CHAPTER 6
NOTE: To convert Fahrenheit to Celsius (centigrade), use 5/9(F–32). For example, –4°F is converted as: 5/9(–4 – 32) = 5/9 X –36 = –20ºC. Celsius to Fahrenheit is converted using 9/5°C + 32. For example, –55°C is converted as: 9/5(–55) + 32 = –99 + 32 = –67°F. The aircraft manufacturer considers these varia- tions in outside air temperature and atmospheric pres- sure when designing the aircraft. ATMOSPHERIC CONSIDERATIONS Pressurization and air-conditioning of aircraft are necessary at high altitudes. With operational ceilings now in excess of 50,000 feet, flight personnel, and in some cases aircraft components, are supplied with an artificial means of maintaining a reasonable pressure around the entire body and/or equipment. This is done by sealing off the entire cabin/cockpit and any equipment area that may require pressurization and maintaining an inside air pressure equivalent to that at substantially lower altitudes. This is known as a pres- surized cabin, cockpit, or compartment, as applicable. In addition to pressurizing them, the cabin, cockpit, and some compartments are also air-conditioned, if the aircraft is to fly at high speeds. This requirement is partly due to the difference in temperatures at various altitudes and also to aerodynamic heating. For example, an aircraft flying at supersonic speeds at an altitude of 35,000 feet may generate a temperature on its skin of 200°F, and twice that temperature at altitudes near sea level. In addition to aerodynamic heating, other factors affecting cabin/cockpit temperatures are engine heat, heat from the sun (solar heat), heat from electrical units, and heat from the body. Through research and test, it was determined that the average total temperature of these five heat sources will raise cabin/cockpit temperature to approximately 190°F (88°C). Through experiments it was determined that the maximum temperature that a person can withstand and maintain efficiency for extended periods is 80°F (27°C); therefore, air-conditioning of the cabin/cockpit area is just as essential as pressurization. Under low-speed operating conditions at low temperature, cabin/cockpit heating may be required. The proper operation of much of today’s aircraft electronic equipment is also dependent on maintaining a reasonable operating temperature that will prolong the life of various components. In most cases, equipment cooling is provided by teeing off the ducting from the cabin/cockpit system. On other aircraft, a separate cooling system may be used primarily for equipment cooling. Q6-1. What is the atmospheric pressure at sea level? Q6-2. As an aircraft ascends to higher altitude, the decrease in atmospheric pressure may affect flight personnel. What is the most noticeable effect? Q6-3. The atmospheric pressure above 35,000 feet is extremely low. This condition may cause what effect on the human body? ENVIRONMENTAL CONTROL SYSTEMS LEARNING OBJECTIVE : Recognize the need for environmental control systems (ECS). The environmental control systems of most aircraft include cabin air-conditioning and pressurization, equip- ment cooling, defogging, windshield washing and rain removal, and equipment pressurization sub- systems. Coverage in this chapter is limited to air cycle air-conditioning and pressurization. There are five requirements necessary for the successful functioning of a pressurization and air-conditioning system: • The cabin must be designed to withstand the necessary pressure differential. This is primarily an airframe engineering and manufacturing problem. • There must be a means of limiting the maximum pressure differential to which walls will be subjected. This is provided by the cabin safety valve. • The aircraft must have an adequate supply of compressed air. This is provided through the compressor section of the jet engine. A separate compressor or supercharger is used on aircraft having reciprocating engines. On all jet aircraft, the air is taken directly from the compressor section of the jet engine. This is generally referred to as bleed air. • There must be a means of cooling the bleed air before it enters the cabin. This is provided by an aircraft refrigeration unit. • There must be a means of controlling the cabin pressure. This is provided by the cabin pressure 6-2
regulator, which regulates the outflow of air from the cabin. In addition to the major components, various valves, controls, and other related units are neces- sary to complete an aircraft pressurization and air-conditioning system. The design, construction, and use of these components may vary somewhat with different manufacturers; however, the systems on all jet aircraft operate on the same principles. Q6-4. List the five requirements necessary for successful functioning of a pressurization and air-conditioning system. Q6-5. List three systems that are included in the environmental control system. AIR CYCLE AIR-CONDITIONING SYSTEMS LEARNING OBJECTIVE : Recognize the components and operating principles of air cycle air-conditioning systems (ACS). Most naval aircraft are designed with an air cycle ACS because it is efficient for the weight and space required and is relatively trouble-free. The name air cycle or air-to-air comes from the principle of cooling the air without the use of refrigerants by compression and expansion of bleed air. The P-3 air cycle ACS is an example of this type of system. DESCRIPTION The P-3 air-conditioning system is comprised of two independent air cycle cooling systems of identical capacity, each with its own temperature control system, and fresh air sources. Fresh air sources are comprised of two engine-driven compressors (EDCs) and the air multiplier package (AMP). Fresh Air Sources In order for the air-conditioning system to function, air at the proper temperature and flow volume must be available. The fresh air sources are the EDCs and the AMP. The EDCs are single-stage compressors with fully automatic controls. They supply air to the air cycle cooling systems during flight and are operable only when no. 2 and no. 3 engines are running. The EDCs also serve as a secondary air source for the air cycle cooling systems during ground operations. The AMP is the primary air source for the air cycle cooling systems only during static ground operations when the auxiliary power unit (APU) is running. The AMP interacts with the APU to such an extent that it is referred to as the APU/AMP combination. EDCs The P-3 aircraft has two EDCs (fig. 6-1), to supply air to each of the two air cycle cooling systems. During EDC operation, there is no interconnect between the flight station and cabin systems until well downstream in the air distribution and exhaust system. The no. 2 engine EDC supplies air to the right (flight station) air cycle cooling system and the no. 3 EDC supplies air to the left (cabin) air cycle cooling system. The duct crossover is in the APU compartment and allows the ducts some flexibility for expansion. The EDC is mounted to a drive pad on the left side of the engine reduction gearbox assembly. The EDCs are adjusted for a maximum power requirement of 81 horsepower (hp) to deliver 60 pounds of air per minute at sea level. APU/AMP The APU/AMP combination supplies air to the air cycle cooling systems during ground operation only. It 6-3 Figure 6-1.—EDC view from left.
