SECTION 1
Engine-Out Aerodynamics
Aerodynamics Effects of an Engine Failure
When an engine failure occurs in a multi-engine aircraft, asymmetric thrust and drag cause the following
effects on the aircraft’s axes of rotation:
Pitch Down (Lateral Axis)
Loss of accelerated slipstream over the horizontal stabilizer causes it to produce less negative lift,
causing the aircraft to pitch down. To compensate for the pitch down effect, additional back pressure is
required.
Roll Toward the Failed Engine (Longitudinal Axis)
The wing produces less lift on the side of the failed engine due to the loss of accelerated slipstream.
Reduced lift causes a roll toward the failed engine and requires additional aileron deflection into the
operating engine.
Yaw Toward the Dead Engine (Vertical Axis)
Loss of thrust and increased drag from the windmilling propeller cause the aircraft to yaw toward the
failed engine. This requires additional rudder pressure on the side of the operating engine. “Dead foot,
dead engine.”
Engine Inoperative Climb Performance
Climb performance depends on the excess power needed to overcome drag. When a multi-engine
airplane loses an engine, the airplane loses 50% of its available power. This power loss results in a loss
of approximately 80% of the aircraft’s excess power and climb performance. Drag is a major factor
relative to the amount of excess power available. An increase in drag (such as the loss of one engine)
must be offset by additional power. This additional power is now taken from the excess power, making
it unavailable to aid the aircraft in climb. When an engine is lost, maximize thrust (full power) and
minimize drag (flaps and gear up, prop feathered, etc.) in order to achieve optimum single-engine climb
performance.
Approximate Drag Factors
-
Full flaps :
Windmilling prop :
Gear extended :
Control deflections :
-300 to –400 fpm vertical speed
-400 to –500 fpm vertical speed
-150 fpm vertical speed
variable
Airspeeds for Max Single-Engine Performance
VXSE
The airspeed for the steepest angle of climb on single-engine.
VYSE
The airspeed for the best rate of climb on single engine. (or for the slowest loss of altitude on drift
down.) Blue line is the marking on the airspeed indicator corresponding to Vyse at max weight.
Sideslip Versus Zero Sideslip
During flight with one engine inoperative, proper pilot technique is required to maximize aircraft
performance. An important technique is to establish a Zero sideslip condition.
Sideslip Condition (Undesirable)
When an engine failure occurs, thrust from the operating engine yaws the aircraft. To maintain aircraft
heading with the wings level, rudder must be applied toward the operating engine. This rudder force
results in the sideslip condition by moving the nose of the aircraft in a direction resulting in the
misalignment of the fuselage and the relative wind. This condition usually allows the pilot to maintain
aircraft heading; however, it produces a high drag condition that significantly reduces aircraft
performance.
Zero Sideslip Condition (Best Performance)
The solution to maintain aircraft heading and reducing drag to improve performance is the Zero Sideslip
Condition. When the aircraft is banked into the operating engine (usually 2 – 5 deg) the bank angle
creates a horizontal component of lift. The horizontal lift component aids in counteracting the turning
moment of the operating engine, minimizing the rudder deflection required to align the longitudinal axis
of the aircraft to the relative wind. In addition to banking into the operating engine, the appropriate
amount of rudder required is indicated by the inclinometer ball being split towards the operating engine
side. The Zero sideslip condition aligns the fuselage with the relative wind to minimize drag and must be
flown for optimum aircraft performance.
Single-Engine Service Ceiling
Single-engine service ceiling is the maximum density altitude at which the single-engine best rate of
climb airspeed (Vyse) will produce a 50 FPM rate of climb with the critical engine inoperative.
Single-Engine Absolute Ceiling
Single-engine absolute ceiling is the maximum density altitude that an aircraft can attain or maintain
with the critical engine inoperative. Vyse and Vxse are equal at this altitude. The aircraft drifts down to
this altitude when an engine fails.
Climb Performance Depends on Four Factors
-Airspeed: Too little or too much will decrease climb performance.
-Drag: Gear, Flaps, Cowl Flaps, Flight Control Deflection, Propeller, and Sideslip.
-Power: Amount available in excess of that needed for level flight.
(engines may require leaning due to altitude for max engine performance)
-Weight: Passengers, baggage, and fuel load greatly affect climb performance.
Critical Engine
The critical engine is the engine that, when it fails, most adversely affects the performance and handling
qualities of the airplane.
