Unformatted text preview: ATA-30 ICE and RAIN protection
Icing is caused primarily by the super cooled water droplets in the atmosphere which are at a temp below freezing pt water looses heat when it strikes say on a/c wing & engine air intake or a propeller. Rime ice: A rough and opaque ice that forms on aircraft while flying through visible moisture such as cloud when temperature is below freezing. It disturbs smooth air flow and adding weight. Glaze ice: That forms when large drops of water strike a surface whose temperature is below freezing. It is clear and heavy. Frost: Ice crystal deposits formed by sublimation when the temperature and due point are below freezing. Super cooled water: Water in its liquid form at temperature well below its natural freezing temperature when it is disturbed it immediately freezes. Sublimation: A process in which a solid material changes directly into a vapor without passing through the liquid state. Ice will not be formed above 40.000 feet (12,000 meter) height. Effects of icing: The building of ice can change the aerodynamic shape of a/c causing a decrease of lift. A change of trim due to change of wt loss of engine power damage to turbine engine blades. Icing also causes loss of forward vision due to ice forming on the wind shield panels. False reading of altimeter & airspeed can occur due to ice formation on pitot probes & static vents. That's is why ice protection is provided in the a/c protection from rain is also provided to improve the visibility in heavy rain by providing wind shield wipers and rain repellant system. There are two methods of ice protection provided in the modern a/c. 1. Anti-icer system: a system that prevents the formation of ice on the aircraft system. 2. De-icer system: a system that removes ice after it has formed on an aircraft structure. Ethylene glycol: A form of alcohol used as a coolant for liquid cooled engines and as an anti-icing agent. Isopropyl alcohol: A colorless liquid used in the manufacture of acetone and its derivatives and as a solvent and anti-icing agent. Methods of prevention of ice formation: y Breaking ice formation by inflatable rubber boots (pneumatic method). y Heating by hot air bleed from the engine. y Electrical heating. y Fluid spray by methane alcohol. Pneumatic deicing system: Uses rubber devices called boots or shoes attached to the leading edges of the stabilizers & wing. The boots consists of series of inflatable pipe or tubes during operation the tubes are inflated with pressurized air& deflated in the alternate cycle. This inflation causes the ice to break & blow off. The pneumatic air supply is obtained from an engine driven pump at a regulated pressure of 18PSI in turbo prop engines the air is taken from a tapping on compressor stage & then regulated during the deflated cycle the boots are held back by connecting to a vacuum line source to a distributor valves, A thin conductive coating is provide over the surface of boots to dissipate the static charge, Electrical anti-icing (or) deicing system: This consists of three principle sections with relevant control protection & indication. The power normally required 115/200V A.C from wild freq alternators & 28 volt D.C for control of the system heating elements vary in design & type depending in type of applications. For propellers they are fine wire type sand witched in insulating & protecting materials which form overshoes. For propeller turbine engine air intakes, leading edges of wings & helicopter rotor blades the elements of sheared foil type. The elements are normally of Ni, CUNI & Ni Cr. metal, adjusted by chemical etching. Electrical power of 200V A.C of variable freq supplied to propeller blades & the spinner through brushes & the slip rings & acyclic timing device so that deicing part of cycle the heat & applied to four blades simultaneously. It is not necessary o deice force keeps the outer halves free from ice. The air intake elements are aligned so that their elements are aligned so that their position in the WE are continuously heated to conform the anti-icing system. For those on inner side are supplied via the cyclic timing s/w to perform deicing. Wind shield system is necessarily anti-icing. In addition the temp of wind shield panel must be higher than the ambient temp during takeoff & flight at low attitude & landing for improving their impact strength against possible bird hit. Engine air intake and propeller anti-icing and de-icing. When the system is s/wed "ON" DC energizers the power relay via close contacts of the overload sensing device allowing 200V to flow directly to the continuous heating element. The unit is s/wed "ON" either fast (or) slow by selected the temp between -6 to +10° C & duration between heat 'ON' to heat 'OFF' is short when slow is selected in the at temp below - 6° C the heat 'ON7 heat 'OFF' is short. When slow is selected at temp below - 6°C the heat 'ON' to 'OFF' time is longer. The