(4) Quickly reverse the input to nose up to bring the pitch attitude back to trim attitude. As the pitch attitude reaches trim attitude, release the stick (but guard it). The aircraft with positive dynamic longitudinal stability will oscillate briefly about the trim attitude before stopping at the trim attitude position.
(5) To test the aircraft for positive DYNAMIC longitudinal stability (long period), begin from trimmed, straight and level flight. Without re-trimming, pull (or push) the stick to a speed about 5 mph/knots off trim and release the stick.
There is no need to stabilize at the new speed. Expect the aircraft to oscillate slowly about the trim airspeed a number of times before the motion dampens out.
If there is significant friction in the control system, the aircraft may settle at a speed somewhat different from the original trim speed.
(6) If the amplitude increases with time, the dynamic longitudinal stability is negative or divergent. This is not necessarily dangerous as long as the rate of divergence is not too great. It does mean, however, the aircraft will be difficult to trim and will require frequent pilot attention.
(7) An aircraft with ‘‘NEUTRAL’’ dynamic longitudinal stability (long period) will continue to oscillate through a series of increasing/ decreasing airspeeds and never return to the original trim airspeed.
f. Lateral-directional Stability Control Tests.
Lateral (Dihedral Effect) and directional stability tests are to determine if the aircraft can demonstrate a tendency to raise the low wing in a sideslip once the ailerons are freed. They also determine if the rudder is effective in maintaining directional control.
CAUTION: This test may impose high flight loads on the aircraft. Do not exceed the design maneuvering speed or any other airspeed limitation.
(1) To check lateral and directional stability, the aircraft should be trimmed for level flight at a low cruise setting and an altitude above 5,000 feet AGL. Slowly enter a sideslip by maintaining the aircraft’s heading with rudder and ailerons. The aircraft should be able to hold a heading with rudder at a bank angle of 10 degrees or the bank angle appropriate for full rudder deflection. The control forces and deflection should increase steadily, although not necessarily in constant proportions with one another (in some cases, rudder forces may lighten), until either the rudder or the ailerons reach full deflection or the maximum sideslip angle is reached.
(2) At no time should there be a tendency toward a force reversal, which could lead to an overbalance condition or a rudder lock.
(3) Release the ailerons while still holding full rudder. When the ailerons are released, the low wing should return to the level position. Do not assist the ailerons during this evaluation.
(4) To check static directional stability, trim the aircraft at a low cruise setting above 5,000 feet AFL. Slowly yaw the aircraft left and right using the rudder. Simultaneously the wings should be kept level by using the ailerons. When the rudder is released, the aircraft should tend to return to straight flight.
g. Spiral Stability. This is determined by the aircraft’s tendency to raise the low wing when the controls are released in a bank. To test for spiral stability, apply 15 to 20 degrees of bank either to the left or right, and release the controls. If the bank angle decreases, the spiral stability is positive. If the bank angle stays the same, the spiral stability is neutral. If the bank angle increases, the spiral stability is negative. Negative spiral stability is not necessarily dangerous, but the rate of divergence should not be too great or the aircraft will require frequent pilot attention and will be difficult to fly, especially on instruments.
NOTE: Friction in the aileron control system can completely mask the inherent spiral characteristics of the airframe.
52 5/24/95AC 90-89A SECTION 4. A WORD OR TWO ABOUT FLUTTER ‘‘Stay up on the edge of your seat.’’ Scott Crossfield, Test Pilot
1. OBJECTIVE. To understand the causes and cures of the condition known as flutter.
2. DESCRIPTION. Flutter in an aircraft structure is the result of an interaction between aerodynamic inputs, the elastic properties of the structure, the mass or weight distribution of the various elements, and airspeed.
a. To most people, the word ‘‘flutter’’ suggests a flag’s movement as the wind blows across it. In a light breeze, the flag waves gently but as the wind speed increases, the flags motion becomes more and more excited. It takes little imagination to realize if something similar happened to an aircraft structure, the effects would be catastrophic. The parallel to a flag is appropriate.
