DC-3 Description

DDA DC-3, reg. # PH-DDA, Courtesy Trev Morson, The DC-3 Hangar.

General description of the aircraft

Note: The descriptions that follow were excerpted from the Netherlands Aviation Safety Board acident report 96-71/A-16 of the crash of a Dutch Dakota Association (DDA) DC-3 on 25 September 1996. Aircraft registration was PH-DDA and all 32 aboard perished. Go to Aviation Safety Network for a summary of the report.

The DC-3 is a twin piston-engine, low wing monoplane of all metal, semi-monocoque construction. It is equipped with a tail wheel undercarriage, of which the main wheels are hydraulically retractable. It has hydraulically operated split trailing edge flaps.

It is a twin-engined commercial aircraft manufactured by the Douglas Aircraft Company, California, USA. DC stands for Douglas Commercial. The fuselage is an almost circular shaped structure built up of frame and made entirely of aluminium. The wings consist of three sections (left-wing, right-wing and the wing's midsection). The left and right wing sections are bolt-connected to the center wing section.

The ailerons, rudder and elevator are fabric covered. The trailing edge of the center section wing and the inner trailing edge of the right and left wing contains four hydraulically operated trailing edge flaps. The ailerons are located outward of the trailing edge right and left wing flaps and are just like the rudder and elevator, made from aluminium alloy frames covered by fabric.

There are two main fuel tanks, 210 gallons each (794 liters), located forward of the center section spar and two auxiliary tanks, 200 gallons each (760 liters), located aft of the wing spar. Each engine nacelle contains one 33-gallon oil tank (126 liters).

The aircraft is equipped with two pilot seats, a foldable observer seat and 21 to 32 passenger seats. It is certified for two-pilot operation.

Two DC generators, one on each engine, and two batteries supply electrical power for the aircraft electrical 24 volt system.

Two hydraulic pumps, one on each engine, supply hydraulic power. Hydraulic pressure operates the retractable main undercarriage, the split flaps, the wheel brakes, the engine cowl flaps and the windshield wipers.

The flight controls are manually operated via steel cables and are equipped with manually operated trim tabs.

Weight and Balance

• Maximum Take Off Weight: 25,200 lb (11,431 kg) – Piedmont
• Maximum Landing Weight: 25,200 lb (11,431 kg) – Piedmont
• CG limits at MTOW: 11.0 to 28.0 % Mean Aerodynamic Chord, MAC.

Engines

The aircraft is equipped with two Pratt & Whitney double row, 14 cylinder radial reciprocating engines, type R1830-92 Twin Wasp, rated at 1,200 BHP at 2,700 RPM each. Each engine is mounted on the aircraft forward of the main landing gear nacelle by an engine support frame and is isolated from the aircraft by a firewall. Each engine is enclosed by a nacelle with hydraulically operated movable cowling flaps.

The 14 cylinders are radially positioned in two banks of seven cylinders around the two crank cases, positioned in tandem; the front and the rear crank case. The crank cases support a fly-weight balanced double-cranked crankshaft. The crankshaft is driven by two master piston rods (cylinders no. 5 and no. 12). The two master piston rod bearings on the crankshaft are of the plain bearing type. Each master piston rod supports six piston rods. Each cylinder is fitted with two spark plugs, an inlet and an outlet valve, rocker arms and push rods. The push rods are operated by a crankshaft driven cam disc through tappets with roller type cam followers. The crankshaft drives the propeller via a reduction gear. This gear is enclosed by the front casing which also supports the propeller thrust bearing.

Each engine has an independent oil system providing oil for internal lubrication and cooling of the engine, oil for propeller governing and oil for feathering and unfeathering of the propeller.

Propellers

Each engine is equipped with a Hamilton Standard 23E50-473 Three Blade Hydromatic Quick Feathering Constant Speed Propeller with light alloy metal blades.

Propeller and engine speed (RPM) are maintained automatically throughout the constant speed range (1,200 to 2,700 engine RPM) by varying the blade angle of the propeller in order to meet changing conditions of airspeed, altitude, attitude and power setting. The engine-driven propeller governor, which is mounted on top of the engine front casing, controls the change in blade angle. This governor boosts up engine oil pressure and uses centrifugal weight forces balanced against a cockpit controlled spring force to regulate this pressurized engine oil to the propeller blade change mechanism in the propeller dome.