serves as a single source of air with a flow rate equal to that supplied by the two EDCs. With the engines operating at normal revolutions per minute (rpm), each EDC supplies air to its respective air cycle cooling system at the rate of approximately 60 pounds per minute (lb/min). The APU/AMP combination supplies air to a duct common to both air cycle cooling systems at a rate of approximately 125 lb/min. The air volume divides in the air cycle cooling system interconnection duct, with half going to the flight station air cycle cooling system and half to the cabin air cycle cooling system. An air-conditioning system that employs a single source of air to supply two air cycle cooling systems that operate at different back-pressures will have air flow problems unless a control is added to balance airflow. If airflow is not properly balanced, the air cycle cooling system with the lower back-pressure (as the result of more air bypass) will rob air from the unit with the higher back-pressure. Two flow-limiting venturis are used to balance airflow when the APU/AMP combination is the air supply source. Figure 6-2 shows an AMP installation. COMPONENTS Components include a heat exchanger package, turbine refrigeration unit, water separator, water spray system, and a flow-limiting venturi. Heat Exchanger Package The function of the heat exchanger package (fig. 6-3) is to reduce the temperature of the supply air furnished by the EDC or AMP. Two heat exchanger packages, each consisting of a primary and secondary section, electric fan assembly, check valve, and ram air duct check valve are installed on each side of the nose wheel well. The left heat exchanger package supplies air for the cabin systems and the right package cools flight station air. 6-4 Figure 6-2.—AMP installation.
The heat exchanger unit is constructed of a brazed core, which contains a series of metal plates separated by layers of fins that form a passage for cooling air and separate passage for supply air. During ground operation, the fan assembly installed on the heat exchanger forces ambient air through the heat exchanger and overboard. A check valve, installed at the fan outlet, directs ambient airflow in the heat exchanger. Another check valve, located in the ram air inlet duct, closes, preventing ambient air from spilling overboard through the ram air duct check valve. The heat exchanger check valve closes, and the ram air is used to cool the supply air. Turbine Refrigeration Unit Each air cycle cooling system has a turbine refrigeration unit (fig. 6-4) installed on each side of the 6-5 Figure 6-3.—P-3 air-conditioning system schematic diagram. Figure 6-4.—Turbine refrigeration unit.
aircraft nose wheel well. The refrigeration unit, along with the secondary section of the heat exchanger, lowers the temperature of supply air so that it may be used for aircraft cooling. The refrigeration unit consists of a rotating assembly and a housing assembly. The rotating assembly mounts a turbine scroll and a compressor scroll, enclosing the rotating assembly. The bottom of the bearing support housing forms a sump for lubricating oil. A sight gauge is provided in the sump for determining the level of lubrication oil. Each of the three sections formed by the housing assembly is sealed to prevent air and oil leakage. Compressed supply air, after it has passed through the primary section of the heat exchanger, is ducted into the compressor section of the turbine where it is further compressed as it passes through the compressor scroll. The compressed air, with a temperature slightly above that of ambient air, is then routed to the secondary section of the heat exchanger, where it is cooled to a lower temperature. Returning from the heat exchanger, the compressed air enters the turbine scroll, expanding as it flows from the nozzle through the turbine wheel to the outlet duct. As the air is expanded, it drives the turbine wheel at high speed. Mechanical energy, which is extracted from the air to drive the turbine wheel, is transmitted to and absorbed by the compressor wheel. This mechanical energy reduces the supply air pressure and temperature to the point where the air becomes usable for aircraft cooling. Lubrication of the rotating assembly bearings is accomplished by an air-oil mist. Oil is absorbed by wicks, which extend from the oil sump to the shaft of the rotating assembly. Oil is distributed on the rotating assembly shaft as a result of capillary action in the wicks. Rotation of the shaft causes the oil to diffuse into an air-oil mist. The action of the oil slingers causes the air-oil mist to pass through the bearings, providing lubrication. Water Separator The water separator (fig. 6-5) removes moisture from the air before it is distributed within the aircraft. Two water separator units are installed in the APU compartment. The cabin system unit is located in the aft upper left section of the APU compartment and the flight station unit is located in the forward upper right section. The water separator consists of a condenser assem- bly and a collector assembly. The condenser assembly is a coalescer. An ice-limiting sensor is installed in the inlet section of the unit to protect against icing and a check valve is installed in the water separator outlet to prevent reverse airflow through the unit. 6-6 Figure 6-5.—Dual check valve, water separator, and ice-limiting sensor.
As supply air passes through the coalescer, moisture particles are condensed into droplets. After the air has passed through the coalescer, hundreds of small vanes create a swirling motion of the air and the airborne water droplets. This swirling motion centrifuges most of the water droplets from the air into the coalescer sump, where the water accumulates and drains overboard. The air, relieved of approximately 70 percent of its moisture, is then ducted into the aircraft and distributed. Water Spray System The water spray system increases basic cooling capacity of the air cycle cooling system by spraying water separator discharge water into the ram air, cooling it by evaporation before the ram air passes through the heat exchanger’s secondary section. Flow-Limiting Venturi Each air cycle cooling system has a flow-limiting venturi installed in the left and right sides of the APU compartment in the air distribution duct between the AMP and the EDC air ducts. The venturi is sized to limit airflow to 67 lb/min from the AMP to the air cycle cooling system to ensure proper flow division and to prevent excessive flow through the aircraft during the heating mode. It functions to limit flow through the refrigeration unit in the event the other refrigeration unit is operating at a different bypass setting; that is, one refrigeration unit is in maximum cooling while the other is modulated toward heating. A check valve is located in the outlet of each venturi to prevent reverse EDC airflow through the venturi. CABIN AND FLIGHT STATION TEMPERATURE CONTROL SYSTEM There are two independent temperature control systems on the aircraft designed to control the temperature of the air cycle cooling system output air. Each system (fig. 6-6) is composed of a selector-indicator, a temperature controller, a master temperature sensor, a duct rate sensor, an ice-limiting 6-7 Figure 6-6.—Temperature control system components.