On most multi engine aircraft, both propellers rotate clockwise as viewed from the cockpit. By
understanding the following factors when flying an aircraft that has both propellers rotating clockwise, it
will be apparent that a left-engine failure makes the aircraft more difficult to fly than a right-engine
failure. The clockwise rotation of the props contributes to the following factors that cause the left engine
to be critical:
P-Factor
Accelerated Slipstream
Spiraling Slipstream
Torque
P-Factor (Yaw)
Both propellers turn clockwise as viewed from the cockpit. At low airspeed and high angles of attack,
the descending blade produces more thrust than the ascending blade due to its increased angle of attack.
Though both propellers produce the same overall thrust, the descending blade on the right engine has a
longer arm from the CG than the descending blade on the left engine. The left engine produces the thrust
closest to the center line. The yaw produced by the loss of the left engine will be greater than the yaw
produced by the loss of the right engine, making the left engine critical
Accelerated Slipstream (Roll and Pitch)
P-Factor causes more thrust to be produced on the right side of the propeller. This yields a center of lift
that is closer to the aircraft’s longitudinal axis on the left engine and further from the longitudinal axis
on the right engine and also results in less negative lift on the tail. Because of this, the roll produced by
the loss of the left engine will be greater than the roll produced by the loss of the right engine, making
the left engine critical.
Spiraling Slipstream (Yaw)
A spiraling slipstream from the left engine hits the vertical stabilizer from the left, helping to counteract
the yaw produced by the loss of the right engine. However, with a left engine failure, slipstream from
the right engine does not counteract the yaw toward the dead engine because it spirals away from the
tail, making the left engine critical.
Torque (Roll)
For every action, there is an equal an opposite reaction. Since the propellers rotate clockwise, the aircraft
will tend to roll counterclockwise. When the right engine is lost, the aircraft will roll to the right. The
right rolling tendency, however, is reduced by the toque created by the left engine. When the left engine
is lost, the aircraft will roll to the left, and the torque produced by the right engine will add to the left
rolling tendency requiring more aileron input, which increases drag, making the left engine critical.
VMC
VMC is the minimum airspeed at which directional control can be maintained with the critical engine
inoperative. VMC speed is marked on the airspeed indicator by a red radial line. Aircraft manufacturers
determine Vmc speed based on conditions set by the FAA under FAR 23.149.
1.
2.
3.
4.
5.
6.
Most Unfavorable Weight and Center of Gravity
Standard Day Conditions at Sea Level (Max Engine Power)
Maximum Power on the Operating Engine (Max Yaw)
Critical Engine Prop Windmilling (Max Drag)
Flaps Takeoff Position, Landing Gear Up (Least Stability)
Up to 5˚ of Bank into the Operating Engine
Any change to the previous conditions changes Vmc speed, possibly significantly. The following
summarizes how Vmc may be affected by the above conditions:
1. Most Unfavorable Weight and Center of Gravity
The certification test allows up to 5˚ bank into the operating engine. In a given bank, the heavier the
aircraft, the greater the horizontal component of lift that adds to the rudder force. As weight increases,
the horizontal component of lift increases, which added to the rudder, decreases Vmc. As the center of
gravity moves forward, the moment arm between the rudder and the CG is lengthened, increasing the
leverage of the rudder. This increased leverage increases the rudder’s effectiveness and results in a lower
Vmc speed.
2. Standard Day Sea level
Standard day conditions yield high air density that allows the engine to develop maximum power. An
increase in altitude or temperature (a decrease in air density) will result in reduced engine performance
and prop efficiency. This decreases the adverse yaw effect. Vmc speed decreases as altitude increases.
3. Maximum Power On The Operating Engine
When the operating engine develops maximum power, adverse yaw is increased toward the inoperative
engine. The pilot must overcome this yaw to maintain directional control. Any condition that increases
power on the operating engine will increase Vmc speed. Any condition that decreases power on the
operating engine (such as power reduction by the pilot, an increase in altitude, temperature, low density,
or aging engine) will decrease Vmc.
4. Critical Engine Prop Windmilling
When the propeller is in a low pitch position (unfeathered), it presents a large area of resistance to the
relative wind. This resistance causes the engine to windmill. The windmilling creates a large amount of
drag and results in a yawing moment into the dead engine. When the propeller is feathered, the blades
are in a high pitch position, which aligns them with the relative wind, minimizing drag. A feathered prop
will decrease drag and lower Vmc.
