Fam 15 Discuss Items

=Retreating blade stall= Blade stall occurs when the angle of attack of a significant segment of the retreating blade exceeds the stall angle. When this condition occurs, increased blade pitch (or collective) will not result in increased lift and may result in reduced lift and increased rotor drag. One of the more important features of the TH-57 two-bladed, semirigid system is its warning to the pilot of impending blade stall. Prior to progressing fully into the stall region, the pilot will feel marked increase in airframe vibration and control vibrations. Consequently, corrective action can be taken before stall becomes severe.

The threshold of stall varies with the following:
 * 1. Airspeed
 * 2. Gross weight
 * 3. Density altitude
 * 4. G loading
 * 5. RPM.

INDICATIONS:
 * Progressively increasing two-per-revolution vibrations.
 * Loss of longitudinal control and severe feedback in the cyclic.
 * Violent vertical nose oscillations independent of cyclic position.

PROCEDURES: Recovery may be accomplished by one or a combination of the following: 1. Severity of maneuver - Decrease. 2. Collective pitch - Decrease. 3. Altitude - Descend. 4. Rotor RPM - Increase. CAUTION: Entry into severe blade stall can result in structural damage to the helicopter.

=Helicopter vibrations= Who knows where this stuff below came from... Maybe this was heisted from a maintenance manual. I know there is no chapter 63 in NATOPS and 57 drivers don't generally refer to the turbines as N1 N2.. Oh yeah... and have any of you done your 'Vibrex C Tracker/Balancer' qual since you've been here? Maybe just a quick look at NATOPS 11.13 like you did when you briefed this for CPT 4. You're a super fam now; life is good (at least for a few more flights).

VIBRATION ANALYSIS.
The Nodal-beam system provides a very low vibration level. When there is a definite deterioration from this level, immediate action should be taken to correct the condition. Sources for these vibrations are the rotating or moving components of the helicopter. The availability of electronic vibration and tracking devices such as the Chadwick and RADS-AT provide both immediate and positive determination of these sources. Extreme low frequency, low frequency, and most medium frequency vibrations are caused by the rotor or dynamic controls. Certain vibrations are inherent in the helicopter, and are considered normal. Two per revolution 2/rev), is the most prominent, with 4/rev or 6/rev the next most prominent. There is always a small amount of high frequency present. For purposes of this manual, vibrations are divided into four general frequencies and are described in the following paragraphs.

EXTREME LOW FREQUENCY VIBRATION.
Extreme low frequency vibration is limited to a two to three cycles per second, pylon rock, which is inherent with the rotor, mast, and transmission system. When this ìrockî is noticed during normal flight, it is an indication that something is wrong with the transmission mounts or transmission restraints and should be inspected to determine cause and corrective action (Chapter 63).

LOW FREQUENCY VIBRATION.
Low frequency vibrations of l/rev and 2/rev are caused by the main rotor. These are of two basic types; vertical or lateral. A l/rev vertical is caused by one blade developing more lift at a given point than the other blade develops at the same point. A lateral vibration is caused by an imbalance condition of the rotor due to: A difference of spanwise or chordwise weight between the blades. The CG alignment of the blades with respect to the spanwise axis which affects the chordwise balance. Hub imbalance: Initially the rotor is brought into a low speed ground track by rolling the grip on the high blade to fly it down in track with the low blade. This is normally accomplished at 70 to 75 percent N2. A high speed reference track is then made at 100 percent N2. Record tracking data for possible use during flight check. Generally, verticals felt predominately in low power descent at moderate airspeeds (60 to 70 knots) are caused by a basic difference in blade lift and can be corrected by rolling the grip slightly out of track. Verticals noted primarily in forward flight, that become worse as airspeed increases, are usually due to one blade developing more lift with increased airspeed than the other (a climbing blade). A slight raising of the trim tab on the low blade will generally correct this condition. Flight test after adjustment is required to determine acceptability. Analysis of flight test data will be required if further action is deemed necessary. The intermittent l/rev is essentially initiated by a wind gust effect, causing a momentary increase of lift in one blade. The momentary vibration is normal but if it is picked up by the rotating collective controls and fed back to the rotor causing several cycles of l/rev, then it is undesirable.

Sometimes during steep turns, one blade will pop out of track and cause a hard l/rev vertical. This condition is usually caused by too much differential tab in the blades and can be corrected by rolling one blade at the grip and removing some of the tab (as much as can be done without hurting the ride in normal flight). When a rotor or rotor component is out of balance, a l/rev vibration called a lateral will be present. This vibration is usually felt as a vertical due to the rolling motion it imparts to the helicopter, causing the crew seats to bounce up and down out of phase. When the pilot seat is going up, the copilot will be going down.