b. Think of a primary surface with a control hinged to it (e.g., an aileron). Imagine that the airplane hits a thermal. The initial response of the wing is to bend upwards relative to the fuselage.
c. 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.
d. In a simple, unbalanced, flap-type hinged control, the center of mass will be behind the hinge line and the inertial lag will result in the aileron being deflected downwards. This will result in the wing momentarily generating more lift, increasing its upward bending moment and its velocity relative to the fuselage. 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.
e. The wing ‘‘bounces back’’ and starts to move downward but, as before, the aileron lags behind and is deflected upwards this time. This adds to the aerodynamic down force on the wing, once more driving it beyond its equilibrium position and the cycle repeats.
f. Flutter can happen at any speed, including take-off speed. At low airspeeds, however, structural
53 and aerodynamic damping quickly suppress the flutter 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.
g. Further SMALL INCREASES will produce a divergent, or increasing oscillation, 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 causes this is the high frequency of the oscillation, typically between 5 and 20 Hz (cycles per second).
It will take but a small increase in speed ( 1⁄4 knot or less) to remove what little damping remains and the motion will become divergent rapidly.
h. Flutter also can occur on a smaller scale if the main control surface has a control tab on it.
The mechanics are the same with the tab taking the place of the aileron and the aileron taking the place of the wing. The biggest difference are the masses involved are much smaller, the frequencies much higher, and there is less feed-back through the control system. This makes tab flutter more difficult to detect. The phenomenon known as ‘‘buzz’’ is often caused by tab flutter. Since flutter is more prevalent at higher speeds, it is not recommended that the flight test plan call for high speed runs within 10 percent of red line.
i. What can be done about it? Having described how flutter happens, the following suggestions should help reduce the possibility of it happening to the amateur-builder’s aircraft:
(1) Perform a mass balance of all flight controls in accordance with the designer/kit manufacturer’s instructions.
(2) Eliminate all control ‘‘free play’’ by reducing slop in rod end bearings, hinges, and every nut and bolt used in attaching flight controls.
(3) Ensure that all rigging and cable tension is set accurately to the design specifications using a calibrated cable tensiometer.
(4) Re-balance any flight control if it has been repaired, repainted, or modified in any way.
NOTE: If the pilot experiences flutter, or believes he did, reduce power immediately and land as soon as possible. Do not attempt further flight until the aircraft has been thoroughly inspected for flutter induced damage. This inspection should include all wing/tail attach points, flight controls, their attach points/hinges, hardware, control rods, and control rod bearings for elongated bolt/rivet holes, cracks, (especially rod end bearings) and sheared rivets.
54 5/24/95AC 90-89A SECTION 5. SPINS ‘‘Go from the known to the unknown -- slowly!’’ Chris Wheal, Military Test Pilot
1. OBJECTIVE. To determine if spin testing is required.
NOTE: All FAA spin tests for type certification require a spin chute attached to the aircraft. Even though amateur-built aircraft have no such certification requirement, use of a spin chute during testing should be considered.
2. CAUTION.
a. If the manufacturer/designer of the aircraft has not demonstrated satisfactory spin characteristics and safe recovery, avoid all types of high angle of attack flight testing and placard the aircraft: ‘‘spins prohibited.’’
b. If the prototype aircraft has satisfactorily demonstrated spin recovery and the builder’s aircraft is identical to the prototype aircraft, the pilot may confirm the aircraft will recover promptly from inadvertent spin entries. Further tests to prove that the aircraft will recover from a fully developed spin (three turns or more) are not necessary unless the aircraft is designed for, and will be routinely flown in, aerobatic flight.
c. During all spin tests, it is strongly recommended that the pilot wear a parachute and that a quick release mechanism to jettison the canopy or door be installed. If the pilot is unable to exit the aircraft because of the design restraints, it is recommended that intentional spins not be conducted even though the design has successfully demonstrated spin recovery.