5. Flaps Takeoff Position, Landing Gear Up
When the gear is extended, the gear and gear doors have keel effect, reducing the yawing tendency and
decreasing the Vmc speed. Extending the landing gear also lowers the CG and may move the CG.
Extended flaps have a stabilizing effect that may reduce Vmc speed.
6. Up to 5˚ Bank into the Operating Engine
When the wings are level, only the rudder is used to stop the yaw produced by the operating engine
(sideslip condition). Banking into the operating engine creates a horizontal component of lift which aids
the rudder force. With this horizontal component of lift and full rudder deflection, Vmc increases with
decreasing bank by a factor of approximately 3 knots per degree of bank angle.
___________________________________________________________
SECTION 2
Aircraft Systems
Engines
The Turbocharged Twin Comanche is powered by two Lycoming IO-320-C series engines rated at 160
bhp each at 2700 rpm. The IO-320-C series are four cylinder, direct drive, wet sump, horizontally
opposed, air cooled, and have 319.8 cubic inches of displacement. These engines are designed to operate
on 100/130 (100LL) minimum octane aviation-grade fuel. Major accessories furnished with the engines
are geared starters, 50 ampere, 12 volt generators, dual vacuum pumps, direct-drive fuel pumps, dry
automotive-type induction air filters, and dual magnetos. An external oil cooler is mounted on the left
rear of each engine baffle.
Cowl flaps are manually operated by two push-pull controls located to the right of the power quadrant.
The cowl flaps should be open during ground operations and in climbs.
The fuel injection system is a Bendix self-purging servo regulator metering system. The system is
equipped with a manual mixture control and idle cut-off mechanism. A fuel flow indicator is installed in
the instrument panel to give an accurate indication of fuel flow. It is important to note that an indication
of increasing or abnormally high fuel flow is a possible symptom of restricted injector lines or nozzles.
Induction air is normally directed through a filter, but the induction system includes a spring loaded door
which opens automatically if the filter becomes blocked to allow air to the engine. This alternate air door
can also be operated manually by a push-pull (ALT AIR) control on the instrument panel. This control
should be operated if induction system icing is suspected.
Turbocharger
The turbo Twin Comanche is equipped with two manually operated Rajay turbochargers. The
turbocharger unit consist of a precisely balance radial inflow turbine and centrifugal compressor
connected by a short shaft. The turbocharger is driven by exhaust gas energy which is normally wasted
overboard, and controlled by vernier type push-pull waste gate controls which are located under the
throttle quadrant.
The turbochargers are normally used only after full throttle is applied, but turbo-normalizing may be
elected when operating the aircraft from a high altitude airport. When increasing power, apply throttles
first, then turbochargers. When decreasing power, reduce turbochargers first, then throttles.
The turbocharger bearings are lubricated with engine oil at 30-50 psi. if the oil pressure falls below 2730 psi, a red warning light will be activated. Should this occur, deactivate the turbocharger immediately.
If the turbo compressor fails for any reason, a check valve door will open automatically and the engine
will revert to normally aspirated operation.
Propellers
The Twin Comanche is equipped with Hartzell two-bladed, controllable pitch, constant speed, full
feathering metal propellers.
Controllable Pitch
Controllable pitch is the ability to control engine rpm by varying the pitch of the propeller blades. When
the blue propeller control handle is moved forward, oil pressure, regulated by a propeller governor,
drives a piston, which moves the blades to a low pitch-high rpm (unfeathered) position. When the blue
propeller control handle is moved aft, oil pressure is reduced by the propeller governor. This allows a
nitrogen-charged cylinder, spring, and centrifugal counterweights to drive the blades to a high pitch-low
rpm (feathered) position.
Constant Speed
After rpm setting is selected with the blue propeller control handles, the propeller governor will
automatically vary oil pressure inside the propeller hub to change the propeller blade pitch in order to
maintain a constant engine rpm. Because of this, changes in power setting (manifold pressure) and flight
attitude will not cause a change in rpm.
Full Feathering
When the propeller blades are in alignment with the relative wind, they are feathered. Feathered
propeller blades reduce the drag caused by the blade area exposed to the relative wind. Feathering the
propeller blades on the twin comanche is accomplished by moving the blue propeller control handle
fully aft past the low rpm detent, into the feather position. The propeller takes approximately three
seconds to feather. The right engine is slaved to the left, and the systems range of control is limited to 50
rpm. When feathering the propeller, the mixture should be placed to cutoff to stop engine combustion
and power production.