indication of the timer s/w operation is by blue / green light on the control panel. On the ground operation the Voltage is reduced to prevent overheating & this is effected by the automatic closing of micro s/w on to the L/G. In the event of A.C. overload the heats elements are protected by the sensing device which is actuated to interrupt a DC supply to the power relay. The current balance relay functions the same way but is actuated when there is an unbalance between the phases. Engine air intake anti-icing system (only DC operated) The circuit of D.C operated system when switched supply passing through the closed contacts of thermostats & through the contacts of oil pressure s/w which closely only above psi, the heater elements normally rated at 500 watts. The thermostat prevents the overheating of element by opening the circuit. When the element temperature exceeds 49 ± 3° C then the oil pressure s/w opens a circuit. The engine oil pressure drops to 50 ± 2 psi functioning is indicated by a indicator light which is illuminated by a current sensing relay which is in series with heater element & operation when the heater current is more than 15 amp. Propeller De-icing system: The propeller blades each have two heater elements bonded to them; one at the outboard section of a blade and the other at the inboard section. The elements are connected to the power supply via slip rings, brushes and an electrically-operated timer which is common to both propellers. The cycling sequence of the tinier is set so that - (I) the outboard elements of each propeller are simultaneously heated before the inboard elements, and (ii) only one propeller is de-iced at a time. The sequence for the right-hand propeller is shown at (a) and (b) of Fig. 10.28 respectively. The segments 3 and 4 respectively connect the supply to the outboard and inboard elements of the left-hand propeller. The timer energizes the elements for approximately 34 seconds and repeats the cycle as long as the control switch is in the "on" position. Operation of the system is indicated by the ammeter, the pointer of which registers within a shaded portion of the ammeter scale corresponding to current consumed (typically between 14 and 18 amperes) at the normal system voltage. Hot air bleed anti-icing system: This is a standard system adopted on larger transport a/c for anti icing of engine air intake nose cowlings wing L/E & L/E device such as Haps hot air is obtained from certain stages of main engine compressor & is then ducted through metal ducting through air intake & L/E's electrical power is solely required for the purpose of operating of motorized control valves in the ducting. The valve post indicating lights & the duct temp sensing devices the motors are limit s/w at full open & closed positions & in most application they are of controlled 115V. Single phase AC type 28V DC supply used for valve controls relay, s/wing & position indicating lights. The L/G micro S/w prevents the ground operation a ground test s/w is provided to check the operation of valve & indicator lights. A-300 Engine deicing, anti-icing by - bleed air Wind screen - electrical Flight controls - bleed air A - 310 & 320 Engine-bleed air Wind screen - electrical Flight controls - bleed air B-747 Engine - bleed air Wind shield - electrical Flight controls - bleed air Beach craft Wind screen - electrical Wing L/E - Pneumatic boots Horizontal stabilizer - pneumatic boots B-737 Engine-bleed airs, Wind electrical, Slats - bleed air. Wind shield anti-icing: The method adopted for wind shield anti-icing is normally thermal sensing consisting of a temp sensing element & a control unit. The element is embedded within the wind shield panel in such a way that it is electrically insulated from the main heating film and yet is capable of responding to temperature changes. The heating film is normally formed chemically by stenos chloride with gold spurting. A control unit comprises mainly bridge circuit of which the sensing element forms a part. An amplifier & a relay when the required power is s/wed on, initially the control unit relay is energized by an unbalanced bridge signal consequently the power control relay is energized to supply the w/s panel. As the panel temperature starts increasing the sensing element resistance also increases. A typical valve of 40°C makes the current to flow through the sensing element which balances the power control relay are de-energized there by interrupting the heating current power supply. As the temp cools down the sensing element resistance also decrease again to unbalance the bridge. The w/s are fitted with over heat sensing element which in the failure of sensing element takes over the function at higher voltage say 50°C. In some few types of A/c w/s are heated by resistance elements of fine wire supplied by 28 VDC but the temp control is by same bridge control. Tempered glass that has been heat treated to increase its strength. it