Minimum blade angle is 16° (fine pitch stop) and maximum 88° (feathered).

Feathering System

General

Each propeller is equipped with a quick feathering system. Feathering of the blades to the maximum angle position of 88° gives a powerful braking effect to stop a running engine, prevents windmilling of the propeller after the engine has been stopped and reduces the aerodynamic drag to a minimum. The feathering system also can be used for unfeathering of the propeller in case an engine has to be restarted after it has been shut down in flight.

Propeller Feathering

The feathering system is powered by an electrically driven gear type oil pump controlled from the cockpit by momentarily pressing the feathering button in the overhead panel. This action turns the blades from the actual blade angle through the coarse blade angle range to a blade angle of 88°, which is the feathered position in which the blades are streamlined in the flight direction. Feathering takes approximately four seconds.

Propeller Unfeathering

This forces the propeller blades towards the fine pitch blade angle. Normally in flight the propeller will start to windmill as soon as it is out of the feathered position. At about 800 RPM the feathering button must be released to stop the feathering pump and the engine can be started. The propeller governor will resume its normal automatic propeller speed control as soon as the engine is running and engine oil pressure is normal. Unfeathering takes approximately twelve seconds.

Cockpit layout

General

The cockpit layout is of the conventional type with an instrument panel in front of each seat and a center pedestal with the engine controls, trim controls and the press to talk buttons for the Receiver-Transmitter. The flight controls comprise a control wheel and fully adjustable rudder pedals at each pilot station. According to statements of DDA pilots, rudder pedal adjustment is sufficient to allow adequate control in n-1 situations.

DC-3 aircraft are not equipped with an audio and/or visual stall warning device.

Flight Instruments

The "Basic-T"

Extensive studies of visual scanning patterns of flight instrument panel layouts have resulted in the adoption of a standard layout, called the basic-T. This layout presents the best arrangement of instruments for fast and accurate scanning of the four basic instruments (on top and horizontally from left to right): air speed, attitude and altitude, and (vertically) below the attitude indicator, the heading. The center of the scanning cycle is the attitude indicator. Additional flight instruments such as turn and bank and vertical speed indicator are positioned to the left and right of the basic-T (see the figure below). The basic-T layout has been commonly applied in most aircraft during many decades.

This screen shot of the Flight Sim R4D panel perfectly illustrates the basic-T arrangement of the flight gauges.

Feathering Checks

To check the operation of the feathering system, a limited feathering check and a full feathering check are used. The objective of a limited feathering check is to check the operation of the feathering pump. The feathering cycle is stopped by pulling the feathering button.

During the engine run-up by the pilots prior to a flight, a limited feathering check is carried out. At 1,700 RPM the feathering button is pressed. When an RPM drop is observed, the feathering button is pulled and an RPM increase should be verified. By this procedure only the feathering pump is checked.

Use of Derated Take-off Power versus Full Rated Take-off Power

In an Engine Operation Information Letter issued by Pratt & Whitney Aircraft, January 15, 1951, several reasons are stated why derated take-off power should not be used: engine-wise there is very little to recommend in support of reduced power.

• Pressure loads in the combustion chambers oppose the RPM produced loads on the reciprocating system because pressure cushions the centrifugal and inertia forces. If the pressure load is reduced, the wear due to high RPM is increased. This factor is further accentuated by the increased time required to reach the RPM reduction point after take-off. Sustained high RPM is a major factor in keeping engines from staying young and it takes more "RPM minutes" and "piston ring miles" along the cylinder walls to complete the first take-off phase if the manifold pressure is reduced.
• It is also advantageous to reach an air speed that provides cooling airflow as soon as possible.
• Reduced manifold pressure means less induction airflow which in turns means a leaner mixture. As the impeller speed remains the same, the mixture temperature is still at its maximum and the slight help from lowered pressure is offset by the leaner mixture.

