CONVAIR 580

CONVAIR

580

507

Used

Engines On Condition. VIP 36-Passenger First Class Configuration. All Leather/Gold Plated Interior. Updated Avionics, TCAS II, GPWS, FDR/CVR. Long Range Tanks. Executive Airline Operated. Beautiful Paint and Interior. All Inspections Current

30000.0

Weights: Maximum Gross Take-Off: 58,156 lb 26,380 kg Maximum Landing: 52,000 lb 24,040 kg Maximum Zero Fuel: 50,000 lb 22,680 kg Approximate Operating (PAX) 37,000 lb Approximate Operating (Cargo)32,000 lb Maximum Fuel Quantity: 13,900 lb (Jet A-1) 16,770 lb (JP-4)

Engine Model: Allison 501-D13 Engine No. 1: Engine No. 2: S/N TSO S/N TSO 501588 501720 Turbine 702142 380 702427 2,950 Reduction Gearbox 601409 2,240 601661 3,830 Tourquemeter A2518 1,740 A5514 1,140 APU 112P-9560 2,870 Propellor WY12388 915 HC2258 1,445 Regulator R134 1,980 R120 1,445 N.B.( ALL TSO FIGURES ARE QUOTED AS AT 24 JULY, 1998 ) Power plant The model 501 engine is a gas turbine, power unit connected to a torquemeter assembly and struts to a reduction gear having a single propeller shaft. The power section consists of a single-entry fourteen stage, axial flow compressor, a set of six combustion liners of the cylindrical through-flow type, and four-stage turbine. An accessory drive housing is mounted on the bottom of the forward end of the compressor. The reduction gear assembly contains two stages of reduction driving a propeller shaft. The first stage of reduction is accomplished by a set of spur gears; the second stage is accomplished by planetary gearing. The two stages provide a 13.54 to 1 reduction in power section-to-propeller shaft speed. Necessary gears and their drive pads are provided on the reduction gear for engine-driven aircraft accessories. The reduction gear assembly also incorporates an automatic propeller brake, a negative torque signal system, a thrust sensitive signal system and safety coupling. The torquemeter assembly transmits power from the power unit to the reduction gear. The twist of torquemeter drive shaft under load is measured electronically and registered as shaft horsepower on a cockpit indicator. Oil system - General Each engine (power unit and reduction gear box) is provided with a separate and independent oil system of an oil tank, oil-to-fuel heat exchanger and oil-to-air cooler. An accessory section oil cooler is provided for the right engine only. A stainless steel oil tank, located above the engine in the nacelle, has a total volume of approximately 11.5 U.S. gallons, which includes expansion and foaming space. The operating level of the oil tank is 8.5 U.S. gallons. An emergency oil shut-off valve in the bottom of the tank is actuated electrically by the emergency handle of the respective engine. Within the oil system proper, two oil supply routes are used in each engine. One supply is used for cooling and lubrication of the reduction gear box, and the other supplies the power section. Oil from the power section and reduction gear box is routed through the oil-to-fuel heat exchanger to provide fuel filter anti-icing before entering the oil cooler. Electrically-operated and automatically-controlled air exit flaps, with manual override, plus a ground air-induction system, are provided for regulation of the cooling air flow. Engine scavenge oil is utilised to cool the right engine8s cabin compressor oil by the accessory oil cooler's oil-to-oil heat exchanger. APU

9650

SMOH

8650

SMOH

Propeller Model: Aero Products Reverse Thrust The propeller reversing system is designed to provide increasing reverse thrust as the power lever is moved toward zero power lever position where maximum reverse thrust is obtained. Time required to obtain full reverse from FLIGHT IDLE is approximately four seconds. By lifting the power lever out of the FLIGHT IDLE detent with the landing gear scissor switch relay energised the propeller stop solenoid is energised, allowing the propeller control lever to move into the taxi range. This de-energises the propeller solenoid valve relay and electrically disconnects the phase synchronising units. The propeller blade angle and fuel flow then become a function of power lever position. Moving the power lever toward the MAXIMUM REVERSE position gives the desired reverse thrust for deceleration.