Is used in bird proof heated wind shields for high speed aircrafts. The heat keeps the thermoplastic vinyl layers from becoming brittle and this prevents the wind shield from shattering if it should be struck by a bird inflight. Wind shield wiper system. The wiper arms & blades of each system are operated by 28 VDC series variable speed motor coupled to converter units. The motors are controlled by 4 position control s/w & the speed variation accomplished by dividing the resistance. In the low post the voltage applied to the field & armature & then to the ground through two resistors the motor runs at lower speed. When high is selected only one resistor come in circuit & motor will operate faster. When the operation of wiper not required the control is s/w is turned to the park position through the off position. These are no detent to the park position fit the s/w is manually held momentarily. In this case supply voltage is initially applied to motor in the normal way but now the connection to the ground is contacts of breaks s/w within the motor. Then it will run at fastest speed. As the wiper blade reaches its parked position the motor operates a cam to change over the break s/w contacts which then short out the armature & stop the motor. The s/w is then released to swing back to off position. The purpose of thermal s/w is to open the motor circuit if the field winding temp is exceeding 150CC or 8 - 10 amps of field current. Convertor changes rotary output of the motor into the reciprocating motion needed for the wiper blades. Testing is to be carried out after making the wind shield glass wet. Rain repellent system: The purpose of these systems is to maintain a clear area on the windshields of an a/c during takeoff, approach & landing in rain conditions. A system consists of a pressurized container of repellent fluid, control s/w a solenoid valve controlling the supply of fluid to a spray nozzle mounted in the fuselage skin in front of each wind shield. The fluid container is common to each wind shield system &: is located in he cockpit, When the control s/w is pushed in, a 28V DC supply is fed to the solenoid valve via the close contact 'B' of the control relay. The spray nozzle solenoid is energized to open the valve & allow fluid to flow under press through the spray nozzle & onto the wind shield the fluid is of a type which causes the surface tension in water to change so that blown off the windshield by the air stream. Through the action of time delay cct approx 5cc of fluid flows through the nozzle for approx 0.25 sec. At the end of this period the time delay cct applies power to the gate of an SCR which then energizes the control relay & in turn du energizes the spray nozzle solenoid valve. If the control s/w remains pushed in, the time delay cct will keep the control relay energized via a hold in cct across the closed contacts 'A' when the s/w is relayed the time delay cct & SCR are returned to their original state. The fluid is contained in a can which when screwed on to the mounting bracket opens a valve to allow fluid to drain into a reservoir & the system tubing. The reservoir is a clear plastic cylinder containing a float type contents indicator. A manually operated shut off valve is provided between the reservoirs & can & is used during can replacement. ICE DETECTION SYSTEMS Pressure Operated Ice Detector Heads. These consist of a short stainless steel or chromium plated brass tube, which is closed at its outer end and mounted so that it projects vertically downwards from a portion of the aircraft known to be susceptible to icing. Four small holes are drilled in the leading edge of this tube, and in the trailing edge are two holes of less total area than those of the leading edge (Figure 1). A heater element is fitted to allow the detector head to be cleared of ice. In some units of this type a further restriction to the air flow is provided by means of a baffle mounted through the centre of the tube. Hot Rod Ice Detector Head. This consists of an aluminum alloy oblong base (called the plinth) on which is mounted a steel tube detector mast of aerofoil section, angled hack to approximately 3(T from the vertical, mounted on the side of the fuselage, so that it can be seen from the flight compartment windows. The mast houses a heating element, and in the plinth there is a built-in floodlight The heating element is normally off and when icing conditions are met ice accretes on the leading edge of the detector mast. This can then be observed by the flight crew. During night operations the built-in floodlight may be switched on to illuminate the mast. By manual selection of a switch to the heating clement the formed ice is dispersed for further observance. Serrated Rotor Ice Detector Head. This consists of a serrated rotor, incorporating an integral drive shaft coupled to a small a.c, motor via a reduction gearbox, being routed adjacent to a fixed knife-edge cutter. The