sensor, a pressure ratio limiter, and three airflow control valves. Air supplied from the ground-air connection or from the auxiliary ventilation system does not pass through the air cycle cooling system, so the aircraft temperature control system has no control of incoming air from these sources. Temperature control is achieved, either automati- cally or manually, through the three modulating valves. These valves route supply air through the various cooling components to produce the desired output air temperature. The three airflow control valves are driven by servomotors that are controlled by servo-amplifiers. The airflow valve schedule is the same for both manual and automatic modes of operation. The basic difference in operating modes is the method of applying sensing control to the servo-amplifiers. Temperature Control System Selector-Indicator Air-conditioning system control input signals are selected with the selector-indicator. The selector-indicator contains a push-pull knob, a cursor at the edge of the indicator face, an indicator needle, and an indicator flag. Sets of dot markings, one-dot, two-dot, and three-dot, are on the indicator face to facilitate temperature or program selection. The mode of operation is selected by moving the push-pull knob: push for manual operation, pull for automatic. The indicator flag indicates which mode has been selected (MAN or AUTO). The system control voltage (manual) or operating temperature (automatic) is selected by rotating the push-pull knob, clockwise for warmer, counterclockwise for cooler. This moves the cursor at the edge of the instrument face to indicate the program position (automatic) or temperature selection (manual). The indicator needle in the center of the dial is the indicator of the voltmeter that is connected to the temperature controller circuitry. There are three potentiometers inside the selector-indicator assembly. In manual mode, each potentiometer provides a command signal to each of the three airflow control valves (valves A, B, and C) by way of the temperature controller. When an automatic mode is selected, valve A potentiometer provides the temperature controller with the command signal for all three airflow control valves. Temperature Controller The temperature controller is the heart of the temperature control system. It receives and integrates the signals from the selector-indicator, master temperature sensor, duct rate sensor, ice-limiting sensor, and pressure ratio limiter. These signals are used to position the three airflow control valves. The temperature controller is composed of four modules: a programming amplifier module and three transistorized servo-amplifier modules. The pro- gramming amplifier contains the temperature control system automatic mode control circuitry. When automatic mode is engaged, this module integrates the sensor signals with the selector-indicator command signal, and produces the appropriate command signals for the three valve servo-amplifiers. In manual mode, the programming amplifier contributes nothing to system operation. The servo-amplifiers control the operation of the airflow valves and are identified as valve A, B, and C servo-amplifiers. In manual mode, the servo-amplifiers respond to command signals from the three potentiometers in the selector-indicator. In automatic mode, command signals come from the programming amplifier module. Master Temperature Sensor A master temperature sensor is mounted in each system exhaust air duct to sense cabin or flight station air temperatures. The flight station master temperature sensor is in an exhaust duct above and aft of the pilot position. The cabin temperature sensor is in the ex- haust duct above the tactical coordinator (TACCO) station. In automatic mode, the sensor senses the temperature within the aircraft. The heart of the sensor is the thermistor, whose electrical resistance varies inversely with its temperature. The changes of resistance provide reference signals to the temperature controller. The controller combines these signals with the signals from the duct rate sensor and the selector-indicator, and produces a command signal. The command signal is used by the temperature controller to regulate the temperature within the aircraft, by positioning the airflow control valves. Duct Rate Sensor A duct rate sensor is installed in the duct upstream of each ice-limiting sensor. It senses the temperature output air from the air cycle cooling system and the hot air bypass valve (valve C). In automatic mode, the duct rate sensor signals are integrated with the master temperature sensor and selector-indicator signals to regulate system temperature. 6-8
Ice-Limiting Sensor The purpose of the ice-limiting sensor is to eliminate ice formation in the water separator. An ice-limiting sensor is installed in the inlet of each water separator. Electrical signals from the ice-limiting sensor are sent to the programming amplifier in the temperature controller. The programming amplifier directs the signals to valve A servo-amplifier. The servo-amplifier directs the signals to the servomotor that drives the valve toward an open position. This allows warm air to enter the water separator, which eliminates ice formation in the water separator. The ice-limiting sensor is operative in the manual or automatic mode. Ice formation on the water separator coalescer causes a pressure drop between the water separator inlet and outlet, which is sensed by the ice-limiting sensor. As this pressure drop increases to 2.9 inches of mercury, low-pressure (yellow) relay K3 is actuated. This relay removes voltage to valve A servo-amplifier, which allows only an opening signal to be received. If ice buildup increases differential pressure to 4.1 inches of mercury, high-pressure (red) relay K5 is actuated. This relay removes servo-amplifier signals routed to valve A and supplies a signal to the valve in the open direction only. This causes the valve to open, allowing hot air to enter the water separator and circulate, thus removing ice. Pressure Ratio Limiter Assembly The pressure ratio limiter is mounted on the EDC, and is part of the EDC surge control system. Its function is to relieve or eliminate that part of the total EDC back pressure imposed by the turbine refrigeration units and water separators when the EDC is operating at maximum capacity or is overloaded. It is intended to function above 18,000 feet or during climb and descent. Airflow Control Valves There are three airflow control valves installed in each air cycle system. Their function is to control system air output. The turbine bypass valves (valve A) are located in the nose wheel well in the primary of each heat exchanger outlet to turbine bypass duct. The turbine shutoff valves (valve B) are located in the nose wheel well in the secondary section of each heat exchanger outlet to turbine inlet duct. The hot air bypass valves (valve C) are located in the forward left and right sides of the APU compartment. Valve C controls the amount of hot primary compressor discharge air that is bypassed around the air cycle cooling system. Valve A controls the amount of warm air that is bypassed around the bootstrap refrigeration unit. Valve B controls the volume of air flowing through the refrigeration turbine for cooling. The 3 1/2-inch diameter valves B and C are identical and have the same part number. Valve A has a diameter of 4 1/2 inches. Each valve assembly consists of a butterfly type valve, an alternating current (ac) servomotor, a planetary gear train, and a follow-up potentiometer. The servomotor, gear train, and potentiometer are combined into a single unit called the actuator assembly, which is mounted on the valve housing. The actuator assembly receives signals from the temperature controller to position the butterfly valve during temperature control system operation. TEMPERATURE INDICATOR The selectors, controls, and monitoring equipment for temperature control are mounted in the upper portion of the panel grouping (fig. 6-7). Control of cabin and flight station temperature is achieved through modulation of the two air cycle cooling systems. 6-9 Figure 6-7.—Air-conditioning, cabin air compressors, and cabin pressurization control panels.