Propeller Overspeed
Propeller overspeed is usually caused by a malfunction in the propeller governor which allows the
propeller blades to rotate to full low pitch. If propeller overspeed should occur, retard the throttle. The
propeller control should be moved to full decrease rpm and then set if any control is available. Airspeed
should be reduced and throttle used to maintain a maximum of 2700 rpm.
Landing Gear
The PA-30 tricycle landing-gear system is fully retractable air-oil, oleo-strut type, and is electrically
operated by a selector switch located on the instrument panel. The three landing gear are mechanically
connected, and move as a unit.
The nose gear is steerable with the rudder pedals through a forty degree arc. The steering mechanism is
disconnected automatically during gear retraction to reduce rudder pedal loads in flight. The nose wheel
is equipped with a hydraulic shimmy damper.
Retraction of the landing gear is accomplished by an electric motor and transmission assembly located
under the floorboard, activating push-pull cables to each of the main gear, and push-pull tube to the nose
gear. Limit switches are installed in the system to cut off the gear motor when the gear is fully extended
or retracted.
To guard against inadvertent movement of the landing gear selector switch while on the ground, a
mechanical guard is positioned just below the switch. The switch handle must also be pulled aft before
being placed in the GEAR UP position. A warning horn will sound if the selector switch is placed in the
GEAR UP position while the weight of the airplane is resting on the landing gear.
To prevent inadvertent retraction of the landing gear while the airplane is on the ground, a safety squat
switch is installed on the left main gear to open the electric circuit to the landing gear motor until the
strut is fully extended.
If manifold pressure on both engines is reduced below approximately 12 inches, and the landing gear is
not down and locked, a warning horn will sound to alert the pilot to the possibility of a gear up landing.
The landing gear warning horn emits a continuos sound.
A green light on the instrument panel is the primary indication that the landing gear is down and locked.
When the gear is fully extended, the series circuit that lights this lamp is completed through a switch
located on each of the three gear. All three gear must be down and locked for the indicator to light. An
amber light above the landing gear selector switch indicates the gear is up. This lamp will flash if the
landing gear is up and manifold pressure of one engine is reduced below approximately 12 inches. A
third white light (installed on later models) will indicate that the landing gear is in transit. It is important
to note that the landing gear indication lights are automatically dimmed when the navigation lights are
turned on.
A removable emergency handle is used to manually extend the landing gear in the event of a
malfunction of the electrical system.
Brake System
The brakes are activated by toe pedals mounted above the pilot rudder pedals, or by a hand lever located
below the left center of the instrument panel. The hydraulic brake system is a self-adjusting, single-disk,
double piston assembly. Each rudder pedal has its own master cylinder, but both share a common
reservoir.
The parking brake is connected mechanically to the master cylinders and may be set by applying the toe
brakes or hand lever and pulling out the parking brake “T” handle. To prevent inadvertent application of
the parking brake in flight, a safety lock is incorporated to eliminate the possibility of pulling out the
“T” handle until pressure is applied by use of the toe brake or hand lever. To release the parking brake,
apply the brakes and push in on the parking brake “T” handle.
Fuel System
The fuel cells of the Twin Comanche consist of rayon-neoprene bladders which are contained in cavities
in the forward section of the wing. The inboard main cells hold a capacity of 30 gallons (27 usable) each
and the outboard auxiliary cells have a capacity of 15 gallons each.
Fuel cells are vented individually by NACA anti-icing vent tubes located beneath the wing. Fuel from
each cell passes through a selector-shutoff valve to a sediment bowl in the lowest part of the fuel system
where it is filtered, and any water or foreign particles are trapped. From there the fuel is drawn to the
fuel injection system by an engine driven pump. In the event of failure of the engine driven pump, an
electric auxiliary fuel pump is provided. In addition to the back up function, this pump is normally
operated when switching fuel tanks and during starting, takeoff, landing, and when operating above
15,000 feet MSL.
The fuel strainer units are located under the floorboard between the pilot and copilot seats just aft of the
fuel selector valves. Daily draining of the sediment bowl is accomplished by opening the hinged access
door and operating the quick drain valve for approximately five seconds with the fuel selector valve on
one cell. Change the fuel selector to the next cell and repeat the process until all cells are checked.
Allow enough fuel to flow to clear each of the lines as well as the sediment bowl strainer. Positive fuel
flow shutoff can be observed by means of the clear plastic tube which carries the fuel overboard.