motor casing is connected via a spring-tensioned toggle bar to a micro-switch assembly. The motor and gearbox assembly is mounted on a static spigot attached to the motor housing, and together with the microswitch assembly, is enclosed by a cylindrical housing. The detector ii mounted through the fuselage side so that the inner housing is subjected to the ambient conditions with the outer being sealed from the aircraft cabin pressure. The serrated rotor on the detector head is continuously driven by the electrical motor so that its periphery rotates within 0-050 mm (0-002 in) of the leading edge of the knife-edge cutter The torque therefore required to drive the rotor under non-icing conditions will be slight, since bearing friction only has to be overcome. Under icing conditions, however, ice will accrete on the rotor until the gap between the rotor and knife edge is filled, where upon a cutting action by the knife edge will produce a substantial increase in the required torque causing the toggle bar to move against its spring mounting and so operate the micro switch, to initiate a warning signal. Once icing conditions cease, the knifeedge cutter will no longer shave ice, torque loading will reduce and allow the motor to return to its normal position and the micro switch will open-circuit the ice warning indicator. Vibrating rod (ultra sonic probe system) The fundamentals of operation are dependent on the phenomenon of magnetostriction, i.e., its sensing probe is caused to vibrate axially when subjected to a magnetic field at specific frequencies. The sensing probe is a ¼ inch diameter nickel alloy (Ni-Span C) tube mounted at its mechanical centre. The inherent resonant frequency of the probe is inversely proportional to its length, the simplified relationship being expressed as f =S/2l where, f = frequency in Hz, S = 1-88 x 10s inches per second (the speed of sound in the tube material) and L ~ length of the probe in inches. Based on this expression, the tube may be cut to a specific length to achieve a desired frequency; in this particular system the designed tube length is 2-3 inches, resulting in a resonant ultrasonic frequency of 41 kHz. This frequency, however, is reduced to a nominal 40 kHz by the brazing of heating elements within the tube and also by capping the tip of the tube. The probe is maintained in its axial vibration by the ultrasonic frequency excitation current produced by an oscillator and passed through a drive coil wound around the probe. The frequency is controlled by a feedback coil circuit such that the drive coil will excite the probe at whatever the natural frequency of the probe might be at the time. When ice forms on the probe the natural frequency is reduced, and the output frequency of the oscillator drive coil is in turn reduced to match the probe frequency. By means of a comparator circuit, the lower output frequency is compared with a fixed frequency output from a reference oscillator. The frequency difference between the two oscillators relates to the ice formation on the probe, and when the difference has reached a preset level (150 Hz or less) determined by a band pass filter and a limiting amplifier, a signal is sent to a switch and delay circuit. When this occurs, two timer circuits are triggered; one controlling the a.c. supply via a logic AND gate, to the probe heater, and the other controlling the duration an icing signal is available to an annunciator light for warning the flight crew. Thus, as will be noted from Fig. 10.30, there is a standing logic 1 input to the AND gate from the 115-volt bus, so when timer "A" is triggered it will supply a second logic 1 input to the gate causing it to switch on the heater for a period of 4.5 seconds. The signal from timer 4tB" is 28 volt d.c. and keeps the annunciator light illuminated for a period of 60 seconds. Melting of the ice from the probe increases the frequencies of the probe, and if no other icing signal is detected within 60 seconds, timer "A" automatically resets to isolate the heater from the a.c. supply. This cycle of operation is repeated while icing conditions prevail. Failure monitoring of the detector is accomplished with uni-junction oscillators which are set at both ends of the maximum difference frequency band. If the probe becomes severely damaged causing a significant change in the resonant frequency, or if an electronic component failure causes a malfunction in the reference frequency circuit, the annunciator light will be continuously illuminated. Ice Formation Spot Light (Ice inspection lamp). Many aircraft have two ice formation spot lights mounted one each side of the fuselage, in such a position as to light up the leading edges of the mainplanes, when required, to allow visual examination for ice formation NOTE: In some aircraft, this may be the only method of ice detection. ...
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- One '11
- power supply, wind shield