The temperature indicator, TEMP °C, and the three-position TEMP SELECTOR switch are connected to read three different temperature sources as follows: • FLT STA COND AIR position. Temperature of the conditioned air leaving the air cycle cooling system, which is controlled to meet flight stations requirements. • CABIN COND AIR position. Temperature of the conditioned air leaving the air cycle cooling system, which is controlled to meet cabin requirements. • CABIN TEMP position. Temperature of the air leaving the cabin, exhaust air temperature, which is actual cabin temperature. OPERATION The P-3 aircraft is equipped with two temperature control systems, one for the flight station and one for the main cabin area. Each temperature control system consists of a temperature controller, a selector-indicator, a master temperature sensor, a duct rate sensor, and three airflow control valves. To operate the temperature control system, the flight crew sets the selector-indicator at the desired temperature (fig. 6-8). This information is transmitted to the temperature controller, along with signals from the sensors that provide the actual cabin or flight station temperature and the rate of temperature change at the water separator inlet. The temperature controller then positions the three airflow control valves in a programmed schedule to properly blend the hot, warm, and cool air flowing in the air cycle cooling system to obtain the selected flight station or cabin temperature. The temperature control system employs the basic air cycle cooling system, two valve-controlled bypass ducts, and one shutoff valve. The hot air bypass valve (valve C) controls the amount of hot primary compressor discharge air that will be bypassed around the air cycle cooling system. The turbine bypass valve (valve A) controls the amount of warm air that will be bypassed around the turbine refrigeration unit. The turbine shutoff valve B controls the volume of air flowing through the turbine refrigeration unit for cooling. When full cold is commanded, all of the airflow into the air cycle cooling system is being cooled in the turbine refrigeration unit. If the control system demands heating, warm or hot air will bypass various portions of the air cycle cooling system until the desired temperature is obtained. The valve B and C operation schedules are similar in response, but opposite in direction. On the other hand, valve A has a relatively complicated operation schedule. This is because valve A controls the bypass of air that is warm, but close to a comfortable temperature. Under certain atmospheric conditions, water separator bag icing will cause reduced airflow and high 6-10 Figure 6-8.—Selector-indicator.
back-pressure on the air cycle cooling system. High back-pressure from the air cycle cooling system will cause the pressure ratio across the EDC to exceed its design limits and cause a compressor surge condition. The compressor surge may cause a loss of cooling air at the dump valve or damage the EDC. The ice limit sensor, a pressure switch, senses the pressure differen- tial across the water separator. One side of the pressure switch senses the air pressure upstream from the water separator bag, and the other side senses the air pressure downstream from the water separator bag. When the bag is covered with ice, the airflow is impeded and a greater-than-normal pressure drop develops across the bag. When the pressure differential exceeds 2.9 inches mercury (Hg), a switch actuates in the ice limit sensor ice limiter indicating partial blockage (yellow condi- tion) of the water separator (fig. 6-9). If the water separator icing condition worsens and the ice limiter senses an increase in pressure drop across the water separator bag to 4.1 inches Hg, the high-pressure switch actuates in the ice limit sensor, indicating heavy block- age (red condition) of the water separator. These ice limiting sensor signals are used by the temperature con- trol system, during the automatic mode only, to control water during icing. Figure 6-10 shows the ECS flow. 6-11 Figure 6-9.—Air cycle cooling schematic diagram. Figure 6-10.—ECS block diagram.
Automatic Mode The automatic mode of operation regulates the cabin or flight station environment at a temperature setting between 65°F (18°C) (full cold, AUTO) and 85°F (29°C) (full hot, AUTO) depending on the setting of the selector bug. The face of the selector-indicator has reference marks on it at settings of 70°F (21°C) (one dot), 73°F (23°C) (two dots), and 80°F (27°C) (three dots). In the automatic mode, the temperature controller receives three inputs and uses them to determine the proper control valve positions. First, the flight crew uses the selector-indicator to select the desired temperature. This tells the controller what the flight crew wants. The master temperature sensor provides the second input to the temperature controller. This device senses the current flight station or cabin ambient temperature. The temperature controller amplifier compares the temperature requested by the flight crew (selector-indicator) with the actual cabin or flight station temperature (master temperature sensor), and develops an output called the program voltage. This signal is a dc voltage command for the three servo-amplifiers to drive the airflow control valves. The selector-indicator needle, called the program position indicator (PPI), is positioned by the command from the program amplifier. The PPI tells the flight crew what the system is going to produce, regardless of the position of the selector bug. The duct rate sensor senses the rate of temperature change in the air supply duct, then it sends a third signal that is proportional to this rate of change to the controller program amplifier. This signal enables the temperature controller to prevent temperature instability when the actual temperature approaches the desired temperature. The duct rate sensor is mounted in the system supply duct at the control valve blending location. This is the point where the hot air bypass, the warm turbine bypass, and the cold turbine refrigeration unit air discharges are blended. As the three airflow control valves move in response to changes in program voltage or position, the temperature of the air will change at a proportional rate. Manual Mode The manual mode of operation for the temperature control system is a backup mode in case the automatic mode fails. In the manual mode, the selector bug commands control valve position rather than setting a temperature. If the flight crew is uncomfortable, the valve positions must be changed by moving the selector bug. The PPI needle should follow the selector bug closely (within one needle width) as the bug is moved. The master temperature sensor, duct rate sensor, and program amplifier inputs are not used in the manual mode. Q6-6. The term air-to-air comes from what principle? Q6-7. The cabin air cycle cooling system EDC is mounted on what engine? Q6-8. What component balances airflow in case of back-pressure? Q6-9. The water separator removes what per- centage of moisture? Q6-10. What component is the heart of the tem- perature control system? Q6-11. When selecting full cold, all of the airflow into the air cycle cooling system is being cooled by what component? Q6-12. What are the modes of operation for the P-3 air cycle cooling system? AIRCRAFT PRESSURIZATION SYSTEMS LEARNING OBJECTIVE : Recognize the purpose and function of an aircraft pres- surization system to include maintenance and troubleshooting operations. As aircraft became capable of obtaining altitudes above that at which flight crews could operate ef- ficiently, a need developed for complete environmental systems. Air conditioning could provide the proper tem- perature and supplemental oxygen could provide sufficient breathable air. The one problem was that not enough atmospheric pressure exists at high altitude to aid in breathing, and even at lower altitudes the body must work harder to absorb sufficient oxygen through the lungs to operate at the same level of efficiency as at sea level. This problem was solved by pressurizing the cockpit/cabin area. PRESSURIZATION SYSTEM The area of an aircraft to be pressurized must be free from all air leaks. This is accomplished by use of 6-12
