVER EXCEED SPEED (Vne) of 210 statute mph
for the RV-3,4 and RV-6/6A, 230 statute mph for the RV-7/7A/8/8A. and 210 statute mph for the RV-9A. This speed
was determined through flutter testing at a speed of 20 mph above Vne. (FAA certification criteria require flutter test-
ing up to Vne plus 10% or about 20 mph) The flutter testing performed consisted of exciting the controls by sharply
slapping the control stick at various speed increments up to this level. Under all conditions, the controls immediately
returned to equilibrium with no indication of divergent oscillations indicative of flutter. This testing was performed on
factory prototype aircraft, and the flutter free flight operation of subsequent amateur built RVs has substantiated pub-
lished Vne.
The “slap-the-stick'' method of exciting the controls for flutter testing is potentially dangerous and requires a very
skilled pilot trained to recognize the subtle control responses which indicate the onset of flutter. For this reason, it is
suggested that amateur builders do not perform flutter testing of their RVs. Rather, the airplane should be con-
structed in strict conformity to the plans with particular attention paid to the control system--- trailing edge radii, skin
stiffness, control linkage free-play, and static balance in particular. Maintaining conformity with the prototype (plans)
will provide an adequate level of assurance against control surface flutter. Any design changes to the control sur-
faces, control system, or primary structure could invalidate the testing which has been done, and require that testing
be re-accomplished.
SECTION 15 FINAL INSPECTION AND FLIGHT TEST
RV AIRCRAFT
15-22
SEC 15r8 12/23/10
G-LOAD TESTING
As with flutter testing, G-load testing should be conducted systematically, progressing gradually to higher and higher
levels. 6 G's is the highest level recommended in testing. This is the maximum load the structure is designed to be
able to withstand indefinitely. While the actual calculated breaking strength is 9 G's, the structure is designed to with-
stand this load for only 3 seconds. Approaching this load level could permanently weaken the structure even if failure
does not occur. The margin between 6 and 9 G's is reserved to compensate for the effects of airframe deterioration
through aging, fatigue, material flaws, or construction errors. G-loads of over 6 should never intentionally be applied
to an RV structure.
MAXIMUM G-LOAD:
The structure of the RV-4 and RV-6/6A have been designed to withstand aerobatic design loads of plus 6 Gs and
minus 3 Gs at an aerobatic gross wt. of 1375 lb, or 1600 lb for the RV-7/7A/8/8A. The RV-9/9A structure had been
designed to withstand utility design loads of +4.4/-1.8G at a utility gross weight of 1600 lbs. Flight testing to the posi-
tive G limit can be done by putting the airplane in a tight turn and applying elevator back pressure. Do this progres-
sively; increasing the load by 1G increments until the load limit is reached. Between each loading acceleration, relax
and look over the airplane. Move the controls to assure that everything is normal.
For operational gross weights above this figure, aerobatic maneuvers should not be performed. This also assumes
that the RV was built in strict conformity with the plans. Any variation in materials used, dimensions of primary struc-
tural parts, or workmanship standards, can cause a loss of strength and cause the actual limit load to be less than the
design limit load.
The 3 G design limit for negative loads also has a built-in 50% margin. Thus the breaking strength would be -4.5 Gs.
If the RV being tested is equipped with inverted fuel and lubrication systems, negative G testing should also be done.
A parachute should be worn while conducting load testing.
MANEUVERING SPEED:
134 mph statute for the RV-4 and RV-6/6A, 142 for the RV-7/7A8/8A, 118 for the RV-
9/9A. By definition, maneuvering speed is the maximum speed at which full and abrupt controls can be applied. It is
also the minimum speed at which limit G-load can be produced. Thus, at any speed in excess of this, full control ap-
plication could result in G-loads in excess of design limits. The maneuvering speed is function of clean (no flap) stall
speed. For utility category aircraft like the RV-9/9A, it is 2.1 (the square root of 4.4) x stall. For aerobatic category
aircraft, it is 2.45 (the square root of 6) x stall. Because RVs have low stall speeds, maneuvering speed is low rela-
tive to cruise and Never Exceed Speed.
Based on the same formula used to determine maneuvering speed, full control application at Vne would produce a
G-load of about 15. From this it should be very obvious that at any speed above maneuvering speed, the pilot be-
comes the limiting factor: he can impose destructive loads on the structure through excessive control application.
Because of its high ratio of top speed to stall speed, the RV is more susceptible to pilot-induced over-
stresses than are most other contemporary aerobatic airplanes.
GROSS WEIGHT:
See Section 14.
AEROBATIC OR UTILITY GROSS WEIGHT:
See Section 14.
FLAP SPEED:
On the RV-4/6/6A/7/7A/8/8A, 110 Statute for 20º and 100 mph statute for full 40º flap deflection. On
the RV-9/9A, it is 100 smph for 15º and 90 smph for 32º.