Furthermore Engine Operation Letter number 25, January 23, 1952, states:

• Of the several individual forces comprising the resultant force that determines the bearing load, the one produced by centrifugal action predominates. If the other forces were absent the load on the bearing would vary as the square of the RPM and would be applied constantly by contact along an unchanging line;
• When the crankshaft is turning, the master piston rod bearing is pressed against the crankpin by a force which is the resultant of the separate centrifugal, inertia and gas load forces. The component of this resultant force that is tangential to the path of crankpin travel produces the torque output of the engine;
• The centrifugal load is opposed and diminished by the gas load which varies with manifold pressure. Also, because of the varying connecting rod angle, the gas load sweeps back and forth over an appreciable arc with the result that the line of contact is constantly changing;
• High RPM with low manifold pressure approaches the condition of high centrifugal load uncushioned by gas load. Also the line of contact remains more nearly constant and local heating at this region becomes serious. Temporary over-speeds can be tolerated only because there is sufficient heat reserve in the surrounding material and oil to absorb, temporarily, the increased rate of heat generation. If the full stabilized local temperatures were reached, permanent damage would probably result;
• A recent rash of master piston rod bearing failures in one training activity is an excellent illustration of the workings of these opposing forces and how the engine suffers when one is absent.

Investigation of the cause of these failures disclosed that a power setting requiring normal rated RPM with closed throttle was being used while in the traffic pattern to simulate emergency conditions. Now, the bearing is designed to take this condition – for a short interval. The acceptance of the engine design and development assumes that this type of operation will be very infrequent (perhaps once an overhaul period) and then, for only a few seconds duration. When such a high RPM with such a low manifold pressure is imposed for relatively long periods and with training program frequency, the results are inevitably bearing failure.

DC-3 Asymmetric Performance

Flight Handling

From a literature study the following flight handling characteristics of the DC-3 were obtained:

Longitudinal Control

In the clean configuration, with METO power, the aircraft is statically unstable throughout the speed range with the CG at its rearward limit. This implies that the aircraft does not return to its trimmed condition after a disturbance, and therefore constant pilot activity is required to ensure stable flight. The unstable characteristics increase with decreasing airspeed. The stick force stability is essentially zero at lower speeds, which degrades speed control (lack of "speed feel").

Lateral Directional Control

Rudder forces at large rudder deflections are in general very high. This characteristic hampers the execution of coordinated turns. Rudder and aileron forces in steady side slips tend to lighten for angles of side slip larger than 10°. Rudder overbalance, resulting in aerodynamic rudder lock, can occur at higher angles of side slip.

Aileron forces during steady side slips and in steady rolling manoeuvres are qualified as moderate.

Single Engine Characteristics

According to the DC-3 AOM the minimum control speed in the air VMC is 76 kt DIAS, which corresponds with 82 kt CAS. VMC is the lowest airspeed at which the aircraft can be flown on one engine, on a constant heading and with a bank angle of 5° towards the live engine, with the propeller of the shut down engine feathered in clean configuration and with maximum except take-off power on the running engine. In general low weight is the critical condition for determining VMC. In the case of the DC-3 this speed is not limited by the maximum rudder deflection, but by the rudder force (max. 180 lb), which the average human is able to exert. Unfeathering of the stopped propeller will significantly increase the actual minimum control speed. Calculations by NLR indicate that this speed, depending on propeller blade angle, can increase up to approximately 10 kt.

Stall Characteristics

In general, power-off stalling characteristics of the DC-3 are qualified as benign. However, in power-on conditions stalls are accompanied with violent rolling (to the left) and a sharp drop of the nose, with considerable loss of altitude before control can be regained. During n-1 stalls these effects increase considerably. An n-1 stall at low altitude may therefore be expected to be unrecoverable.

Performance

In the performance section 4.4.2 of the DDA DC-3 AOM the following rates of climb are listed in relation to aircraft AUW, with one engine at METO power and the other engine shut down and the propeller feathered, wheels and flaps up, at 1,000 ft and an airspeed of 88 kt DIAS (92 kt CAS):

Note: deviations from the above mentioned speed seriously degrade climb performance.

It should be taken into account that these single-engine performance figures are based on the results of test flights, during the original certification using aircraft in factory-new condition and flown by test pilots.

Single Engine Performance PH-DDA

According to the DDA AOM, it is required to cruise during single engine operation with the so called Minimum Comfortable Airspeed (MCA), which is in fact 1.05 V(L/D)n". For the clean configuration this is 106 kt DIAS, which equals 111 kt CAS. Flying at lower speed than MCA has a negative influence on aircraft performance and flight characteristics. Based on the AOM performance data of the DC-3, the power required to sustain level flight at constant speed during single engine operation has been calculated as a function of airspeed. This has been done for the stopped propeller feathered and unfeathered at the fine pitch stop (16°). Drag data of the unfeathered propeller have been provided by Hamilton Standard. It is established that the calculated required power to fly with MCA matches well with the value given in the AOM, which is 920 BHP for the present configuration.