seals around tubing, ducting, bolts, rivets, and other hardware that pass through or pierce the pressure-tight area. All panels and large structural components are assembled with sealing compounds. Access and removable doors and hatches have integral seals. Canopies are constructed with inflatable seals. The pressurizing air is the air from the aircraft ACS. The S-3 aircraft incorporates a cabin pressurization subsystem. This regulates the outflow of air from the cabin to control the cabin pressures according to a predetermined schedule. Cabin air is drawn through the internal avionics racks by the cabin exhaust fan and is modulated by the cabin pressure regulator valve. A cabin pressure regulator control provides the pres- surization schedule. SYSTEM OPERATION The cabin pressurization subsystem is managed on the pressure regulator control, which provides five modes of operation: unpressurized, isobaric, differential cabin-to-ambient pressure, dump, and re-pressurization. The cabin pressure schedule is designed to satisfy the requirements of a maximum cabin pressure-to-ambient differential of 6.7 (±0.1) psi and a 5,000 feet cabin altitude at flight altitudes between 5,000 and 25,000 feet. The cabin is normally unpressurized below 5,000 feet. Table 6-1 shows cabin pressures and altitudes with actual flight altitude. During the unpressurized mode of operation, the pressure regulator control directs low-pressure air to the pressure regulator valve to command it to the full open position. This mode of operation occurs at all altitudes below 4,350 feet. In this mode, cabin pressure is maintained at a near ambient pressure. The pressure is slightly above ambient because of the duct pressure losses, the quantity of air flowing into the cabin, and the pressure across the internal avionics ventilation subsystem. During flight operations between 5,000 and 24,000 feet, the isobaric mode maintains the cabin altitude between 4,350 and 5,000 feet. The pressure regulator control, using the sensed ambient pressure as a low-pressure source and the sensed cabin pressure as the high-pressure source, modulates the pressure regulator open or closed to maintain cabin pressure at the specific altitude. The differential mode of operation overrides the isobaric mode when the aircraft is flying at altitudes in excess of 24,000 feet. As cabin-to-ambient differential pressure reaches 6.7 ±0.1 psi, a spring-loaded diaphragm in the pressure regulator control positions a poppet valve to supply this differential pressure as a control pressure to the pressure regulator valve. The 6-13 Flight Altitude (ft) Cabin Pressure Differential Cabin Pressure Altitude Min (psi) Max (psi) Min (ft) Max (ft) 0 0 0.25 –500 0 5,000 0 0.30 4,350 5,000 10,000 2.12 2.42 4,350 5,000 15,000 3.94 4.24 4,350 5,000 20,000 5.48 5.78 4,350 5,000 *24,300 6.60 6.80 4,500 5,000 25,000 6.60 6.80 5,000 5,380 30,000 6.60 6.80 7,400 7,870 35,000 6.60 6.80 9,600 10,100 40,000 6.60 6.80 11,520 12,050 *Maximum flight altitude for a 5,000 feet cabin altitude Table 6-1.—Cabin Altitude vs Flight Altitude Schedule
pressure regulator valve compares this control pressure to cabin pressure, and it positions the butterfly to maintain the required differential pressure. The cabin pressurization system also makes provision for dumping cabin pressure in an emergency. By setting the cabin pressure switch on the environmental control panel to the DUMP position, the latching solenoids on both the cabin outflow pressure regulating valve and on the cabin safety valve are actuated to the dump position. In addition, the re-circulation air shutoff valve will be actuated to the full open position, provided electrical power is available. A secondary method of achieving cabin depressurization is to turn the air-conditioning switch to the OFF/RESET position and select the auxiliary vent mode. This selection will cause the cabin outflow pressure regulator valve to open, but it will not actuate the cabin safety valve to the open position. The re-pressurization mode of operation is used when returning to the normal mode from the dump mode or during a rapid descent in excess of 4,000 feet per minute. In this mode, the pressure regulator control modulates the rate of cabin re-pressurization with an integral isobaric and differential pressure control system. The pressure regulator control compares the existing cabin pressure to a lagging cabin pressure reference. If the result of this comparison exceeds the calibrated rate, control pressure output from the pressure regulator control is reduced. This causes the pressure regulator valve to sense a relatively higher pressure on the opening side of its actuating diaphragm, allowing the diaphragm to open the pressure regulator valve butterfly. This reduces cabin pressure and the rate of pressurization. Precautions for operating the S-3 cabin pressurization subsystem on the ground, where the elevation is 5,000 feet or higher, are required because the cabin pressurization subsystem does not have provisions for automatic depressurization. Therefore, the cabin will pressurize whenever the ground elevation is above 5,000 feet. To ensure adequate cooling of the internal avionics during operations at ground elevations above 5,000 feet, one of the following steps must be used: • Keep the cabin pressurized as in flight. • Set CABIN PRESS switch to DUMP to ensure a full-open pressure safety valve. • Turn AUX VENT selector to ON if outside air temperature is below 80°F, and open the cabin entry door to ensure an adequate supply of cooling air. COMPONENTS The S-3 cabin pressurization subsystem consists of five primary components (figs. 6-11 and 6-12). Four of them are shown in figure 6-11. The fifth component is located in the cockpit. Each component is discussed in the following paragraphs. If you are to troubleshoot effectively, it is important to know the relationship of each component to the system as a whole. Cabin Pressure Regulator Valve The cabin pressure regulator is a pneumatically actuated butterfly valve mounted in the cabin exhaust ducting downstream of the cabin exhaust fan. The butterfly is spring-loaded to the closed position. The pressure regulator valve consists of the butterfly valve, which is actuated by a pressure-controlled diaphragm, and a solenoid valve to control the air pressure on the 6-14 Figure 6-11.—Cabin pressurization subsystem schematic.