A severe lateral can be felt as a definite sideward motion as well as a vertical motion. Laterals existing due to imbalance in the rotor are of two types; spanwise and chordwise. Spanwise imbalance is caused simply by one blade and grip being heavier than the other (i.e., an imbalance along the rotor span).

A chordwise imbalance means there is more weight toward the trailing edge of one blade than the other. Both types of imbalance can be caused by the hub as well as the blades. Generally, a chordwise lateral imbalance condition is more pronounced at 95 percent N1 and a spanwise lateral is more pronounced at 100 percent N2. If spanwise imbalance is indicated, a wrap of one or two turns of 2-inch (50 mm) masking tape (or equivalent weight of another type) around one blade, a few inches in from the tip so that it will not be easily torn off by the wind. Hover the helicopter, either in or out of ground effect, wherever the lateral was most pronounced, and note the effect. An increase in vibration means that the tape was applied to the wrong blade. Once the correct blade is determined, further tape is added in amounts depending on the severity of the vibration. Utilize one-half wraps of tape until best balance is obtained. If the lateral remains excessive or if the tape is of no help on either blade, a chordwise imbalance exists and it is necessary to sweep a blade. One blade is arbitrarily selected and swept aft 1/4 point of blade latch nut. When sweeping a blade aft, always loosen leading edge and tighten trailing edge latch nuts the same amount; each nut one-quarter point (paragraph 18-26). To determine the effect of this sweep adjustment, hover the helicopter. When it is determined that the proper blade is being swept, continue sweep adjustments in amounts based on the severity of the vibration until the lateral imbalance is eliminated or further sweep fails to help. When it is necessary to sweep either blade more than two points, the rotor assembly should be removed, aligned, and statically balanced. If this action does not correct the problem, it will be necessary to return to taping and adjust tape and sweep until the optimum combination is obtained. If the lateral is still not eliminated, a small amount of grip rolling should be attempted as in the l/rev vertical procedure, being careful not to adversely affect forward flight. Should the lateral still be present, a small amount of tab may be tried. If the condition still exists, the hub and blades should again be removed and a careful check of the alignment and a static balance should be accomplished. Two per rev (2/rev) vibrations are inherent with a two bladed rotor system and a low level of vibration is always present. A marked increase over the normal 2/rev level can be caused by two basic factors: a loss of designed damping or absorption capability or an actual increase in the 2/rev vibration level of the rotor itself. The loss of damping can be caused by such factors as deteriorated transmission mounts, nodal beam attachments, or an airframe component loosening and vibrating in harmonics with the inherent 2/rev. An increase in the 2/rev level of the rotor itself can be caused by worn or loose components in the rotor hub or looseness in the rotating controls. Occasionally tab settings and sweep will affect the overall 2/rev level. If no mechanical cause of excessive 2/rev can be found, an attempt to decrease the level by rotor adjustments may be made. Tabbing both blades down (usually) or up (rarely) a few degrees sometimes helps. A recheck of boost off forces should be made. Sometimes both blades may be swept in the same direction in small amounts and thus decrease the 2/rev.

MEDIUM FREQUENCY VIBRATION.
Medium frequency vibrations (4/rev and 6/rev) are another inherent vibration associated with most rotors. An increase in the level of these vibrations is caused by a change in the capability of the fuselage to absorb vibration. Contributing factors may be a loose airframe component, such as skids, vibrating at that frequency. Changes in the fuselage vibration absorption can be caused by such things as fuel level, external stores, structural damage, structural repairs, internal loading or gross weight. Abnormal vibration levels can nearly always be attributed to one of these conditions. The vibration is felt as a rattling of the fuselage structure. The most common cause is loose skids caused by loose, worn or improper skid attachments. Loose skids can be detected by shaking the helicopter with cyclic and watching the skids vibrate. (Excessive or severe shaking is not recommended and may even make tight skids vibrate.) Many times skids will cause excessive vibration during turns and maneuvers if they are extremely loose. Other sources of medium frequency vibrations are caused by the elevator, access doors, cargo hook, electronic equipment, a safety belt hanging out the door, and engine/transmission cowling. Occasionally portions of the cabin roof, side panels or doors will ìoil canî rapidly in flight, giving the same sensation as a medium frequency vibration.

