6-7 experiences no acceleration and remains in a steady trimmed condition until the force or moment balance is disturbed by an atmospheric irregularity or by pilot input.
Figure 6-3. Static Stability
6.6 Definitions.
6.6.1 P ositive Static Stability. When an aircraft tends to return to the state of initial equilibrium position following a disturbance.
6.6.2 Neutral Static Stability. When an aircraft remains in equilibrium in a “new” position, following a disturbance from an initial equilibrium position.
6.6.3 Negative Static Stability. When an aircraft tends to move further in the same direction as the disturbance that moved it from the initial equilibrium position.
Note: See Figure 6-3 for a g raphic of these three stability states.
6-8 Figure 6-4. Time
6.6.4 Dynamic Stability. The time history of the movement of the aircraft in response to its static stability tendencies following an initial disturbance from equilibrium. See Figure 6-4.
6.6.5 Test for Static Longitudinal Stability.
6.6.5.1 This test should be done first. All tests should be conducted with the aircraft in the forward of center CG. Climb to at least 6,000 feet AGL and trim the aircraft for zero stick force in straight and level flight at low cruising speed.
Note: Do not retrim the aircraft once the test has begun.
6.6.5.2 Apply a light “pull” force and stabilize at an airspeed about 10 percent less than the trimmed cruise speed. At this reduced airspeed, it should require a “pull” force to maintain the slower speed.
1. If it requires a “pull” force, pull a little further back on the stick and stabilize the airspeed at about 20 percent below the initial cruise trim speed.
2. If it requires a still greater “pull” force to maintain this lower airspeed, the aircraft has positive static longitudinal stability.
3. If at either test points, no “pull” force is required to maintain the reduced airspeeds, the aircraft has neutral static longitudinal stability.
4. If either of these test points requires a “push” force to maintain the reduced airspeed then the aircraft has negative static longitudinal stability.
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6.6.5.3 Repeat another series of static longitudinal stability tests using a “push” force on the control stick. At an airspeed 10 percent above the trim cruise speed, the control stick should require a “push” force to maintain the airspeed. If a “pull” force is required, the aircraft has negative static longitudinal stability.
Warning: If the aircraft exhibits negative static longitudinal stability, seek professional advice on correcting the problem before further flight.
6.6.5.4 After confirming the aircraft has positive static longitudinal stability, the pilot can check for short-period positive dynamic longitudinal stability. First, trim the aircraft to fly straight and level at normal trim cruise speed. With a smooth, but fairly rapid motion, push the nose down a few degrees.
6.6.5.4.1 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 keep your hands close to it. The aircraft with positive dynamic longitudinal stability will oscillate briefly about the trim attitude before stopping at the trim attitude position.
6.6.5.4.2 To test the aircraft for long-period positive dynamic longitudinal stability, 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.
You don’t have 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.
6.6.5.4.3 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.
6.6.5.4.4 An aircraft with long-period neutral dynamic longitudinal stability will continue to oscillate through a series of increasing/decreasing airspeeds and never return to the original trim airspeed.
6.6.6 Lateral-Directional Stability Control Tests. Lateral (dihedral effect) and directional stability tests show if the aircraft tends 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 (VA) or any other airspeed limitation.
6.6.6.1 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
6-10 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.
6.6.6.2 At no time should there be a tendency toward a force reversal, which could lead to an overbalance condition or a rudder lock.
6.6.6.3 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.
6.6.6.4 To check static directional stability, trim the aircraft at a low cruise setting above 5,000 feet AGL. 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.
6.6.7 S piral 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: F riction in the aileron control system can completely mask th e inherent spiral characteristics of the airframe.
6-11 SECTION 4. A WORD OR TWO ABOUT FLUTTER
6.7 Objective. Understand the causes and cures of the condition known as flutter.
6.8 Description. Flutter in an aircraft structure is an interaction between aerodynamic inputs, the elastic properties of the structure, the mass or weight distribution of the various elements, and airspeed.
6.8.1 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 flag’s 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.
6.8.2 Think of a primary surface—a wing—with a control, such as an aileron, hinged to it.
Imagine that the airplane hits a thermal. The initial response of the wing is to bend upwards relative to the fuselage.
6.8.3 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.
6.8.4 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.
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6.8.5 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.
6.8.6 Flutter can happen at any speed, including takeoff speed. At low airspeeds, however, structural and aerodynamic damping quickly suppresses 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.
6.8.7 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 (¼-knot or less) to remove what little damping remains and the motion will become divergent rapidly.
6.8.8 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 differences are the masses involved are much smaller, the frequencies much higher, and there is less feedback 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.
6.8.9 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 check 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 they believe they 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.
6-13 SECTION 5. SPINS
6.9 Objective. To determine if spin testing is required.
6.10 Caution.
6.10.1 If the manufacturer/designer of the aircraft has not demonstrated satisfactory spin characteristics and safe recovery, avoid all types of high AOA flight testing and placard the aircraft: “spins prohibited.”
6.10.2 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.
6.10.3 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.
6.10.4 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.
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6.10.5 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 additional experience, aerobatic training with an emphasis on spins from a qualified instructor is highly recommended.
6.11 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 (CG) OF THE AIRCRAFT IS AT THE FORWARD CG LIMIT AND ANY BALLAST USED SHOULD BE SECURELY ATTACHED TO THE AIRCRAFT.