diaphragm. The solenoid valve is electronically con- nected to the cabin pressurization switch on the en- vironmental control panel (fig. 6-13). There are three ports leading into the pressure regulator valve diaphragm chamber. The first port is located on the spring-loaded closing side of the diaphragm. It admits pressure from the cabin pressure regulator control. The second port is the ambient vent port. It is also located on the spring-loaded closing side of the diaphragm. The third port is located on the opening side of the diaphragm. A sensing line is attached to the third port to connect the cabin pressure regulator control and the cabin pressure exhaust duct. The pressure admitted to the diaphragm through the third port is equivalent to cabin air pressure. The difference between them causes the pressure regulator valve to modulate between the open and closed positions. Cabin Pressure Regulator Safety Valve The pressure regulator safety valve is an independent, pneumatically operated, balanced type of poppet valve that limits cabin-to-ambient pressure differentials to 7.07 (+0.2 and –0.0) psi. If the difference between cabin pressure and ambient pressure reaches the calibrated limit, the change in pressure acting on the limit control diaphragm overcomes the metering valve spring-load and allows 6-15 Figure 6-12.—Cabin pressurization components. Figure 6-13.—Environmental control panel.
the metering valve to open. This also opens a passage in the cabin pressure safety valve head, which causes the head pressure to be slightly lowered. Since the cabin pressure is greater than head pressure, it opens the pressure-balanced main poppet to allow cabin air to be vented overboard. When the cabin pressure differential is restored to normal, the limit control metering valve closes, and the pressure safety valve returns to its normally closed position. Cabin Pressure Regulator Control The pressure regulator control is a pneumatic control that provides four modes of cabin pressure operation. In addition to the modes of operation, a test valve is included with three manually set positions (FLIGHT, DIFF ON, and ALL OFF). The test valve is normally lockwired in the FLIGHT position for all cabin pressurization modes. The DIFF ON position permits a ground test of the normal delta-P setpoint. The ALL OFF position permits a ground test of the set point of the pressure safety valve. These test are accomplished with pressure supplied by support equipment. Four pneumatic ports are provided on the pressure regulator control for interfacing with various sensed pressures and the pressure regulator valve. These ports are different sizes to prevent improper plumbing connections. The pressure regulator control contains an isobaric bellows, which is calibrated to maintain an aircraft cabin pressure of 5,000 feet while the aircraft is flying at altitudes between 5,000 and 24,000 feet. The isobaric bellows, which modulates a control pressure, uses cabin air as a pressure source and low pressure in the environmental control system compartment as a negative pressure. Control pressure is delivered to one side of the pressure regulator valve diaphragm, and cabin pressure is connected to the opposite side. Because control pressure is normally less than cabin pressure, the pressure regulator valve becomes more open to decrease cabin pressure. The pressure regulator control contains provisions for controlling the rate of cabin re-pressurization when recovering cabin pressure after using the cabin dump mode, or during a rapid descent in altitude. The control pressure modulated by the isobaric bellows is further modulated by the re-pressurization diaphragm to limit cabin re-pressurization to an equivalent 4,000 feet per minute change. The pressure regulator valve is held open until normal pressure characteristics are sensed. Cabin Low-Pressure Switch The low-pressure switch is installed below the center console to sense cabin absolute pressure. The normally open low-pressure switch closes at 13,000 (±500) feet and reopens at 11,000 (±500) feet. The CAB PRESS indicator light on the annunciator panel illuminates when the low-pressure switch closes. The indicator light goes off when the low-pressure switch reopens. Cabin Air Pressure Sensing Filter The air pressure-sensing filter is located in the line that connects the cabin exhaust air duct, the cabin pressure regulator control, and the cabin pressure regulator valve. The replaceable filter element, which is connected to the air sensing tube, is mounted with clamping rings on the fuselage frame. The filter element is a cylindrical plug of treated paper and fabric in a metal housing. The clamping rings confine the air entry to the dome-shaped end to trap the entry of tobacco tar and dust particles greater than 10 microns in diameter. MAINTENANCE AND INSPECTION Very little maintenance is required on most pressurization and ACSs other than making the required periodic inspections and operational checks. In most instances, a maladjusted or malfunctioning component simply must be removed and replaced. There are, however, certain components that require periodic servicing, cleaning, and inspection so the component will function properly and efficiently. Specific requirements for servicing, cleaning, and inspection are listed in the daily, postflight, and special/conditional maintenance requirement card (MRC) decks as well as the maintenance instruction manual (MIM) for each aircraft. Electrical Failures Since all pressurization and ACSs have electrically controlled components, maintenance of these systems must include the related electrical circuits. Although an Aviation Electrician’s Mate (AE) is generally called upon to locate and correct electrical troubles, the AME should be able to check circuits for loose connections, and even perform continuity checks when necessary. A knowledge of electrical symbols and the ability to read circuit diagrams is therefore necessary. Figure 6-14 6-16
illustrates the electrical symbols commonly found in schematic diagrams. Loose connections are located by checking all connectors in the circuit. A connector that can be turned by hand is loose and should be tightened hand-tight. A continuity check is simply a matter of determining whether the circuit to the valve or other electrically controlled unit is complete. To perform a continuity check, the connector at the electrically controlled unit is first disconnected. Then, with all necessary switches and circuit breakers closed, a test lamp is connected into the circuit at the electrical connector. The lamp indicates whether or not the circuit is complete. Continuity checks may also be made with the use of a multimeter, an instrument used for measuring resistance, voltage, or amperage. Troubleshooting Troubleshooting is the process of locating a malfunctioning component or other unit in a system or mechanism. For the AME, troubleshooting is an important responsibility and one which will require a lot of squadron time. When a malfunction is reported concerning any of the components or systems that are maintained by the AME, the difficulty must be located and corrected quickly. To troubleshoot intelligently, the AME must be familiar with the system(s) at hand, knowing the function of each component in the system and with a mental picture of the location of each component in the system in relation to other components, as well as the location of the component in the aircraft. This can be achieved best by studying the installation and schematic diagrams of the system found in the applicable MIM. Troubleshooting procedures are similar in practically all applications. The procedures covered in this section are adaptable to almost all aircraft systems. Auto mechanics use these steps to find and repair automobile malfunctions. The AME can use these procedures to find and repair malfunctions within aircraft systems. There are seven distinct steps to follow during troubleshooting, as follows: 1. Conduct a visual inspection. This inspection should be thorough and searching—checking all lines, linkages, and components for obvious damage, 6-17 Figure 6-14.—Electrical symbols.