Collins Public Address System with 7 speakers Dual Collins (VIR 31A) Navigation Receivers One Collins DF-206 Automatic Direction Finder System Dual Collins DME-40 Distance Measuring Equipment (DME) Units Two Transponder/Encoder Systems with full Dual TCAS Bendix Weather Radar Radar Altimeter JET Standby Instruments (Gyro) Two Collins 200 HF Radios GPWS King KLN 90B GPS Flight Controls - General The Aircraft is equipped with a flight control system that utilises push-pull rods, cables and torque tubes connected to conventional cockpit controls. Aileron and rudder control systems are inter-connected to improve lateral stability during low speed flight. Trim tabs are provided for all axis of flight and a servo tab system is used to reduce the force required to move corresponding control surfaces. Aileron, elevator and rudder trim tab controls and position indicators are installed on the control pedestal, and each trim tab control is installed with its axis parallel to the axis of trimming motion. Curtain seals are installed ahead of the rudder and elevator control surfaces. This increases control efficiency by preventing airflow through the slots between the control surfaces and the stabilisers.

Pressurisation and Air Conditioning The cabin pressure and air conditioning systems provide cabin ventilation, pressurisation and temperature control during flight. These functions are normally automatic; however, ventilation and temperature may be controlled on a selective, non-automatic basis when required or preferred. Pressurisation is provided by the cabin air compressor on the right engine. The compressor receives air from an opening in the engine inlet duct. The compressed air flows through the firewall shut off valve to the area under the floor. It is then forced through heat exchangers, a secondary compressor, a refrigeration unit, a water separator and temperature control valves to outlets in the fuselage. The air normally vents overboard through the outflow valve which restricts flow to provide cabin pressurisation. Cabin Pressurisation - General The cabin pressurisation system is designed to supply all compartments used by personnel with an airflow of 73 pounds of air per minute at sea level. The Aircraft holds a sea level cabin altitude to an airplane altitude of 8,900 feet. At altitudes above 8,900 feet, a maximum pressure differential of 4.16 psi is maintained. This results in an 8,000 foot cabin altitude at 20,000 feet and a 10,000 foot cabin altitude at 23,000 feet. Presetting of controls permits operating at high rates of climb or descent with a minimum rate-of-change of cabin altitude. Airflow is maintained automatically against all normal loads imposed upon it by the constantly changing demand for pressurisation and refrigeration. Outflow from the cabin is controlled by a pressure regulator and outflow valve to maintain pressurisation. In the event of regulator malfunction, a combination dump, pressure relief and vacuum relief valve limits the pressure to 4.21 psi. The fuselage structure is designed for an ultimate differential pressure of 8.9 psi. The pressurisation area is effectively sealed to reduce air leakage to a minimum. Cabin Air Conditioning system Cabin temperature control is effected automatically at any selected temperature within a range of 16°C - 27°C (60°F - 80°F). Cooling is obtained from the engine-driven compressor (EDC). Heating is obtained from the heat of compression and can be supplemented by an electric heater. This heater incorporates elements electrically connected for combination of 4, 3, 2 and 1 kW which supply heat in ten increments. The electrical heater is located in the ducting between the heat exchanger and the cabin air distribution ducting. Note: To operate the electric heater, both the left- and righthand alternators must be operating; therefore, both engines must be operating at high rpm. A refrigeration unit bypass valve allows regulated amounts of the EDC air to bypass the refrigeration unit and the secondary heat exchanger. The refrigeration unit is powered by the compression force of the EDC air and incorporates an expansion turbine which can cool the air substantially below outside air temperature. Positioning of all the modulating valves is determined by the electronic temperature control unit. The heat generated in the EDC air, as it is compressed, is available for cabin temperature control. This is usually sufficient to maintain a comfortable temperature. Should ventilation by alternate airflow be selected in flight, the air compressor output is bypassed, and ventilating air is obtained at the ram air scoop in the left wing leading edge. The air then passes through the primary heat exchanger but bypasses the refrigeration unit and secondary heat exchanger on its way to the ventilation outlets. In this condition, cooling below outside air temperature is not possible; however, heating air is available from the electrical heater. Heating or cooling is available during ground operation if the GTCP is operating. Cooling air across the heat exchangers is provided by 2 electric fans. Electronic Temperature Regulator The cabin temperature regulator automatically regulates the cabin temperature within the limits of the