d. If any modifications or alterations have been made to the airframe’s original design or configuration (e.g., adding tip tanks or fairings), it is not safe to assume that the aircraft still has the same spin recovery characteristics as the prototype aircraft. Spins in a modified aircraft should not be attempted without consulting a qualified test pilot and/or flight test engineer.
e. The pilot who conducts the spin tests should have experience in entry into and recovery from fully developed spins, preferably in makes and models similar to the aircraft being tested. If the pilot needs
55 additional experience, aerobatic training with an emphasis on spins from a qualified instructor is highly recommended.
3. PLANNING THE FLIGHT. At this point, nearly all the preparatory work for spin testing has been accomplished. Planning the next flight should be identical to planning for the first flight through stalls. IT IS EXTREMELY IMPORTANT THAT
THE CENTER OF GRAVITY OF THE AIRCRAFT
IS AT THE FORWARD CG LIMIT AND ANY
BALLAST USED SHOULD BE SECURELY ATTACHED TO THE AIRCRAFT.
a. The aircraft should be tested with landing gear (if applicable) and flaps in the up position. The pilot’s minimum entry altitude for these tests should be no less than 10,000 feet AGL with the cockpit secured.
NOTE: The following procedure is one way, but not the only way, of conducting a spin test and executing a recovery. Non-conventional aircraft may require significantly different spin recovery control applications.
The pilot should evaluate these procedures and determine if they are compatible with the aircraft before attempting any spin testing.
b. The basic technique used to get a clean spin entry is to continue to reduce airspeed at about a 1 mph/knot a second rate in level flight, carburetor heat on, and the power at idle.
(1) As the aircraft stalls, APPLY FULL RUDDER in the desired spin direction, followed immediately by full aft movement of the control stick keeping the ailerons neutral.
(2) The transition from a horizontal to a vertical flight path takes approximately three or four turns and is referred to as the incipient stage of the spin.
(3) During the incipient spin, the dynamic and inertia forces have not achieved equilibrium.
Many aircraft can recover from the incipient spin phase, but may not be able to recover from a steady spin.
(4) The normal spin recovery technique is to apply full rudder opposite to the direction of yaw (check the turn needle). Move the control stick smoothly and fairly rapidly forward towards the instrument panel until the rotation stops.
(5) Quickly center the rudder and ease out of the dive. Do not attempt to pull up too rapidly because the structural limits of the aircraft can easily be exceeded, or the aircraft can stall again. Recover from the first deliberate spin after a half a turn.
c. If the aircraft is not built for aerobatics, no further spin testing is required, It is recommended the instrument panel be placarded ‘‘SPINS PROHIB- ITED.’’
d. If further spin testing is required, it is strongly recommended the services of a professional flight test pilot be used.
56 5/24/95AC 90-89A SECTION 6. ACCELERATED STALLS ‘‘Does it pass the Common Sense test?’’ U.S. Air Force, Thunderbird
1. OBJECTIVE. To further explore the stall characteristics of the aircraft.
a. An accelerated stall is not a stall reached after a rapid deceleration. It is an in-flight stall at more than one g, similar to what is experienced in a steep turn or a pull up.
NOTE: Do not attempt this or any other extreme maneuver unless the designer or kit manufacturer has performed similar tests on a prototype aircraft identical to the amateur-builder’s aircraft.
b. The two standard methods for accelerated stalls are the constant g (constant bank) and constant speed (increasing bank). Most preferred of the two is the constant bank method in which the airspeed is decreased and the angle of bank is held constant, until the aircraft stalls. It is the most preferred because the potential violence of any accelerated stall is largely governed by the increasing g load and airspeed.
c. As with every test, plan the sequence of events. Start with small bank angles -- 30 degrees will produce 1.15 g. Decelerate slowly, ball in the center, do not over control. Work up incrementally to a two g, 60 degree bank.
d. The aircraft does not have to develop a deep stall each time. The pilot needs only to record the airspeed and bank angle in which the aircraft hits the pre-stall buffet. Recover by adding power and reducing the angle of bank.