evidence of leakage, looseness, security, material condition, and proper installation; and servicing when applicable. 2. Conduct an operational check. The mal- functioning system or subsystem is checked for proper operation. This may be done by using special support equipment such as the environmental control test set or by using aircraft power and equipment with the engine running. Each aircraft maintenance manual provides the steps to be taken in performing the operational checkout of all the aircraft’s systems. The operational checks and troubleshooting charts for each system are numbered so that when a malfunction occurs during a step in the operational checkout, the malfunction can be located under the same step number in the troubleshooting chart. The troubleshooting chart will provide a list of possible causes of the malfunction in the order of probability, along with a recommended remedy. In any case, the AME must check the system out thoroughly, observing proper operation, sequence of events, etc. 3. Classify the trouble. Malfunctions usually fall into three basic categories—electrical, mechanical, and/or improper installation. Using the information acquired in steps 1 and 2, the AME determines under which category the malfunction occurs. Proper use of the test set or multimeter will identify whether the trouble is electrical or mechanical. Use of the MIM when performing all maintenance tasks should prevent improper installation. Something affecting the flow of gas or liquid (as could be the case in a vapor cycle ACS) could be categorized as a combination electrical/mechanical failure. Most mechanical failures should be found on the visual inspection; however, drive shaft failure is not readily apparent until the valve is operated. In some cases it may even be necessary to disconnect the valve from the ducting so that the butterfly valve can be observed through the end opening. The position indicator on some valves can indicate that the valve is changing positions, which can be a false indication if the shaft is broken after the indicating mechanism, or if the butterfly valve was damaged in such a manner that the shaft would rotate without actually repositioning the valve. 4. Isolate the trouble. This step calls for sound reasoning and a full and complete knowledge of how the system and each component operate. During this step, the AME can make full use of their knowledge and the system schematics to trace system operation and systematically eliminate components. They can arrive at a reasonable conclusion concerning the cause of the malfunction based on facts and deductive reasoning. Usually the trouble can be pinned down to one or two areas. By checking each individual area or component, the trouble can be isolated. 5. Locate the trouble. This step is used to eliminate unnecessary parts removal, saving time, money, and man-hours. Once the AME has isolated the trouble to a certain area or component, a closer observation of the valve or component in operation should provide some obvious indication that it is not operating as specified in the MIM. If all evidence indicates that the problem is electrical, the assistance of an AE should be requested. 6. Correct the trouble. This step is performed only after the trouble has been definitely pinpointed and there is no doubt that the AME’s diagnosis is correct. Removal and replacement, or repair of the unit or system is done using the instructions provided in the applicable aircraft MIM. NOTE: While performing maintenance on any system, ensure the step-by-step procedures outlined in the MIM, including cautions, warnings, and safety notes concerning the specific procedures, are strictly complied with. 7. Conduct a final operational check. The affected component or system must be given an operational check following installation or repair to verify proper system or component operation. The MIM will provide the procedures for conducting the operational check. It will usually require operation of the system in various modes (manual and automatic for air-conditioning and pressurization systems) or through several cycles, as applicable. Specified steps throughout the repair procedure and operational check must be observed and certified by a quality assurance representative or a collateral duty quality assurance representative from the work center performing the work. These steps are usually identified in the MIM by underlining, italics, or some other obvious method. Q6-13. In order for the human body to operate at the same level of efficiency as at sea level, what solution was developed? Q6-14. The area of an aircraft to be pressurized must be free from all air leaks. How is this accomplished? Q6-15. For the S-3 pressurization system, what component provides the pressurization schedule? 6-18
Q6-16. During flight operations between 5,000 and 24,000 feet, what mode maintains the cabin altitude between 4,350 and 5,000 feet? Q6-17. When is the re-pressurization mode used? Q6-18. The S-3 cabin pressurization system will pressurize whenever ground elevation is above what altitude? Q6-19. The S-3 cabin pressurization subsystem consists of how many primary components? Q6-20. The pressure regulator safety valve limits cabin-to-ambient pressure differentials to what psi? Q6-21. The pressure regulator control maintains cabin pressure of 5,000 feet while the aircraft is flying between what altitudes? Q6-22. State the purpose of performing a continuity check. Q6-23. How many distinct steps should be followed during troubleshooting? Q6-24. Malfunctions usually fall into three basic categories. What are they? Q6-25. During troubleshooting, all evidence in- dicates that the problem is electrical. Whom should you call for assistance? 6-19