air conditioning system. It accomplishes this function by positioning the refrigeration bypass valve and valve that controls the flow of cooling air across the primary heat exchanger. Two temperature anticipators operate in conjunction with the temperature regulator to sense changes in refrigeration discharge air temperature and ambient air temperature. Cabin temperature variation is sensed by a cabin pickup thermostat. This thermostat also transmits a signal to the temperature regulator. Electrical Systems The Aircraft is equipped with two main electrical systems. One is an alternating current (AC) system and the other is a direct current (DC) system. The AC system is used primarily to supply the larger electrical requirements of the propeller feathering motors, the electrical anti-ice and de-ice systems, and the cabin resistance heating elements. In addition, it has been designed to furnish AC power during normal flight conditions for aircraft flight instruments, certain engine instruments, fuel temperature trim controls, and the torquemeter. A transformer rectifier is also incorporated which can, when selected, supply DC power to the Aircraft DC essential distribution bus and DC essential bus, should trouble develop in the Aircraft DC system. The DC system is used primarily to supply the electrical requirements of the Aircraft electrical components. An inverter, powered by the DC system, can supply AC power for essential flight and engine instruments should trouble develop in the AC system in flight and prior to starting the engines on the ground. Airplane Lighting - General Exterior lights for landing, taxi, wing illumination, position and ant-collision purposes are provided. Strobe lights are provided at the wing tips. Interior lights include: pedestal, dome, compass, instrument panel, switch panels, door warning, cabin aisle and reading map and emergency lights. Fuel system Each wing has a fuel-tight compartment which serves as a fuel tank for the respective engine installed on that wing. Fuel capacity is 848 imperial gallons in each wing tank. Refuelling is over wing. Integral fuel tanks in the wings are joined only by crossfeed plumbing. Fuel is supplied to the engines by engine driven pumps and electrical boost pumps. Electrically operated shutoff valves are provided at the tank outlets. Mechanically actuated shutoff valves are provided at the firewalls and are operated by the emergency handles on the firewall control panel. Failure of the fuel system in one wing does not affect the system in the other wing. Either system is capable of supplying one or both engines through the crossfeed. Hydraulic System The hydraulic system supplies fluid under pressure to the following hydraulically-operated units: 1. Landing gear 2. Wing Flaps 3. Brakes 4. Nose wheel Steering 5. Windshield wipers 6. Main entrance door and stairs The main system pressure is provided by two variable displacement pumps, one engine-driven pump mounted on the left reduction gear box, and one 280V AC electric-driven pump mounted in the right wheel well. A DC electrically-driven auxiliary or emergency pump with independent fluid supply is installed for use in the event of failure of the main system or when on the ground with both engines stopped. A compressed air system is provided for emergency release of the landing gear up-latches and emergency braking. Brake System Goodyear Single Disc brakes are used with one brake assembly for each main landing gear wheel. A brake shuttle and lockout valve is installed in each wheel well. Its function is to prevent loss of pressure should a single brake or brake line failure and to prevent air from entering the hydraulic system when the emergency air brake is used. Anti-Ice and De-Icing The Aircraft employs both hot air and electrical ice control methods. The following table shows the surfaces which are ice controlled and the method employed: Ice Controlled Surfaces Location Method Airfoil (wing and tail) Hot Bleed air from the engine Leading edges (full span)compressor diffuser (14th stage) Bleed air from the engine Engine and aft inlet duct Hot bleed air from the compressor diffuser (14th stage) Propeller and inlet anti-icing and de-icing Electric heating element (from alternator) Pitot heads Electric heating element(from the essential DC bus 28 volts) Windshield Electric heating element (Nesa)(from alternators) Window defrosting Conditioned air from cockpit heater (from generators main DC bus) Automatic Airfoil Anti-Ice System The Aircraft are provided with an automatically controlled airfoil anti-icing system. An automatic temperature control causes the firewall bleed air shut-off valves to modulate as necessary to maintain airfoil temperatures at approximately 210° C. Two over-temp lights are provided to illuminate if the automatic modulation system allows the temperature to reach 232° C in the ducts. Also, just aft of the ducts, six limit switches are installed in the spar to provide structural temperature over limit protection for the leading edges and ducts and are reset to de-energise the anti-ice button if an overheat of 121° C occurs aft of the duct. Fire Detection and Extinguishing Fire Zones Fire Detection system

Lavatory, Galley and Oven