57 CHAPTER 6. PUTTING IT ALL TOGETHER: 36 HOURS TO —————?
‘‘Beware of false knowledge; it is more dangerous than ignorance.’’ George Bernard Shaw SECTION 1. MAXIMUM GROSS WEIGHT TESTS
1. OBJECTIVE. To develop aircraft performance data across the weight and CG ranges.
a. Up until this point, all tests have been performed well below the test aircraft’s maximum gross weight, with the possible exception of single seat aircraft designs. A complete series of flight tests at maximum gross weight from stalls, rates of climb, angles of climb, stability, retraction tests, slow flight, through accelerated stalls should be investigated.
b. These tests should demonstrate that the aircraft has been successfully flown throughout the CG range, and will operate in and at the full range of aircraft weights from minimum to full gross weight.
The findings should be documented in the aircraft’s flight manual.
c. Each phase of the testing should be done slowly, incrementally, with the same careful attention to detail that should characterize all the flight testing.
d. Increases in the aircraft weight should be done in a series of steps. Usually, 20 percent increments of the maximum payload (e.g., sandbags, lead shot) are added in the aircraft to simulate passengers or baggage weight. The pilot should carefully weigh and secure the ballast. A new weight and balance and CG location must be worked for each new increase in weight. Stop testing when the aircraft’s maximum gross weight is reached.
e. The testing up to this point has been done at, or near, the forward CG limit. During these tests, the CG should be slowly, but progressively, moved aft between each test flight. Limit the change to the CG range to about 20 percent of the range. Again the pilot should weigh the ballast and work a new weight and balance for each flight. With each CG change the aircraft longitudinal static stability and stall characteristics should be carefully evaluated by using the same technique discussed earlier. Stop testing when the designer’s or kit manufacturer’s aft CG limit is reached.
f. If the aircraft develops either a neutral or negative longitudinal stability problem, or the aircraft displays unsatisfactory stall characteristics at any CG location being tested, STOP FURTHER TESTING!!
g. These tests should confirm the designer’s aft CG limit or establish the last satisfactory aft CG location. If the aft CG range is not satisfactory, consult with the kit manufacturer, aircraft designer, or a flight test engineering consultant.
h. The pilot should avoid the temptation to take a live ballast weight up for a ride for three reasons:
(1) The aircraft has not been proven safe for the higher gross weights.
(2) The pilot and passenger are at great risk. It is a sure sign the pilot has become complacent and sloppy in his flight test program.
(3) The pilot will be breaking a contract (Operating Limitations) with the U.S. Government, which is known not to look kindly on such matters.
i. Pilots should ensure that the added ballast weight in the cockpit is secured. A seat belt over some sand bags will not stop the weight from shifting and getting loose in a cockpit. The last thing a test pilot needs is a 20-pound lead-shot bag free in the cockpit during a climb test, a landing, or a spin. Tie each weight down individually, and cover all the weights with a cargo net.
j. Ensure the ropes/nets and airframe attach points are strong enough to take the added load.
Make sure the passenger seat can take that much localized weight safely.
k. The maximum gross weight test results should be recorded in the flight manual. If there are any changes to the stall speed initially marked on the airspeed indicator, it should be changed to reflect the aircraft stall speed at maximum gross weight.
58 5/24/95AC 90-89A SECTION 2. SERVICE CEILING TESTS ‘‘Man is made for error; it enters his mind naturally and he discovers a few truths only with the greatest effort.’’ Frederick the Great
1. OBJECTIVE. To determine the highest altitude at which an aircraft can continue to climb at 100 feet per minute (Service Ceiling).
a. Pilots who wish to determine the actual service ceiling of their aircraft are offered the following suggestions:
(1) Ask the local Flight Standards District Office (FSDO) to amend the Operating Limitations to permit a climb to the aircraft’s service ceiling, if that altitude is above 18,000 feet.