CHAPTER 6 ANSWERS TO REVIEW QUESTIONS A6-1. 14.7 psi A6-2. Breathing A6-3. The human body tends to burst. In some cases, blood vessels near the surface may burst, causing hemorrhages in the ears, eyes, and breathing passages. A6-4. 1. The cabin must be designed to withstand the necessary pressure differential. This is primarily an airframe engineering and manufacturing problem. 2. There must be a means of limiting the maximum pressure differential to which walls will be subjected. This is provided by the cabin safety valve. 3. The aircraft must have an adequate supply of compressed air. This is provided through the compressor section of the jet engine. A separate compressor or supercharger is used on aircraft having reciprocating engines. On all jet aircraft, the air is taken directly from the compressor section of the jet engine. This is generally referred to as bleed air. 4. There must be a means of cooling the bleed air before it enters the cabin. This is provided by an aircraft refrigeration unit. 5. There must be a means of controlling the cabin pressure. This is provided by the cabin pressure regulator, which regulates the outflow of air from the cabin. A6-5. The envir onmental control systems of most aircraft include cabin air-conditioning and pressurization, equipment cooling, defogging, windshield washing and rain removal, and equipment pressurization subsystems. A6-6. The name air cycle or air-to-air comes from the principle of cooling the air without the use of refrigerants by compression and expansion of bleed air. A6-7. No. 3 A6-8. Two flow-limiting venturis A6-9. 70 percent A6-10. Temperature controller A6-11. Turbine refrigeration unit A6-12. Automatic and manual A6-13. Pressurizing the cockpit/cabin area A6-14. By the use of seals around tubing, ducting, bolts, rivets, and other hardware that pass through or pierce the pressure tight area. All panels and large structural components are assembled with sealing compounds. A6-15. Cabin pressure regulator A6-16. Isobaric A6-17. When returning to the normal mode from the dump mode or during a rapid descent in excess of 4,000 feet per minute. A6-18. 5,000 feet 6-20
A6-19. Five A6-20. 7.07 (+0.2 and –0.0) psi A6-21. 5,000 and 24,000 feet A6-22. To determine whether or not the circuit to the component is complete A6-23. Seven A6-24. Electrical, mechanical, and/or improper installation A6-25. An AE 6-21
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CHAPTER 7 OXYGEN SYSTEMS A dependable supply of oxygen is an essential element for maintaining life. Oxygen systems aboard naval aircraft sustain the lives of the pilot and aircrew so they can perform their missions. AME personnel service and maintain aircraft oxygen systems. Therefore, it is important that AME personnel understand how and why oxygen systems function as they do. This chapter provides an overview of the operating characteristics and maintenance require- ments for several aircraft oxygen systems, stressing safety and use of the applicable maintenance instruc- tion manual (MIM). IMPORTANCE OF OXYGEN LEARNING OBJECTIVE: Identify the im- portance of oxygen to include types, charac- teristics, and effects of a lack of oxygen. No one can live without sufficient quantities of food, water, and oxygen. Of the three, oxygen is by far the most urgently needed. If necessary, a well-nourished person can go without food for weeks, living on what is stored in the body. The need for water is more immediate, but still does not become critical for several days. The supply of oxygen in the body is limited to a few minutes. When the supply is exhausted, death is inevitable. Oxygen starvation affects a pilot or aircrewman in much the same way that it affects an aircraft engine. Both the body and the engine require oxygen for the burning of fuel. An engine designed for low-altitude operation loses power and performs poorly at high altitudes. High-altitude operation demands a means of supplying air at higher pressure to give the engine enough oxygen for the combustion of fuel. A super-charger or compressor satisfies the engine’s demands. What about the demands of the human body? The combustion of fuel in the human body is the source of energy for everything the aviator is required to do with muscles, eyes, and brain. As the aircraft climbs, the amount of oxygen per unit of volume of air decreases, and the aviator’s oxygen intake is reduced. Unless the aviator breathes additional oxygen, the eyes, brain, and muscles begin to fail. The body is designed for low-altitude operation and will not give satisfactory performance unless it is supplied the full amount of oxygen that it requires. Like the engine, the body requires a means of having this oxygen supplied to it in greater amounts or under greater pressure. This need is satisfied by use of supplemental oxygen supplied directly to the respiratory system through an oxygen mask, and by pressurizing the aircraft to a pressure equivalent to that at normal safe-breathing altitudes, or both. For purposes of illustration, an aviator’s lungs are like a bag of air since the air in the lungs behaves in the same way. If an open bag is placed in an aircraft at sea level, air will escape from it continuously as the aircraft ascends. The air pressure at 18,000 feet is only half that at sea level; therefore, at 18,000 feet the bag will be subjected to only half the atmospheric pressure it was subjected to at sea level. For this reason, it will contain only half the oxygen molecules it had when on the ground. Similarly, an aviator’s lungs contain less and less air as the aircraft ascends, and correspondingly less oxygen. Thus the use of supplemental oxygen is neces- sary on high-altitude flights. Up to approximately 35,000 feet, an aviator can keep sufficient oxygen in the lungs to permit normal activity by use of oxygen equipment that supplies oxygen upon demand (inhalation). The oxygen received by the body on each inhalation is diluted with decreasing amounts of air up to approximately 33,00 feet. Above 33,000 feet and up to approximately 35,000 feet, this equipment provides 100 percent oxygen. At approximately 35,000 feet, inhalation through the demand oxygen system alone will NOT provide enough oxygen. Above 35,000 feet and up to 43,000 feet, normal activity is only possible by use of pressure demand equipment. This equipment consists of a super-charger arrangement by which oxygen is supplied to the mask under a pressure slightly higher than that of the surrounding atmosphere. Upon inhalation, oxygen is forced (pressured) into the mask by the system. Upon exhalation, the oxygen pressure is shut off automatically so that carbon dioxide can be expelled from the mask. Above 43,000 feet, the only adequate provision for the safety of the aviator is pressurization of the entire body. 7-1