HIGH FREQUENCY VIBRATION.
High frequency vibrations are too fast to count and may be manifested as a buzz in the pedals. These vibrations may come from the engine or accessory gearbox components, improper drive shaft alignment, malfunctioning couplings, dry or excessively worn bearings, or an out-of-track or damaged rotor. If the vibration is associated with a malfunctioning tail rotor, the time from onset of vibration to complete loss of tail rotor thrust may be extremely limited. Recognition of suspected failure of the tail rotor drive system relies heavily on pilot judgement. Indications of suspected failure include, but are not limited to, sudden increase in amplitude of vibrations, unusual noises, and illumination of T/R CHIP caution light.

PROCEDURES: *1. ECS - Off If vibrations continue: *2. Land as soon as possible If vibrations cease: *3. Land as soon as practical WARNING: Be prepared to execute Complete Loss of Tail Rotor Thrust procedures. Increased power settings required to accomplish a normal approach may ultimately precipitate the complete failure of a malfunctioning tail rotor. Be prepared for uncommanded right yaw in the event of complete loss of tail rotor thrust during the approach. Consideration should be given to maintaining an autorotative profile or low-powered approach.

=Rotor Droop= Droop is a term used to denote a change in power turbine speed (Nf) and rotor speed that occurs with a demand for increased power with the governor at a constant speed setting. Droop may be further categorized as either transient or steady state. Transient droop is the momentary change in power turbine speed and rotor speed resulting from an increased power demand, and it is compensated for by the power turbine governor control. Steady-state droop is the decrease in power turbine speed and rotor speed that results from an increased power demand when the engine is already operating at maximum gas producer speed. This condition should be avoided during normal operation.

=Uncommanded Right Roll During flight below 1G= INDICATIONS: Uncommanded right roll Reduced cyclic effectiveness

PROCEDURES: *1. Cyclic — Immediately Apply AFT to Establish positive G Load on Rotor, then Center Laterally. WARNING: Lateral cyclic is decreasingly effective below 1g and increases main rotor flapping, which can result in mast bumping. When main rotor returns to a positive thrust condition: *2. Controls — As required to regain balanced flight. If mast bumping has occurred: *3. Land immediately.

NATOPS 11-2

=Tail Rotor malfunctions= Helicopter pilots should guard against the tendency to place all tail rotor malfunctions and their corrective actions into a single category. The successful handling of a tail rotor emergency lies in the pilot's ability to quickly analyze and determine the type of tail rotor malfunction that has occurred and then to execute the proper emergency procedures. It is as equally important that the pilot understand how the helicopter will react as it is to know the procedure. This requires forethought and discussion so that the pilot can more ably discern the situation and take the proper corrective action.

There are four basic types of tail rotor malfunctions which are covered in the following paragraphs. Of the basic types, the pilot can make some generalizations:
 * (1) 	in the case of right rotations, a low-powered approach or autorotation is the most likely course of action
 * (2) 	in the case of left rotations, a powered approach will usually be possible.

A controllability check at cruise flight should be performed, determining what torque is required for balanced flight. A high-torque setting (above hover power) will usually indicate a stuck left situation, and a low-torque setting (below hover power) will usually indicate a stuck right situation.

Complete Loss of Tail Rotor Thrust
Probable Causes:

1)Tail rotor driveshaft severed

2)Loss of tail rotor blades

Helicopter Reaction: In this situation, the nose of the helicopter will swing rapidly to the right in a hover with an accompanying sideslip in forward flight.

Procedures:

In a hover:

*1. Twist grip - Close *2. Cyclic - Eliminate Drift *3. Collective - Increase to cushion landing

During transition to forward flight or hover/air taxi:

*1. Twist grip - Close *2. Cyclic - Eliminate Sideward Drift *3. Collective - Increase to cushion landing

At altitude:

*1. Collective - Reduce to minimize yaw *2. Cyclic - Adjust for best airspeed to control yaw If yaw is not controllable: *3. Autorotate *4. Twist grip - Closed prior to flare If yaw is controllable: *5. Continue powered flight and set up to a suitable landing area at or above minimum rate of descent autorotational airspeed. *6. Autorotate *7. Twist grip - Closed prior to flare

WARNING: Airspeed indications during sideslip are unreliable. At airspeeds below approximately 50 knots, the side-slip may suddenly become uncontrollable, and the helicopter will begin an unrecoverable vertical axis "flat spin". If attempting to achieve higher airspeeds, care must be taken to avoid excessive cyclic inputs coupled with large power settings that could lead to mast bumping or rapid nose tucking.