(2) Contact the local Flight Service Station (FSS) or ATC facility, and reserve a time and airspace to make the test.
(3) Install a transponder (reference FAR § 91.215) or get a waiver.
(4) Install a portable oxygen bottle, if plans are to go above 12,000 feet. (Recommend the pilot becomes familiar with the symptoms and cures of hypoxia and hyperventilation.)
(5) Review the engine manufacturer’s mixture leaning procedures.
(6) Maintain communications with an air traffic facility at all times.
b. The climb to the aircraft service ceiling should be made in a series of step climbs during which engine performance, temperatures and pressures are recorded. At the slightest indication of engine performance or aircraft control problems, the pilot should terminate the test and return to the airport.
59 SECTION 3. NAVIGATION, FUEL CONSUMPTION, AND NIGHT FLYING ‘‘That’s one small step for man, one giant leap for mankind.’’ Neil Armstrong
1. OBJECTIVES. To ensure all the small but important aspects of flight have been tested and found reliable.
a. The Magnetic Compass. The magnetic compass should have been checked for accuracy prior to the first flight. However, the addition and removal of equipment, changing of wire bundle routing, and other airframe modifications may have affected the accuracy of the instrument. The following recommendations are offered:
(1) The magnetic compass can be checked for accuracy by using a compass rose located on an airport, or using a hand held ‘‘master compass.’’ The master compass is a reverse reading compass with a gun-sight mounted on the top of it. With the aircraft facing north and the pilot running the engine at 1,000 rpm, a second individual standing 30 feet away facing due south ‘‘shoots,’’ or aligns, the master compass with the aircraft’s centerline. Using hand signals, the pilot aligns the aircraft with the master compass. The pilot then runs the aircraft engine up to approximately 1,700 rpm to duplicate the aircraft’s magnetic field and reads the compass.
NOTE: Conventional gear aircraft builders will have to position the magnetic compass in a straight and level position for this test.
Raise the tail or mount the compass level with the horizon.
(2) If the aircraft compass is not in alignment with the master compass (start at north), correct the error by adjusting the north/south brass adjustment screw with a non-metallic screwdriver (can be made out of stainless steel welding rod, brass stock, or plastic) until the compass reads correctly. Go to the reciprocal heading (south) and remove half the error. On the east/west headings, use the other brass adjustment screw to make the corrections using the same procedures that was used to correct the north/ south errors.
(3) Check again for errors at each cardinal heading. Record the last readings and prepare a compass correction card. The maximum deviation (positive or negative) is 10 degrees on any one heading.
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(4) If the compass cannot be adjusted to meet this requirement, install another one. If the new compass is not available, try a different location in the cockpit, away from all ferrous metals and electrical bundles.
NOTE: A common error that affects the compass’s accuracy is the mounting of magnetic compass on/in the instrument panel with steel machine screws and nuts rather than brass.
(5) If the aircraft has an electrical system it is recommended that two complete compass checks be made, one with all electrical accessories on (e.g., radios/nav lights), and one with all electrical accessories off. If the deviation in level flight is more than 10 degrees on any heading with the accessories on, make up a separate compass correction card that shows the magnetic heading with the equipment on.
(6) Record the findings in the aircraft’s flight manual and create a compass correction card, mounting it near the magnetic compass in the cockpit. Make two cards; one with radios on and one with radios and non essential electrical accessories off.
b. Very High Frequency (VHF) Omni-directional Radio Range (VOR) Check. The best guide to check the accuracy of the VOR on board equipment is the VOR Receiver Check found in the Airman’s Information Manual (AIM), available from the Superintendent of Documents. The following is an abbreviated summary of the VOR procedure in the AIM.