Available single engine METO power (1,050 BHP) is insufficient to sustain level flight, in case the stopped propeller is fully unfeathered.

Standard Operating Procedures DDA (Single Engine Operations)

Commercial aviation companies routinely use standard practices and procedures with regard to n-1 training on the actual aircraft or on the simulator. The emphasis is put on the most critical situation, the occurrence of an engine failure during takeoff. Procedures to cater for this occurrence are incorporated in the relevant AOM's and training syllabi, as well as for a n-1 approach and landing. Information for engine failure during other phases of flight is usually limited to performance figures for the one or multi engine-out conditions.

DDA closely follows these standard practices and procedures and has incorporated these in the DDA Flight Training Curriculum of the DC-3. Emphasis here is also given to the n-1 situation occurring during take-off.

All training is done on the actual aircraft as a DC-3 simulator does not exist. For safety reasons the engine is not shut down and the propeller not feathered. To simulate the n-1 condition the engine and the propeller are set for zero drag. As a consequence hands on training for-in flight engine shut down, propeller feathering and engine re-start is not practiced.

Relevant AOM Standard Operating Procedures (SOP) are:

"The Pilot Flying (PF) always occupies the left seat, the Pilot Not Flying (PNF) always the right seat. Under normal circumstances PNF handles the Receiver-Transmitter."

When an emergency occurs, such as an n-1, the AOM states that it is considered of utmost importance, that one pilot is clearly charged with the control of the aircraft. The main task of the PF is to fly the aircraft, he must not be distracted by conversation or actions with respect to the trouble shooting. The PNF performs the actions according to the Emergency Check List (ECL). Which pilot is handling the Receiver-Transmitter during the execution of the ECL by PNF, is not covered in a SOP. When performing the emergency checklist procedures, in principle the PF will take over ATC communications by calling: "My R/T". "Only the Pilot-in-Command is authorized to declare an emergency, and it is up to him to decide, if and when such an emergency is declared. If an emergency is declared, the only appropriate manner is to give a "Mayday" call and to select 7700 on the ATC transponder. ATC must be informed as soon as possible about the consequences of an emergency and/or abnormal situation. Do not hesitate to call "Mayday" to declare an emergency, when the safety of the aircraft and/or its occupants is, or is likely to become, endangered. The captain considers all operational consequences for the remainder of the flight, including abnormal system procedures, airport facilities, landing weather, maintenance and emergency procedures."

Overboost

Overboosting will occur when the engine speed is low in relation to the applied amount of power (Low propeller RPM/high manifold pressure). When engine overboost occurs the reciprocating loads are high and will not be sufficiently opposed and cushioned by the centrifugal loads. Overboost will cause the master piston rod bearings to wear down quickly and may cause ovalising of the piston pin holes.

Underboost

This phenomenon has much in common with an engine overspeed. The engine speed/applied power ratio is too high which causes the centrifugal forces to be predominant. As it takes place in the lower RPM range valve/piston striking will not take place, but damage to the master piston rod bearings is usually more severe than with an overspeed. Underboost may be caused by unintentional power decrease due to carburetor icing, or by a too low power setting with high propeller RPM selected, such as during an emergency descent or when reducing power too much during the approach. Underboost will usually result in wear down of the bearing material of the master piston rod bearings over a period of time, enabling the maintenance crew to detect bearing damage when inspecting the oil filters.

Derating

According to Pratt & Whitney Aircraft the use of derated take-off power is not supported. The use of derated take-off power increases the wear of the master piston rod bearings. Derated take-off power was introduced by DDA pilots, who were used to jet engine techniques. By using derated take-off thrust the turbine inlet temperature is lowered. This diminishes creep and increases engine durability. However this technique is not valid for piston engines and may be a cause of master piston rod bearing failure.

Performance

In the AOM of the PH-DDA it is stated that for best performance and flight characteristics during single engine operation, it is required to cruise at the Minimum Comfortable Airspeed (MCA), which equals 111 kt CAS. The calculated required power to maintain this airspeed is approximately 920 BHP (matching with the value given in the AOM).

Power ratings of the engines are:

    Take-off 1,200 BHP at 2,700 RPM
    METO 1,050 BHP at 2,550 RPM