NOTE: Depending on the nature of the failure and degree of damage, airspeeds between 50-72 KIAS may provide the best opportunity to maintain level flight. Due to yaw stiffness provided by the vertical fin at higher airspeeds, it may be possible to continue at faster airspeeds in cruise or descent depending on power requirements. A non-typical nose down attitude may be required to achieve a desired airspeed due to increased drag on the tail. Turns to the right may provide greater controllability of airspeed and potentially minimize altitude loss. Banking to the left will aid in counteracting torque.

WARNING: In the autorotation, maintain airspeed above minimum rate of descent airspeed until flare to avoid loss of yaw control. Once the engine is secured, in the absence of torque, the lift produced by the vertical fin may tend to yaw the nose to the left at faster speeds. As airspeed slows and Nr decays, the decelerating rotorhead and swashplate friction will create additional left yaw, increasing the chance for rollover. Depending on the landing profile, consideration should be given to leaving twist grip open until pulling collective at the bottom of the autorotation to allow control of yaw with twist grip.

Fixed Pitch Right Pedal Applied
Probable Causes:

1)Pedals locked in fixed position because of FOD

2)Control linkage failure during a right-pedal applied situation

Helicopter Reaction: The pilot will be unable to control right yaw with pedal input. If power is increased it will tend to aggravate the degree of yaw or sideslip.

Procedures:

In a hover:

If rate of rotation is not excessive and landing surface is smooth and firm: *1. Collective - Decrease to effect a power-on landing If rate of rotation is excessive or landing surface is unsuitable for a power-on landing: *2. Twist grip - Reduce as nose approaches windline *3. Cyclic - Eliminate drift *4. Collective - Increase to cushion landing

At altitude:

1. Maintain airspeed and engine rpm to streamline the aircraft. 2. Plan an approach to a smooth level surface into the wind or with a slight left crosswind if possible. 3. Establish a shallow approach, maintaining 60 KIAS until on final. Note: In such an approach profile, it is not unusual for the nose to be yawed slightly to the left. 4. At 50 to 75 feet AGL and when the landing area can be made, start a slow deceleration to arrive over the intended landing point with minimum forward speed required for directional control. 5. At approximately 2 to 3 feet skid height, increase collective to slow the rate of descent and coordinate twist grip to maintain nose alignment.

WARNING: If necessary, a waveoff should be made early in the approach, using cyclic to increase forward airspeed. If it becomes necessary to use large collective inputs to wave off near the deck, the nose will yaw right and possibly enter uncontrolled flight.

Note: If nose swings right after touchdown, follow the turn with cyclic to prevent the aircraft from rolling over.

Fixed Pitch Left Pedal Applied
Probable Causes:

1)Pedals locked in fixed position because of FOD

2)Control linkage failure during a left-pedal applied situation

Helicopter Reaction: The pilot will be unable to control left yaw with pedal input. If power is decreased it will tend to aggravate the degree of yaw or sideslip.

Procedures:

In a hover:

If rate of rotation is not excessive and landing surface is smooth and firm: *1. Collective - Decrease to effect a power-on landing If rate of rotation is excessive or landing surface is unsuitable for a power-on landing: *2. Twist grip - Slowly reduce while increasing collective to stop rotation. *3. Collective - Coordinate with twist grip to maintain heading and allow aircraft to settle.

At altitude:

1. Maintain airspeed and engine rpm to streamline the aircraft. 2. Plan an approach to a smooth level surface into the wind or with a slight left crosswind if possible. 3. Establish a normal approach and maintain 60 KIAS during the initial part of the approach. 4. On final approach, maintain engine rpm within limits and begin a slow deceleration in order to arrive at a point about 2 feet above the intended touchdown area as effective translational lift is lost. 5. Apply collective pitch to slow the rate of descent and align the helicopter with the intended landing path. If the aircraft is not aligned after pitch application, adjust the twist grip to further help with the alignment. Allow the aircraft to touch down at near zero groundspeed maintaining alignment with the twist grip.

Note: In a fixed-pitch left-pedal situation, it is possible for the pilot to slow the aircraft to a hover and effect such a recovery.

Loss of Tail Rotor Effectiveness
Four aircraft characteristics during low-speed flight have been identified through extensive flight and wind tunnel tests as contributing factors in unanticipated right yaw.