(1) For a ground test of the VOR, a VOR Test Facility (VOT) must be used. To use the VOT service, tune in the VOT frequency on the VOR receiver. It is normally 108 Mhz. With the Course Deviation Indicator (CDI) centered, the omni-bearing selector should read 0 degrees with the to/from indicator showing ‘‘from,’’ or the omni-bearing selector should read 180 degrees with the to/from indicator showing ‘‘to.’’ The maximum bearing error should never be more than four degrees.
NOTE: The VOT facilities closest to the flight test location can be found in the Airport/Facility Directory. It is available by subscription from NOAA Distribution Branch N/CG33, National Ocean Service, Riverdale, MD 20737, or contact the nearest FAA FSS.
(2) For the airborne test, select a prominent ground point along the selected radial, preferably more than 20 miles from the VOR. Maneuver the aircraft directly over the point at a reasonably low altitude.
(i) Note the VOR bearing indicated by the receiver when over the ground point. The maximum permissible variation between the published radial and the indicated bearing is six degrees.
(ii) If the aircraft has dual VOR’s, the maximum permissible variation between the two receivers is 4 degrees.
c. Fuel Consumption: a good indication of how much the engine is working for each rpm produced. For a new or recently overhauled engine, the fuel consumption should improve each flight hour until the engine finishes its ‘‘break in’’ period, i.e., after approximately 100 hours of operation.
(1) To determine the aircraft fuel consumption, lay out a race track course with 8 to 10 mile legs. If the aircraft has one fuel tank or cannot switch tanks, do the following: Determine the approximate fuel burn to reach 1,000, 3,000, 5,000, 7,000, and 9,000 feet of altitude. With full tanks, climb to 3,000 feet and run the race track course for half an hour at 55 percent power.
(2) Land and measure the fuel used by dipping the tanks with a calibrated fuel stick, or by adding measured amounts of fuel to the tank until the tank is full. Subtract the approximate fuel burn to altitude, and multiply the remainder by two to get the fuel burn per hour.
(3) The tests are much easier and the results more accurate if the aircraft has two independent fuel tanks. Take-off on one tank and switch to the opposite tank at the test altitude. At the completion of the test, switch back to the first tank; land and measure the amount of fuel added in both tanks and multiply the quantity by two to get the amount of fuel used per hour.
(4) Run the same test at 65 percent and 75 percent power at the same altitude, using the same procedures. Move up to the next altitude and run the same tests.
d. Night Operations: should be conducted in accordance with the aircraft’s FAA Operating Limitations and limited to normal climbs and descents (e.g., 500 feet per minute), pitch angles of less than 5 degrees, straight and level flight, and coordinated turns of no more than 20 degrees of bank angle.
(1) The main concern for night testing should be the availability of a horizonal reference
(e.g., bright moon or artificial horizon).
(2) Prior to every night flight, ensure a reliable flashlight with fresh batteries and a set of FLIGHT TEST PLAN procedures are on board.
Some night testing requirements should have already been determined on the ground. For example:
(i) The electrical load review of all the lights, pumps, instrumentation, and avionics did not exceed 80 percent of the aircraft’s charging system capacity.
(ii) The cockpit instrumentation lighting is adequate and was tested for reliability of operation during daytime flights.
(iii) The pilot has at least 1⁄2 hour of night time taxiing the aircraft. This practice is needed to familiarize the pilot with a different operating environment. Do not exceed engine operating temperatures during taxiing.
(iv) The position and brightness of instrument panel lights, anti-collision strobe lights, and rotating beacons will not adversely affect the pilot’s night vision.
(3) A suggested night flight test plan is a series of takeoffs and landings and traffic pattern entries and exits. The tests should begin while there is still enough light to read a newspaper and transition to true night flying. The actual night flight will consist of an evaluation of the effectiveness of the taxi/landing light system, during taxi, take-off, and landing. The pilot should note any glare on the windshield or light flicker on the instrument panel.
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