Fuel quantity is indicated by two electric gauges located in the engine instrument cluster. The fuel
quantity gauges will indicate the amount of fuel in the cells that are selected by the selector shutoff
valves.
A crossfeed is provided for emergency single engine operation. To use fuel from the side opposite of the
operating engine, place the fuel selector for the inoperative in the “MAIN” or “AUXILIARY” position
and place the fuel selector for the operating engine in the “CROSSFEED” position.
Never place both fuel selectors in the crossfeed position at the same time, and the fuel system should be
taken off crossfeed before executing a single engine landing.
Electrical System
Electrical power for the Twin Comanche is supplied by a 12-volt, direct current, negative ground
system. The primary electrical power source is dual 12-volt 70 ampere alternators protected by an over
voltage relay. Secondary power is provided by a 12-volt, 35 ampere hour battery which supplies power
for starting, and is a reserve power source in the event of alternator failure.
The battery is mounted in a stainless steel box either immediately aft of the baggage compartment or in
the nose section. The voltage regulator is mounted on the aft bulkhead of the nose section. The ammeter,
located on the instrument panel near the engine gauge cluster, indicates battery discharge.
Vacuum System
The vacuum system provides the suction necessary to operate the attitude indicator and the directional
gyro. The engine-driven system consist of two vacuum pumps, each of which has its own vacuum relief
valve with filter, a system inlet air filter, and a suction gauge. If suction is lost from one of the vacuum
pumps, a check valve closes and adequate suction is supplied by the remaining pump. A mechanical
warning indicator located in the suction gauge will indicate which pump has failed. The vacuum pumps
are dry-type pumps. A shear drive in the pump assembly protects the engine from damage. Caution
should be exercised to insure that the propeller is never pulled through backward, as doing so will
damage the rotary vanes in the vacuum pump and potentially render the gyros inoperative.
The vacuum regulator is adjustable to a normal reading of 5.0 inch Hg plus .1 or minus .2 inch Hg.
Proper adjustment is important because higher settings will damage the gyros, and the instruments will
be unreliable with a low setting.
Pitot-Static System
The pitot-static system provides ram air pressure to the airspeed indicator and static pressure to the
airspeed indicator, vertical speed indicator, and altimeter. The system is composed of a heated pitot tube
mounted on the lower surface of the left wing, a pair of static ports located on either side of the fuselage
aft of the baggage compartment, and the associated piping necessary to connect the instruments to the
sources.
An alternate static source is available. Airspeed and altitude readings can be expected to be higher than
normal when operating from the alternate air source.
Heating And Ventilating System
There are four individual controls located on the lower right side of the instrument panel which regulate
the flow of heating, defrosting and ventilating air.
Heat for the cabin interior is provided by a gasoline heater installed in the nose section. Heated air for
the defroster system is provided by the same heater, but has an individual control. Caution should be
used when operating the defroster on the ground as prolonged application of heat may cause damage to
the Plexiglas windshield.
The cabin heater consumes approximately one gallon of gasoline an hour when operating, and its source
is the right engine fuel supply. If the cabin heater is use, this factor should be computed when figuring
fuel consumption.
Fresh air is supplied to the cabin by air inlets located in the fuselage nose. Two adjustable ventilators are
located near the floor forward of the front seats. In addition, there is a fresh air scoop located in the
dorsal fin which provides air to the rear seating positions.
The airplane is equipped with a janitol heater. Controls are provided to direct the airflow to both the
front and rear cabin. The heater uses gasoline supplied from the right engine fuel system. If the right fuel
selector is off, the heater is inoperative. A temperature limit switch will override heater operation if a
malfunction occurs resulting in excessive temperatures.
To operate the Janitol heater, the manual fuel valve must be ON and the three position switch on HEAT.
No priming is required. Heat is regulated by a thermostat and no excess hot air is exhausted overboard.
When the Janitol heater is turned off, and the manual fuel valve is closed, combustion stops and no
purge time is required.
Stall Warning
An approaching stall is indicated by a stall warning horn which is activated between five and ten knots
above stall speed. The stall warning horn emits an interrupted sound to distinguish it from the landing
gear warning horn. Mild to moderate airframe buffeting and gentle pitching may precede the stall. The
stall warning lamp is activated by a lift detector installed in the leading edge of the left wing. The stall
warning system is inoperative with the master switch off.