For this occurrence, certain relative wind velocities and azimuth (direction of relative wind) must be present. The aircraft characteristics and relative wind azimuth regions are:

1. Weathercock stability (120 to 240°)

2. Tail rotor vortex ring state (210 to 330°)

3. Main rotor vortex disk interference (285 to 315°)

4. Loss of translational lift (all azimuths).

5. Angle of Attack Reduction (060-120°). (out of Aero book pg. 9-29... now pg 6-12)

The aircraft can be operated safely in the above relative wind regions if proper attention is given to controlling the aircraft. However, if the pilot is inattentive for some reason and a right yaw is initiated in one of the above relative wind regions, the yaw rate may increase unless suitable corrective action is taken.

Weathercock Stability (120 to 240°)

Winds within this region will attempt to weathervane the nose of the aircraft into the relative wind. This characteristic comes from the fuselage and vertical fin. The helicopter will make an uncommanded turn either to the right or left depending upon the exact wind direction unless a resisting pedal input is made. If a yawrate has been established in either direction, it will be accelerated in the same direction when the relative wind enters the 120 to 240° shaded area of Figure 11-3 unless corrective pedal action is made. The importance of timely corrective action by the pilot to prevent high yaw rates from occurring cannot be overstressed.

Tail Rotor Vortex Ring State (210 to 330°)

Winds within this region, as shown in Figure 11-4, will result in the development of the vortex ring state of the tail rotor. The tail rotor vortex ring state causes tail rotor thrust variations that result in yaw rates. Since these tail rotor thrust variations do not have a specific period, the pilot must make corrective pedal inputs as the changes in yaw acceleration are recognized. The resulting high pedal workload in tail rotor vortex ring state is well known and helicopters are operated routinely in this region. This characteristic presents no significant problems unless corrective action is not timely. If a right yaw rate is allowed to build, the helicopter can rotate into the wind azimuth region where weathercock stability will then accelerate the right turn rate. Pilot workload during tail rotor vortex ring state will be high; therefore, the pilot must concentrate fully on flying the aircraft and not allow a right yaw rate to build.

Main Rotor Disk Vortex (285 to 315°)

Winds within this region, as shown in Figure 11-5, can cause the main rotor vortex to be directed onto the tail rotor. The effect of this main rotor disk vortex is to change the tail rotor angle of attack. Initially, as the tail rotor comes into the area of the main rotor disk vortex during a right turn, the angle of attack of the tail rotor is increased. This increase in angle of attack requires the pilot to add right pedal (reduce thrust) to maintain the same rate of turn. As the main rotor vortex passes the tail rotor, the tail rotor angle of attack is reduced. The reduction in angle of attack causes a reduction in thrust and a right yaw acceleration begins. This acceleration can be surprising, since the pilot was previously adding right pedal to maintain the right turn rate. Analysis of flight test data during this time verifies that the tail rotor does not stall. The helicopter will exhibit a tendency to make a sudden, uncommanded right yawwhich, if uncorrected, will develop into a high right turn rate. When operating in this region, the pilot must anticipate the need for sudden left pedal inputs.

Loss of Translational Lift

The loss of translational lift results in increased power demand and additional anti-torque requirements. If the loss of translational lift occurs when the aircraft is experiencing a right turn rate, the right turn rate will be accelerated as power is increased, unless corrective action is taken by the pilot. When operating at or near maximum power, this increased power demand could result in rotor rpm decay. This characteristic is most significant when operating at or near maximum power and is associated with unanticipated right yaw for two reasons. First, if the pilot’s attention is diverted as a result of the increasing right yaw rate, he may not recognize that he is losing relative wind and, hence, losing translational lift. Second, if the pilot does not maintain airspeed while making a right downwind turn, the aircraft can experience an increasing right yaw rate as the power demand increases and the aircraft develops a sink rate. Insufficient pilot attention to wind direction and velocity can lead to an unexpected loss of translational lift. The pilot must continually consider aircraft heading, groundtrack, and apparent groundspeed, all of which contribute to wind drift and airspeed sensations. Allowing the helicopter to drift over the ground with the wind results in a loss of relative windspeed and a corresponding decrease in the translational lift produced by the wind. Any reduction in translational lift will result in an increase in power demand and anti-torque requirements.

Angle of Attack Reduction (060 to 120°)

In a right crosswind, the relative wind shifts toward the rail rotor blade's chordline because of effectively increased induced velocity. The shifted relative wind impacts at a lower angle of attack, which develops lower lift and results in less thrust. The pilot will automatically compensate by adding more left pedal, but in some cases can reach pedal travel limits before adequate thrust can be generated.

PROCEDURES: *1. Pedals - Maintain full left pedal *2. Collective - Reduce (as altitude permits) *3. Cyclic - Forward to increase airspeed If spin cannot be stopped: *4. Autorotative landing - Execute