6.11.1 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.
6.11.2 The basic technique used to get a clean spin entry is to continue to reduce airspeed at about a 1-mph/knot-per-second rate in level flight, carb 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.
6. If the aircraft is not built for aerobatics, no further spin testing is required. It is recommended the instrument panel be placarded “SPINS PROHIBITED.”
7. If further spin testing is required, it is strongly recommended the services of a professional flight test pilot be used.
6-1 SECTION 6. ACCELERATED STALLS
6.12 Objective. To further explore the stall characteristics of the aircraft.
6.12.1 An accelerated stall is not a stall reached after a rapid deceleration. It is an in-flight stall at more than 1 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.
6.12.2 The two standard methods for performing 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.
6.12.3 Performing recommended constant bank method. First, review the techniques described for wing level stall tests on the flight test card and apply them here as well. If you are not spin-qualified and comfortable, hire a professional test pilot for these tests. Start with a light aircraft weight and forward CG. The amount of fuel in the tanks while performing the stall test should be enough fuel for 1 hour minimum. Follow the procedure below:
1. Normal takeoff and climb to 8,000 feet AGL.
2. Verify engine readings in the green.
3. Trim airplane to 1.5 times estimated stall speed.
4. Set the flaps (and landing gear) appropriate to test.
5. Power to idle (carb heat if applicable).
6. Establish—and maintain—a coordinated 30-degree bank turn.
7. Decelerate at 1 knot/mph per second.
8. Note positive stick/yoke force: it should increase.
9. Note control reversal: abort if evident.
10. Note speed at pre-stall buffet.
11. Recover, climb to starting altitude, and continue series until complete.
12. At stall, note indicated airspeed, pitch changes, roll direction, and amount.
13. Recover, climb to starting altitude, and continue series until complete.
14. You should have a detailed description of the expected behavior in an accelerated stall available (either from the kit manufacturer or through design analysis) prior to beginning the test. The objective is to verify (or not) that the behavior of your aircraft matches the expected behavior.
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6.12.4 Make separate test flights to evaluate the airplane’s performance at different weights and CG locations. Start with lighter gross weight and forward CG locations. Each test series includes six runs, left and right coordinated turns with the indicated flap settings. If your airplane has retractable gear, run the series with the gear up and down.
6.12.4.1 For each stall test fly the aircraft at a safe altitude (8,000 feet AGL or as appropriate for aircraft type, in smooth air), trim the airplane to 1.5 times the predicted stall speed, and set the flaps as required for the test (this should be slower than maximum flap-extended speed (VFE)).
6.12.4.2 Apply carb heat (if required), pull the power to idle, and establish a coordinated turn in a 30-degree bank; ensure the slip/skid ball is centered.
Reduce speed to 10 knots/mph above the straight and level stall speed for that flap setting. Then decelerate at 1 knot/mph per second.
1. As the airplane slows, make sure it requires an increasingly greater stick/yoke pull force. If the force lightens or changes to a push force— abort the test. This may indicate an aft CG location or insufficient elevator authority. Either of these may cause the nose to pitch up at stall.
2. As the airplane slows, normal control inputs should maintain the 30-degree bank attitude and nose position. When you feel the pre-stall buffet, note the speed, roll level, and recover to straight and level flight.
Note: Performing accelerated stalls in the airplane is a personal decision that should be based on the wing drop it exhibited in your wings-level stall tests. Accelerated stalls usually exacerbate this roll, especially when you’re turning to the direction the wing drops. Even a properly rigged airplane may roll up to 60 degrees into the turn, or up to 30 degrees in the opposite direction. If the possibility of a roll to 60 degrees is deemed unsafe—do not perform this test.
7-1 CHAPTER 7. PUTTING IT ALL TOGETHER: 36 HOURS TO —————?
SECTION 1. MAXIMUM GROSS WEIGHT TESTS
7.1 Objective. To develop aircraft performance data across the weight and CG ranges.
7.1.1 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.
7.1.2 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 AOH.
7.1.3 Each phase of the testing should be done slowly, incrementally, with the same careful attention to detail that should characterize all the flight testing.
7.1.4 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 W&B and CG location should be worked for each new increase in weight. Stop testing when the aircraft’s maximum gross weight is reached.
7.1.5 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 W&B 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.
7.1.6 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!
7.1.7 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.
7.1.8 The pilot should avoid the temptation to take a live ballast weight up for a ride for three reasons:
7.1.8.1 The aircraft has not been proven safe for the higher gross weights.
7.1.8.2 The pilot and passenger are at great risk. It is a sure sign the pilot has become complacent and sloppy in their flight test program.
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7.1.8.3 In most cases, unless otherwise approved by the FAA, the pilot will be in violation of the Operating Limitations for Phase I flight testing. If this is discovered and determined to be an intentional and/or reckless deviation from the regulations, legal action would be appropriate (refer to FAA Order 2150.3, FAA Compliance and Enforcement Program).
7.1.9 P ilots should ensure that the added ballast weight in the cockpit is secured. A seatbelt 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.
7.1.10 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.
7.1.11 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.
SECTION 2. SERVICE CEILING TESTS
7.2 Objective. To determine the highest altitude at which an aircraft can continue to climb at 100 feet per minute (service ceiling).
7.2.1 Pilots who wish to determine the actual service ceiling of their aircraft are offered the following suggestions: