Electrical.docx

2A-24--10: General Electrical Power Purpose: The Gulfstream 650ER electrical power system (see Figure 1. Electrical Power System Component Locations and Figure 2. Electrical Power System Simplified Block Diagram) provides AC and DC power and a means of control, protection and distribution of...

2A-24--10: General Electrical Power Purpose: The Gulfstream 650ER electrical power system (see Figure 1. Electrical Power System Component Locations and Figure 2. Electrical Power System Simplified Block Diagram) provides AC and DC power and a means of control, protection and distribution of electrical power required for ground and in-flight operation of the aircraft. Primary electrical power is provided by the AC power system, comprised of two engine-driven Integrated Drive Generators (IDGs) and an Auxiliary Power Unit (APU)-driven generator. DC power is provided by five Transformer Rectifier Units (TRUs) and supplemented by two 53 Amp/Hour Ni-Cad Main Ship Batteries. A DC-powered Ground Service Bus (GSB) is provided to allow routine aircraft servicing without powering other aircraft systems. Selected electrical loads are controlled and protected by a solid state Secondary Power Distribution System (SPDS). A Ram Air Turbine (RAT) provides AC power for the flight control system and aircraft essential loads if the APU generator and both mainAC generators are not available. If all of the previously mentioned power sources fail, emergency power battery packs provide power to support equipment essential for safety of flight. There are also two 60 Hz converters to provide power to outfitter-installed equipment. General Description: Note Detailed descriptions of the AC POWER and DC POVVER 2/3 window synoptic display and 116 window AC I DC system display can be found in Section 2B-09-00: Engine Instruments and Crew Alerting System. AC Electrical Power System: The Electrical Power Generation System (EPGS) is based on constant-frequency Alternating Current (AC) power supplied from an IDG mounted on each main engine's HP stage gearbox and a constant-speed generator mounted on the APU. Constant frequency AC power from the IDG is obtained by converting variable input speed from the HP gearbox through hydro-mechanical means into a constant output speed which drives a 4-pole, 12,000 rpm generator that outputs fixed frequency, 400 Hz AC power. Both the IDG and APU generators provide power in the form of 115/200V AC, 400 Hz, 3-phase power and are rated at 40 kVA each. For ground operations during maintenance and flight preparation, power can also be derived from the APU generator (APU GEN) or from external AC or DC Power sources. Additionally, the APU generator can deliver sufficient power to both main AC buses to permit dispatch, with some restrictions,in the event of the loss of one IDG. The APU generator can serve as an alternate source of power during flight if one or both IDGs should suffer failures. DC Electrical Power System: For use of equipment requiring Direct Current (DC) power (except when the aircraft is beingsupplied by an external DC source via the external DC power receptacle), DC power is derived by converting AC power to DC power via unregulated TRUs. If all AC power is lost, theTRUswill be inoperative. DC power for essential flight functions is furnished by the two 53 Amp/Hour Ni-Cad aircraft batteries, designated as "Main Batteries". They are provided to start the APU, run the Auxiliary Hydraulic (AUX) pump, power battery bus loads and one (or both) of the Essential DC buses in the event power is not available from the Essential TRUs. The main batteries are normally used only for starting the APU or for powering the Ground Service bus in preparation for flight. Each battery has a high-capacity charger to return the battery to fullcapacity after use. An external DC power supply may be used on the groundfor powering the ground service bus for servicing, maintenance or preflight inspections. Other Power Systems: Secondary Power Distribution System (SPDS): The SPDS performs a control and protection function, effectively replacing elements of relay logic and traditional electro-mechanical circuit breakers. The control and protection functions are logically independent. High power AC and DC feeders provide the power source for the distribution units referred to as Modular Power Tiles (MPTs). Control of the system and the interface with other aircraft systems is performed by the Remote Interface Units (RIUs). Solid State Power Controllers (SSPCs) within the MPTs function as electronic circuit breakers to distribute power to the required electrical loads. Relay functions as well as select control functions reside in this system, thereby reducing the quantity of relays and separate LRUs. The system consists of two RIUs located under the floor boards in the pressurized cabin, three MPTs located under the floor boards inthe pressurized cabin, three MPTs located on the front of the LEER and two MPTs located on the aft side of the Aux Pressure Bulkhead. Approximately 400 loads are controlled from this system. (2) 60 Hz: A 60 Hz power system is installed to provide 115V AC, 60 Hz power for outfitter-installed equipment. On the ground, the 60 Hz system can be powered with a conventional external power cart using the service panel 60 Hz power receptacle. In flight a 60 Hz converter changes aircraft 400 Hz power to a normal household frequency level to power plug-in ouUets for the cabin and kitchen equipment in the galley. Distribution and protection of 60Hz power to the aircraft loads is performed by a 60Hz MPT within the SPDS. The 60 Hz Power System consists of two 3.5 kVA converters that convert 3-phase 115V AC 400 Hz power into single phase 115V AC 60 Hz. These converters are identified as Left and Right, butonly one converter is typically providing power at any given time. 50 Hz (Optional): Limited wiring provisions have been installed for a 50 Hz 250V AC converter. These provisions require removing the left 60 Hz converter and installing the 50 Hz converter in its tray. The 60 Hz MPT cannot support 250V AC so the wiring from the terminal strip in the forward Baggage Electronic Equipment Rack (BEER) to the 60 Hz MPT will require removal (capped and stowed). Ram Air Turbine (RAT): For non-time limited emergency power, a generator mounted on a RAT can provide 3-phase 115/200V AC, 324-440 Hz, emergency AC power to flight-critical loads while in flight and when primary AC power is not available. Operability of the RAT is limited by both minimum and maximum airspeeds. Deployment of the RAT is by manual means (handle/cable) only. Electrical Power System Controls and Indications: Pushbutton switches on the MAIN BATTERIES panel, ELECTRICAL POWER CONTROL panel, FLT CTRL BATTERIES panel, RAM AIR TURBINE panel and EMERGENCY POWER panel control the functions of the electrical system. The four panels are located together on the cockpit overhead. For simplicity in describing system operation, the panels will be referred to collectively as the ELECTRICAL PANEL throughout this chapter. The Crew Alerting System (CAS) supplies warning, caution and advisory messages to alert the crew to abnormal electrical system operation.2/3 synoptic displays for both AC and DC power and a AC I DC 1/6 system display provide views of electrical system status and configuration. Subsystems of the Electrical Power System: The electrical power system is divided into the following subsystems: 2A-24-20: AC Electrical Power System 2A-24-30: DC Electrical Power System Figure 1.Electrical Power System Component Locations LEGEND: 1) Rlf#lte-mlmDCnw.l'l:H1rarR8cllller'Ul'll RldrtMllnDC -Ull AUXT...-.bnnarRocllller'lkll 4) L.rt MllnDC RKCl'llir Urlt AC n..tonnerRedll..Unlt 7) ..._.,,._ _ (FOP) 8) APtl a...tor CbnlrU Uni B) Rlf#ltG....tDrcantralUnlt 10) Rlltit:EluaPar.rCamnill.lnl 11) Lall GaiwrmtorCortn:1Wrlt 12) LoftBuol'owwc.rt..!Urlt 13) Rlf#ltPOlllllrDllltJbJllDnBm 14) Loft-DldrlnOl..,Bm '.".' .... TIL-002985 Figure 2. Electrical Power System Simplified Block Diagram TIL-0053488 2A-24-20: AC Electrical Power System 1. General Description: Alternating Current (AC) power is produced by engine-driven Integrated Drive Generators (IDGs) and/or an APU driven generator. Generated power is routed over cables and wires to the electrical system buses where the power is distributed to components or subsystems.Distribution wiring on the electrical buses incorporates Circuit Breakers (CBs) to protect components and subsystems from damage due to over-amperage conditions. AC power is the primary electrical source, with Direct Current (DC) power derived from AC power by Transformer I Rectifier Units (TRUs).AC power is usually used by aircraft components incorporating some type of movement or motor function. Although DC power may also operate motor installations, it is normally used for control functions or steady-state operations. The AC electrical power system is composed of the following units and components: Engine-mounted Integrated Drive Generators (IDGs) Auxiliary Power Unit (APU)-mounted Generator Generator Control Units (GCUs) Bus Power Control Units (BPCUs) AC Power Distribution Buses 60 Hz Converters 50 Hz Converter (Optional) External AC power Ram Air Turbine (RAT) Static Inverter Description of Subsystems, Units and Components: Engine-Mounted Integrated Drive Generators (IDGs): The engine-driven generators are IDGs; "L GEN" for the left engine mounted generator and "R GEW for the right engine mounted generator. These generators provided fixed-frequency power by controlling their own speed via hydro-mechanical means internal to the IDG assembly and regulated by the unit's GCU. During normal operations, two 40 kVA IDGs, one driven by each engine, supply electrical power to the aircraft's main AC buses (left engine IDG to power the Left Main AC Bus, and right engine IDG to power the Right Main AC Bus). The aircraft engines each tum an IDG though a transmission mounted on the engine accessory gearbox at the compressor section of each engine. The IDGs are driven at a constant speed of 12,000 rpm (regardless of engine speed) through a hydro-mechanical transmission coupling the generator to the engine. A constant rotational speed enables the GCU to maintain AC power parameters of 3-phase 115/200V AC at 400 cycles-per-second (Hz) and deliver up to 40,000 volt I amps (40 kVA). The IDG rotates whenever the engine is operating. Under normal operating conditions, there is no provision for disconnecting the IDG from the engine while in flight, however; a thermal disconnect will decouple the engine reduction gearbox rotational power from the IDG in the event of a high oil temperature. This transmits the input shaft power to the IDG. An over-temperature condition in the IDG or external oil cooler circuit resulting in a rise of IDG internal temperature to 354 ±3"F (179 ±2"C)will cause the solder ring to melt. This liquefied solder flows into a cavity, where it remains. The IDG has an input shaft section, with a reduced diameter that will shear to protect the IDG when the input torque reaches 3,400 +250 pound-inches. After the shaft shears, there is no damage to the gearbox output spline and no contamination of the gearbox oil system. The gearbox can continue to operate. Auxiliary Power Unit (APU) - Mounted Generator: The APU-mounted generator provides fixed frequency by virtue of the APU operating at constant speed.If one or both of the engine generators fail, the load is transferred to the APU's generator (APU generator or auxiliary generator). The APU is primarily used during ground operation, but it is also rated for airborne operation at altitudes up to 45,000 ft. Note If one of the engine generators fails and the APU is unavailable, the load is picked up by the operating engine generator. The APU generator is controlled by the APU Generator Control Unit (APU GCU) and provides 40 kVA of 115/200V AC, 400 Hz, 3-phase power to the AC Tie Bus via the APU AC Contact.or ACC. The APU GCU provides primary control of the contactor and either prevents its closure if power quality from the APU generator is not within specified limits or opens the contactor if power quality at the APU generator Point of Regulation (POR) in the Left Power Distribution Box (LPDB) is not within specified limits. It also opens the contactor if power quality falls outside of specified limits. ON I OFF control and reset of the APU generator (and its output contactor) is provided by the APU GEN switch located in the Cockpit Overhead Panel (COP). A "Ready to Load" signal from the APU Ready to Load relay permits the GCU to energize the AAC once the APU has reached 99% of it's normal operating speed. Generator Control Units (GCUs): The GCUs, located in the Left and Right Electronic Equipment Racks (LEER and REER), provide control, monitoring, test and indication functions for the generators. The GCUs are electrically isolated from the generators in order to perform control and monitoring functions. The GCUs are initially powered by 28V DC from the Essential DC buses (L GCU and APU GCU from the Left Essential DC bus, R GCU from the Right Essential DC bus) until the generators reach normal operating speed.The GCUs then convert generated three phase AC power from a Permanent Magnet Generator (PMG) into 28V DC power with an internal diode transformer/rectifier. Each generator control unit (Left, Right, and APU) is identical in construction. The appropriate functioning for each generator is defined by a programmingjumper installed in the aircraft wiring harness.The GCUs are microprocessor-based and protect the AC buses by monitoring the quality of power produced by the associated generator. If generator performance does not meet specified parameters,the GCU will not permit power to be supplied to the AC buses. GCUs provide the following protection: Overvoltage: Overvoltage protection is provided by samplingthe three phasesatthe POR. If the highest phase voltage exceeds 125.5 +1V AC, the GCUwill open the generator contactor (LAC, RAC, and AAC) and de-energize the generator in accordance with an inverse timecurve. A UR/APU Generator Failadvisory message on the Crew Alerting System (CAS), change in the electrical power system synoptic page indicating affected generator, and the extinguishingof the "ON" indication inthe generator control switch (L GEN, R GEN, and APU GEN) will alertthe crew tothe loss of the affected generator channel. A reset of the generator can be attempted by cycling the generator control switch inthe COP. Undervoltage: Undervoltage protection operates similarty to overvoltage protection, with the exception of two distinct trip points/timedelays. The first trip level is set for a POR voltage of 102.5 +2.5V AC with a triptime of approximately 4.5 seconds. The second trip level is set for a POR voltage of 70 +1.5VAC with a trip time of approximately 160 milliseconds. A (advisory) and/or {advisory) CAS message, change in the electrical power systemsynoptic page indicating affected generator, and the extinguishing of the "ON" indication in the generator control switch (L GEN, R GEN, and APU GEN) will alert the crew to the loss of the affected generator channel.A reset of the generator can be attempted by cydingthe generator control switch inthe COP. Overfrequency: The overfrequency protection will sense the PMG (which hasa 3:1frequency ratiotothe output of the generator).There are three levels of overfrequency protection described as follows: If the generator's output frequency exceeds 420 +0,-5 Hz, the generator contactor (LAC, RAC, AAC) is opened in approximately four seconds and the generator is de-energized. If the generator's output frequency exceeds 440 +5 Hz, the generator contactor {LAC, RAC, AAC) is opened in approximately 160 milliseconds and the generator is de-energized. Additionally for the engine generators only, the Servo Valve {SRV) is de-energized. If the generator's output frequency exceeds 454 +8 Hz, the generator contactor {LAC, RAC, AAC) is opened in approximately 80 milliseconds and the generator is de-energized. Additionally for the engine generators only, the SRV is de-energized. Underfrequency: The generator is disconnected from buses and de-energized if the generator PMG frequency drops below 380 +5/-0 Hz. OEM Provided Data Electrlc:al Power Revision 13 20233-23 2A-24-00: B of 44 Overcurrent: The overcurrent protection will sense the three phases ofthe generator current transformers. If the highest phase current, as sensed by the GCU, exceeds 122 +/- 0.25 amps, the GCU will dear the fault in accordance with an inverse time curve. The GCU will first open the UR AC Cross Tie Contactor. If this dears the fault, no further action isrequired.If the overcurent condition isstillpresent for 200 milliseconds after opening the UR AC Cross Tie Contactor, the UR Essential AC Bus relay will belocked out. If the overcurrent is deared, no further action is required.If the condition persists for 200 milliseconds after UR Essential AC Contactor lockout, the GCU will open the left AC I right AC I APU alternator contactor and de-energize the generator output. Current Transformers (CTs) within the GCU monitor all three phases of generator output. If the current of any phase of generator power supplied to a bus exceeds 116-0/+5 amps, the GCU attempts current reduction by removing power - first from the main AC bus cross-tie relays, next by changing the power source of the essential bus, or finally by opening the contactor between the generator and the normally assigned main AC bus, thuslocking out the generator from all buses. An APU generator overcurrent will result in the generator first dropping power to the right main AC bus, then left main AC bus, before locking out the APU generator from all buses. (If another power source for the mainAC buses is operating, that sourcewill pick up and power the buses after APU generator power is dropped.) Phase Sequence: The GCU checks the three power phases at the POR during power up to vertfy they are in proper sequence. If the GCU detects that the phase sequence is not A-B-C, the GCU prevents the dosure of the APU alternator contactor or UR AC Contactor and the Generator Control Relay (internal to the GCU) will be tripped after a time delay of 60 milliseconds maximum. Differential Fault: If the CTs sense that current output of the generator differs from the current delivered to the assigned main AC bus by more than 20-25 amps, the generator is disconnected from the bus.If removinggenerator power to the bus clears the fault, a bus fault is indicated and the main AC bus is isolated from all power sources. If the differential fault does not clear, a generator malfunction is the source and the generator is blocked from powering the bus and bus power may be transferred to another AC source. Some generator protective features incorporate a time delay after the generator is removed from the aircraft AC buses. VVhen a GCU determines that the power output from the generator is again within acceptable parameters, the generator is made available to power the aircraft buses. Normal generator-to-bus configuration may be restored by cycling the RESET pushbutton on the overhead ELECTRICAL POWER CONTROL PANEL. If an aircraft engine is shut down for a malfunction or fire, the GCUs will de-energize the associated generator if a FIRE handle is activated or if a FUEL CONTROL switch is selected off. Bus Power Control Units (BPCU): Distribution of electrical power from the power sources are accomplished via a Left Power Distribution Box (LPDB) and a Right Power Distribution Box (RPDB), located in the bottom of the left and right electronic equipment racks. AC power from the L GEN, APU GEN and external AC power is routed to the LPDB located in the forward, left section of the aircraft beneath the LEER. AC power from the R GEN, APU GEN and external AC power is routed to the RPDB located in the forward, right section of the aircraft beneath the REER. The BPCUs dose and open contactors to p01Ner AC (and DC) buses in response to the COP pushbutton switch positions. BPCUs also illuminate the appropriate legend in the switch, indicating electrical system performance. VVith the right and left bus tie switches (L BUS TIE, R BUS TIE) in the normal (AUTO) position, the BPCUs provide automatic switching for the left and right main AC buses. In a standard aircraft engine starting sequence, the APU generator would first power both main AC buses. When the right engineis started and correct power is available from the right generator, the right BPCU switches the right main AC bus to right engine generator power, and the APU generator continues to power the left main AC bus. V\lhen the left engine is started and correct generator power is available, the left BPCU switches the left main AC bus to left engine generator p01Ner, andtheAPU generator revertsto a backup power source until the APU is shut down. When the BPCUstransfer mainAC buses to available power sources,the BPCUs first ensure that the power source parameters match in terms of frequency and phase angle. The BPCUs select one of the power sources as a reference for comparison, then signals the GCU to adjust IDG frequency and/or phase angle so that both power sources are congruent. This process is termed "paralleling" and enables the BPCU to achieve a No-Break Power Transfer (NBPT) that maintains power on the buses while switching p01Ner sources. All automatic p01Ner transfers will be NBPTs except those between an external AC power source and the APU generator, since the APU generator is not an IDG and the GCU cannot adjust frequency or phase angle without a hydro-mechanical drive unit. The BPCUs provide monitoringand bus protectionwhen external sources are providing electrical power to the aircraft. The left BPCU monitors external AC power; the right BPCU monitors external DC power. The BPCUs act in a manner similar to the GCUs, checking external AC for under/over voltage, under/over frequency, phase sequence and overcurrent. External DC is monitored for over/under voltage, overcurrent and polarity. If external power quality is within limits, the BPCU will signal the acceptable quality by illuminating the AVAI L section of the EXT PV\/R pushbutton on the COP. The AVAIL light will extinguish when external power is selected ON (external power quality can be monitored on the synoptic and system window displays).The BPCUs will trip the external power contactor if power quality falls outside of usable parameters. Electrical Power Distribution Buses: The Electrical Power System (EPS) is capable of providing electrical power throughout all operational conditions of the airaaft (ground and flight). The EPS generates 40 kVA of 3-phase 115/200V AC, 400 Hz electrical power from each of the engines on the aircraft (two mainengine and an Auxiliary Power Unit [APU]). The generated power is distributed from two Power Distribution Boxes (PDBs) to the aircraft's user equipment. These PDBs are provided for distribution and control of primary AC and DC power. A portion of the generated power is converted to 28V DC power via five TRUs capable of providing an output of 250 Amps each. A third PDB, located in the tail compartment, is provided for the distribution and protection of power to the Flight Controls Electrical Backup Hydraulic Actuators (EBHAs). The primary PDBs are rack-mounted units designed for easy removal and installation, allowing quick field servicing of the PDB contactors. The EBHA PDB is of a simplified design to allow quick field servicing. AC electrical power is distributed throughout the aircraft by the following buses: Left main AC Right main AC Emergency AC 60 Hz AC The left main bus is located in the Left Power Distribution Box (LPDB). The right main bus is located in the Right Power Distribution Box (RPDB). Both PDBs are installed in the bottom shelf of the respective Electronic Equipment Racks (e.g., LEER and REER). The Emergency AC box is located in the upper REER. The two 60 Hz power converters are located in the tailcompartment. The left main AC bus powers the Emergency AC bus during normal operation. The left and right main AC bus normally power the 60 Hz converter. Redundancy is provided by bus tie relays in the electrical system that allow any of the engine or APU-mounted generators to power the main AC buses. The Emergency AC bus may be powered by the RAT if the main AC buses cannot be powered by an aircraft generator. Left and right bus tie switches on the cockpit overhead panel can prevent a bus from receiving power from another generator source. Selecting the R BUS TIE or L BUS TIE switches to ISLN (isolation) allows only the respective right or left engine IDG to power the associated bus. 60 Hz Converters: Note The 60 Hz power system consists of one 3.5-KVA, 400 Hz to 60/50 Hz converter as active and one 3.5 kVA, 400 Hzto 60/50 Hz converter as standby. One 60 Hz external power receptacle and up to 24 internal power receptacles (and associated protection and distribution circuitry) are provided. Control and distribution of the 60 Hz power is provided by the solid state Secondary Power Distribution System (SPDS). During operation, when the aircraft's main power system is providing power, a 400 Hz to 60/50 Hz frequency converter, located in the tail compartment, takes 1151200V, 400 Hz, 3-phase power from the Right Main AC bus and converts it to 115/230V, single-phase AC power. The converter's input power is protected by a 20 Amp, 3-phase circuit breaker labeled "R FREQ CONV' located in the Right Main AC section of the Right Power Distribution Box (RPDB). The converter and associated 60/50 Hz circuit breakers are located in the baggage compartment EER. Note The aircraft is equipped with provisions to supply both AC and DC power to a galley installation(s) in accordance with the individual requirements of the customer. The specific arrangement of electrical supplies is contingent upon the customer-selected options determined duringfinal outfitting of the aircraft. The 60 Hz converters are controlled by the CABIN/GALLEY Master and the 60 Hz Master switches on the COP. The 60Hz Master is in series with the CABIN/GALLEY switch which allows the CABIN/GALLEY Master to act as a single-point load shed, while also allowing the 60 Hz Master to control the 60 Hz System without affecting the rest of the Cabin and Galley power. Both of these switches must be activated (OFF extinguished) before the 60 Hz converters will produce power. Once the cockpit overhead switches are activated, the virtual switches in the MCDU allow individual control of the Left and Right 60 Hz Converters. The MCDU switches would normally be left in the active (ON) position, but if the crew wishes to inhibit one (or both) of the converters from coming on, these switches allow individual control. 50 Hz Converter (Optional): Limited wiring provisions have been installed for a 50 Hz 250V AC converter. These provisions require removing the L 60 Hz converter and installing the 50 Hz converter in its tray. Since the 60 Hz Modular Power Tile (MPT) cannot support 250V AC, the wiring from the terminal strip in the forward Baggage Electronic Equipment Rack (BEER) to the 60 Hz MPT will require removal (capped and stowed). External AC Power: An external AC power source may be connected to the aircraft for ground servicing and/or maintenance. The AC receptade is located on the lower right side of the aircraft fuselage, forward of the right wing. When external AC power is plugged into the receptacle, the left BPCU first monitors the power source for quality parameters and, when requirements are met, illuminates the AVAIL legend on the EXT PVVR pushbutton on the ELECTRICAL POWER CONTROL panel. (External power parameters may be visually monitored on the AC Power 2/3 synoptic page or the AC I DC Power Summary 116 system page.) Selecting external AC power ON with the EXT PWR switch will power all buses of the aircraft electrical system. The left BPCU continues to monitor external AC for fault protection, and will drop external AC from the aircraft. electrical system if a malfunction occurs. Ram Air Turbine (RAT): If both engine generators fail and the APU is not available, the RAT is capable of supplying Emergency AC power by means of converting airstream energy to electrical energy. The assembly consists of a two-bladed 19 inch diameter turbine driving an air-cooled generator on a common shaft in a housing supported on a strut with a downlock. One 15 kVA, 115/200V AC, 400 Hz, 3-phase RAT is provided. The electrical generator for the RAT is a 6-pole, 3-phase, 115 / 200V AC, 400 Hz-nominal, 4-wire component driven at a speed range of 6,800 rpm to 8,800 rpm. The generator is air-cooled and rated at 15 kVA continuously when driven above 6,800 rpm. The RAT is able to provide power through its GCU to the Left Essential DC bus via the Left Essential TRU and the Right Essential DC bus via the Right Essential TRU during emergency conditions.The turbine is manually deployed from a recessed handle on the copilot side of the center pedestal (see Figure 3. Ram Air Turbine (RAT) Deployment Handle). The handle connects to an uplock: mechanism via a cable assembly. When the handle is pulled, the uplock: releases the spring-loaded, hydraulically-damped deployment actuator. This will force the RAT bay door open and place the RAT into the deployed position. The RAT system is capable of providing time-unlimited emergency electrical power in the event of a total loss of the other AC electrical power sources. Note The RAT is capable of operating at speeds of up to VMO I MMo at an altitude of up to 51,000 ft. The minimum airspeed for operation is approximately 180 KCAS. A RAT test can be initiated from a RAT TEST momentary cockpit switch. While the test is in progress, the SPDS provides a discrete output to illuminate the RAT TEST switch. Also while the self-test is in progress, the l;t!il§§1ijjl (caution) and (caution) CAS messages illuminate. The respective CAS message is extinguished if the RAT GCU I contactors or RAT heater passes their BIT test. j. Static Inverter: The static inverter, located in the REER, converts DC power from the L ESS DC Bus to a single phase A for the Emergency AC Bus to operate the Cabin Pressure Control System (CPCS). Controls and Indications: TheAC electrical system components are integrated intothe ModularAvionics Units (MAUs) through wire connectors to Input/Output (1/0) modules. The 1/0 modules translate analog system data into digital information for transmission on the ASCB-D bus. ASCB-D information is shared with the Monitor and Warning System (MWS) for CAS messages andaural annunciations, and sent to the Accelerated Graphics Modules (AGMs) for formatting prior to presentation on the Display Units (DUs) 2/3 synoptic pages and 1/6 window system pages. The synoptic and system window pages are valuable tools in circumstances requiring the use of electrical system abnormal or emergency procedures. Note For a detailed description of the AC Power 213 synoptic page and AC I DC Power Summary 116 system page, see Section 28-09-00. (See Figure 4. Cockpit Overhead Panel: Electrical Power Control Panel through Figure 6. Electrical Power Control Panel:TRU I MASTERS Section.) Circuit Breakers (CBs): The AC electrical power system is protected by the following CBs: Left Main AC Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) ACC 1 LPDB (LEFT MAINAC) ACC 2 LPDB (LEFT MAINAC) AUX TRU-LAC LPDB (LEFT MAINAC) EMG AC-LAC LPDB (LEFT MAINAC) L BATT CHARGER LPDB (LEFT MAINAC) L ESS TRU LPDB (LEFT MAINAC) L FREQ CONV LPDB (LEFT MAINAC) L MAIN TRU LPDB (LEFT MAINAC) PILOT 1 LPDB (LEFT MAINAC) PILOT 2 LPDB (LEFT MAINAC) PILOT 3 LPDB (LEFT MAINAC) PILOT 4 LPDB (LEFT MAINAC) PILOT S LPDB (LEFT MAINAC) R MN TRU-LAC LPDB (LEFT MAINAC) RBPCU-LAC LPDB (LEFT MAINAC) Right Main AC Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) ACC 1 RPDB (RIGHT MAINAC) ACC 2 LPDB (LEFT MAINAC) Circuit Breaker Name CB Panel Location (PDB Section) ACC 3 RPDB (RIGHT MAINAC) ACC 4 RPDB (RIGHT MAINAC) COPILOT 1 RPDB (RIGHT MAINAC) COPILOT 2 RPDB (RIGHT MAINAC) COPILOT 3 RPDB (RIGHT MAINAC) COPILOT 4 RPDB (RIGHT MAINAC) COPILOT 5 RPDB (RIGHT MAINAC) R ESS TRU RPDB (RIGHT MAINAC) R FREQ CONV RPDB (RIGHT MAINAC) R MAIN TRU RPDB (RIGHT MAINAC) STAB CH 2 RPDB (RIGHT MAINAC) R BATT CHRGR RPDB (RIGHT MAINAC) L MAINTRU-RAC RPDB (RIGHT MAINAC) AUX TRU-RAC RPDB (RIGHT MAINAC) RBPCU-RAC RPDB (RIGHT MAINAC) Crew Alerting System (CAS) Messages: CAS messages associated with the AC electrical pcmer system are: Main AC Electrical Power: Area Monitored CAS Message UR BPCU ll;fJd;J§@t (caution) UR BPCU L-R EPS NBPT Fai (caution) UR GCU / BPCU L-R Main AC Vol (caution) UR BPCU L-R AC Power Fai {caution) UR BPCU L-R AC Power Faul (caution) UR GCU / BPCU L-R Generator Overloa (caution) GCU / BPCU APU Generator Overloa (caution) GCU / BPCU APU Power Faul (caution) GCU MQllQ·@§ittJ (caution) GCU f;1ijllffilll ilj!I (advisory) GCU APU Generator Fail{advisory) GCU APU Generator Maint Req (advisory) UR BPCU L-R AC lsolatio (advisory) UR GCU l!;lijijllC (advisory) UR GCU L-R Generator Fai (advisory) Area Monitored CAS Message UR GCU L-R Generator (advisory) UR BPCU UR BPCU L-R EPS Maintenance Reqd (advisory) 60 Hz AC Power: Area Monitored CAS Message L or R BPCU L-R 60 Hz Power Fai (advisory) 50 Hz AC Power (Optional): Area Monitored CAS Message L or R BPCU 50 Hz Power Fai (advisory) NOTE: Generally installed only on aircraft registered outside of the U.S.A. External AC Power: Area Monitored CAS Message L BPCU l@j,j,.fi!f4dCl (caution) L BPCU External AC Power Faul (caution) L BPCU External AC Power Over1oa (caution) L BPCU External AC Availabl (advisory) External DC Power: Area Monitored CAS Message R BPCU External DC Availabl (advisory) R BPCU External DC Fai (caution) Limitations: Flight Manual Limitations: RAT: Operative Items: When operating on the RAT system, consult the list of operative items in the G650ER Airplane Flight Manual I Emergency Procedures I Electrical Power System I Dual Generator Failure. Use V\lhen Nonnal AC Power Is Available: Do not operate the Ram Air Turbine Generator when nonnal AC power is available, except as directed in abnonnal or emergency procedures. Figura 3. RamAir Turbine (RAT) Deployment Handle SEE DETAIL A DETAIL A TIL-006190 Figure 4. Cockpit Overhead Panel: Electrical Power Control Panel TIL-0000558 Figure 5. Electrical Power Control Panel: ELECTRICAL POWER CONTROL Section (1 of 3) UR BUS TIE AUTO: (Blue) UR Main AC Bus can receive power as a function of priority. SEE DETAILA ISLN: (Amber) UR Main AC Bus is isolated from lower priority AC power sources. UR GEN ON: (Green) Allows UR generator to provide power to aircraft. OFF: (Amber) Allows UR generator has been isolated. SWITCH NOT SELECTED: (Extended) UR generator not selected. No legends displayed. DETAIL A TIL--0000578 Figura 5. Electrical Power Control Panel:ELECTRICAL POWER CONTROL Section (2 of 3) SEE DETAIL A RESET AC/DC (Amber): Indicates a lockout due to protective trip. Cycling switch resets the lockout and allows the tripped contactors to dose if conditions are normal. APUGEN ON (Amber): Allows APU generator to provide power to aircraft. GND SVC BUS ON (Amber): The Ground Service Bus is powered by external power or the right main airaaft. battery. DETAIL A TIL-0000588 Figure 5. Electrical Power Control Panel: ELECTRICAL POWER CONTROL Section (3 of 3) SEE DETAIL A EXT PWR AVAIL: (Blue) AC or DC external power is available, but has not been selected ON. ON: (Amber) If AVAIL legend is illuminated, depressing switch causes the ON legend to illuminate, the AVAIL legend to extinguish and allows AC or DC external power aboard the aircraft. NOTE: If both AC and DC external power are available at the same time, the AVAIL legend will remain illuminated to indicate DC external poweris still available although AC external power has been selected. DETAIL A TIL-0000598 Figure 6. Electrical Power Control Panel:TRU I MASTERS Section SEE DETAIL A (@) IMASTERS 1 CABIN GAU£Y DETAIL A TRU LMAIN: OFF: Left Main TRU powered by left Main AC bus. R AC (Amber): Left Main TRU powered by right MainAC bus. R MAIN: OFF: Right Main TRU powered by right MainAC bus. LAC (Amber): Right Main TRU powered by leftt Main AC bus. MASTERS CABIN/GALLEY: When selected on, 115V AC power from the Left and Right Main AC busses and 28V DC power from the Left and Right Main DC busses supply power to CABIN/GALLEY systems. OFF (Blue): No power is available for CABIN/GALLEY systems. 60HZ: Wired in series with CABIN/GALLEY. Allows for control of the 60HZ system without affecting the rest of the Cabin and Galley power. TIL-002961 2A-24-30: DC Electrical Power System General Description: The Direct Current (DC) power system provides a means of generating, controlling, protecting and distributing 28V DC power for aircraft systems that require DC electrical power. DC power is supplied from one of the following: Transfonner-Rectifier Units (TRUs) Aircraft Batteries Ground Service Bus External DC Power Description of Subsystems, Units and Components: Transfonner-Rectifier Units (TRUs): Five identical and interchangeable TRUs provide primary DC power to the electrical power system. They are as follows: Left Main TRU Right Main TRU Left Essential TRU Right Essential TRU Auxiliary TRU The TRUs provide DC power to Left I Right (LIR) essential DC buses, L I R main DC buses, and the L I R auxiliary DC buses. The main and essential TRUs supply p01Ner to its associated DC buses. The auxiliary TRU provides power for the L IR auxiliary DC buses. The TRUs convert 3-phase, 115V AC, 400 Hz power to a 28V DC nominal output. The output voltage of the TRUs is unregulated and therefore dependent on loads. Each TRU operates in a nonnal range of 26V to 29V DC with loads in the range of 20 to 250 amps, respectively. Normal input power to the TRUs is supplied via the Integrated Drive Generators (IDGs), APU generator or external AC power through L I R main AC buses (dependent upon system configuration). Under normal configuration, the LmainAC bus powersthe L mainand L essential TRUs while the R main AC bus powers the R main and R essential TRUs. Inthe event of a main AC busfailure,the LIR mainTRUscan be powered by the opposite main AC bus through the LIR Main TRU AC contactor (L IR MTAC) During emergency operation, the left essential TRU can be powered by the Ram Air Turbine (RAl) through the Left Essential Transfonner Rectifier Unit Emergency AC Contactor (LETEAC). The auxiliary TRU receives its input power from the LIR main AC buses through Auxiliary TRU AC Contactor 1 (ATAC1). The power source priorities are the left main AC bus, foll01Ned by the right main AC bus. In the event of mainor essential TRU failure, the Aux TRU will pick up the affected bus. Once this happens, the Auxiliary (AUX) DC busses will be dropped (LADC and RADC will be de-energized). It should be noted that AUX TRU does not shed the AUX buses: it is the software in the Secondary Power Distribution System (SPDS) which de-energizes the two previously-mentioned contactors (based on cross-tie contactor status). The five TRUs supply power to DC buses through their associated contactors: left main DC contactor, Left Essential DC Contactor (LEDC), right main DC contactor, right essential DC contactor (REDC) and L I R auxiliary DC contactor (UR ADC). Each TRU powers its associated bus except for the auxiliary TRU, which powers the left and right auxiliary DC buses. The overtemperature warning discrete signal is output from each TRU for display as a (advisory), Forward Cabin Temp !rnttJI(advisory), o( (advisory) CAS message in the event of a sensed overtemperature condition or a detected fan failure. TRU Hall Effect Sensors (HES): The HES measure the intensity of the magnetic field produced by the flow of DC current thru the TRU system. This measured information is provided to the Bus Power Control Unit (BPCU) for EICAS and overcurrent protection. Aircraft Batteries: The G650ER aircraft utilizes the following batteries: Two 53 Amp/Hour Ni-Cad Main Ship Batteries One 53 Amp/Hour Ni-Cad Electrical Backup Hydraulic Actuation (EBHA) Battery One Flight Control Uninterruptable Power System (UPS) Battery (10.5Ah Sealed Lead Acid) Two Emergency Batteries These batteries come with an integrated battery charger.This equates to an overall weight savings of approximately 200 pounds. Main Batteries: Mainbattery power is provided by two 28V DC, 53 Amp/Hour NI-Cad batteries with an external charger. These batteries provide power for APU starting, running the AUX pump, powering Battery Bus loads and one or both Essential DC buses in the event p01Ner is not available from the Essential TRUs. Each of these batteries has an external charger capable of supplying 40 Amps at 28V DC for charging the batteries and supporting loads on their outputs. The batteries and battery chargers also provide indications to the Crew Advisory System (CAS) to alert the crew if there is a battery fault. During APU starting, the Left Main Battery is isolated from the Right Main Battery and all loads and is dedicated to starting the APU. This is done to ensure that sufficient voltage is maintained to all flight essential loads. Both Main Batteries also serve as the power source for the Auxiliary Hydraulic Pump. During flight, operation of the Auxiliary Hydraulic Pump by direct manual command of the crew is limited to two minutes before a reset is required. This is done to prevent inadvertent depletion of the batteries caused by the crew leaving the Auxiliary Hydraulic Pump running continuously. V\lhen the pump's system is in automatic mode and there has been loss of hydraulic pressure in a primary system, the Auxiliary Hydraulic Pump can start and stop when the flaps or landing gear is commanded to move and there is a disagreement between command and actual position. Automatic operation of the Auxiliary Hydraulic Pump is accomplished via the Secondary Power Distribution System (SPDS). Note Left and Right Main Battery current indications may read erratic values on overhead display (see Figure 8. Electrical Power Control Panel: MAIN BATTERIES Section) and DC Synoptic page during HF Transmission. Battery current indications will return to stabilized values when HF transmission is terminated. Electrical Backup Hydraulic Actuator (EBHA) Battery: A single 28V DC, 53 Amp/Hour Ni-Cad battery, designated as the "EBHA Battery", is provided to power seven FCS EBHAs. This battery has two external chargers in parallel, capable of supplying 80 Amps at 28V DCfor chargingthe battery and supporting loads on its output. The battery and battery charger also provide indications to the Crew Advisory System (CAS) to alert the crew if there is a battery fault. The EBHA battery is sized with sufficient reserve capacity and its charging controlled to ensure that sufficient excess battery capacity will exist to absorb all energy from all EBHAs under all foreseeable conditions. Flight Control Uninterruptable Power System (UPS) Battery: A single, 24V DC, 10.5 Amp-Hour Sealed Lead Acid (SLA) battery, designated as the "UPS Battery", provides redundant, uninterruptable power to flight control computers and servo valves. This battery incorporates an internal DC-DC charger for charging the battery and supporting loads on its output. An external TRU powered by the L MAIN or EMER AC Bus provides the 28V DC input to the UPS Battery. Integral charging, discharging, temperature and out gassing protection is also incorporated to ensure safe operation and performance of this battery. The battery and battery charger also provide indications to the Crew Advisory System (CAS) to alert the crew if there is a battery fault or if the battery does not have sufficient state of charge to dispatch. Emergency Batteries: Two 24V DC, 10.5 Amp-Hour Sealed Lead Acid batteries, designated as the "Emergency Batteries·(FWD and AFT), provides redundant, unintenuptable power to L and R Emergency Susses as well as to the Flight Instrument Bus. The Emergency Battery system is manually enabled (COMMAND ON) when: Either of the aircraft 28V DC bus feeds drop below 20V DCwhen the battery switch is ARMED The impact sensor input disaete is activated Manually selected on There are nine total power outputs provided by the Emergency Battery: Five emergency lighting outputs Two DC BUS outputs Two navigation system outputs The IRS outputs are not switched and are available whenever the battery is available. The emergency lighting outputs are segregated from the DC BUS and navigation system outputs. c. Ground Service Bus (GSB): (See Figure 7. Ground Service Bus Switch Locations.) The GSB is provided for nonnal aircraft servicing without activation of external power or operation of the APU or engines. Power for the GSB is supplied by the right battery. The bus can be activated through GND SVC BUS switches in the cockpit on the System Monitor Test Panel and also aircraft external locations (GND SVC Control Panel located in the tail compartment, Refueling Panel, Fwd Security I GND Service Switch Panel). When activated, the ground service bus provides po.ver to the following systems that may be individually turned ON or OFF: Aisle lights Airstair lights VVheel well lights Galley dome lights Vestibule dome lights Aft lavatory dome lights Baggage compartment lights Ramp area lights Utility compartment DC lights External water level indicator External service compartment lights Engine oiler Left and right fueling shutoff valves Left and right fuel quantity Water system DC VVindow Shades OEM Provided Data Elec:trtc:al Power Revision 13 2023-03-23 2A-24-00: 'll of 44 Switch Name Panel Location GND SVR BUS SW Ground Service Panel Aft Equipment Compartment (Tail) GND SVC BUS SWITCH Security I GND SVC Panel Left Side Of Nose (Exterior) GND SVC BUS SWITCH System Monitor and Test Panel To Left of Observer Audio Control Paneland the Right Electronic Equipment Rack (REER) External DC Power: External DC power provides 28V DC through an external DC power receptacle located just forward of the external AC power receptacle. This power source can power the main and essential DC buses through the External DC Contactor (EDC), the battery tie bus through the Auxiliary Power Contactor (APC) and the ground service bus through Ground Service Bus Contactors (GSBCs 1 and 2). When an external DC power cart is connected to the aircraft and turned on, the voltage and interlock are checked by the Right Bus Power Control Unit (RBPCU). If voltage and interlock are good,the RBPCU energizes the available relay and the AVAIL light will illuminate (assuming the battery switches are on).V\lhen the external power switch is depressed, an ON command is sent to the RBPCU. The RBPCU commands the EDC I ADC to close. With the EDC I ADC closed, power from the external source is supplied to the main and essential buses, and the external power ON light is illuminated. VVith external power applied, the RBPCU monitors voltage, polarity, current and interlock. The small pin on the external power connector provides interlock protection. This ensures that the plug is fully engaged before current flows through it. A BPCU lockout caused by any of the protective functions will result in illumination of the DC reset switch and opening of EDC I ADC. DC Interlock Protection: The DC external power receptacle has three pins: +28V DC/BPCU, +28V DC sense/interlock, and the +28V DC return. The +28V DC sense/interlock pin is shorter than the other two. The 28V DC signal on this pin is monitored by the RBPCU for power quality, in addition to supplying the BPCU power and interlock. This pin is shorter so connection tothe DCexternal power supply is established prior to the RBPCU sensing the 28V DC. The BPCU commands the EDC to open in 20 milliseconds if the interconnect is lost. DC External Power Protective Functions: The RBPCU provides overcurrent: over/under-voltage, reverse polarity, and DC interlock protection. The RBPCU is located in the REER. Polarity Protection: Vv'hen external DC power is first applied through the external DC power plug, the RBPCU verifies the power quality before it allows the EDC to close. One of the power quality checks is for correct polarity. The RBPCU is able to sense incorrect polarity for a voltage of more than 5 ±1V DC and respond within 60 milliseconds maximum. If incorrect. polarity is sensed, the RBPCU will inhibit EDC dosure, and a fault message will be generated as long as a source of power is available to the BPCU. Overvoltage Protection: The RBPCU overvoltage (OV) protection function monitors the external DC power voltage. If the voltage is more than 32.2 ±1V DC, the overvoltage protection function will trip the EDC after an inverse time delay. Undervoltage Protection: The RBPCU undervoltage (UV) protection function monitors the external DC power voltage. If the voltage is less than 21±1V DC, the undervoltage protection function will trip the EDC after 5.0 ±0.25 seconds time delay. Overcurrent Protection: A HES is located on the source side of the EDC to monitor external DC power. An overcurrent condition exists when the current is more than 320 ±20 amperes for an inverse time delay. The RBPCU overcurrent protective function sends a request to the Left Bus Power Control Unit (LBPCU) to open the Right Main DC Cross Tie Connector (RMDXC). If the fault is cleared, no further action is necessary. If the overcurrent fault is still present 100 milliseconds after the request was sent to the LBPCU, then the RBPCU opens Left Main DC Cross Tie Connector (LMDXC). If the fault is deared, no further act.ion is necessary. If the overcurrentfault is still present 100 milliseconds later, then the RBPCU sends a request to the LBPCU to open the REDXC. If the fault is cleared, no further action is necessary. If the overcurrent fault is still present 100 milliseconds after the request was sent to the LBPCU, then the RBPCU opens the LEDXC. If the fault is deared, no further action is necessary.If the overcurrent fault is still present 100 milliseconds later, then the RBPCU will open the EDC side of the EDC / ADC. Controls and Indications: (See Figure 8. Electrical Power Control Panel:MAIN BATTERIES Section and Figure 9. Electrical Power Control Panel: EMERGENCY POWER Section.) Note For a detailed description of the DC Power 2/3 synoptic page or AC/DC Power Summary 116 system page, see Section 26-07-00. Circuit Breakers (CBs): The DC electrical power system is protected by the following CBs: Left Main DC Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) PILOT 1 LPDB (LEFT MAIN DC) PILOT 2 LPDB (LEFT MAIN DC) PILOT 3 LPDB (LEFT MAIN DC) PILOT 4 LPDB (LEFT MAIN DC) PILOT 5 LPDB (LEFT MAIN DC) PILOT 6 LPDB (LEFT MAIN DC) PILOT 7 LPDB (LEFT MAIN DC) PILOT S LPDB (LEFT MAIN DC) PILOT 9 LPDB (LEFT MAIN DC) PILOT 10 LPDB (LEFT MAIN DC) PILOT 11 LPDB (LEFT MAIN DC) CAB 1 LPDB (LEFT MAIN DC) CAB 2 LPDB (LEFT MAIN DC) Right Main DC Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) COPILOT 1 RPDB (RIGHT MAIN DC) COPILOT 2 RPDB (RIGHT MAIN DC) COPILOT 3 RPDB (RIGHT MAIN DC) COPILOT 4 RPDB (RIGHT MAIN DC) COPILOT 5 RPDB (RIGHT MAIN DC) COPILOT 6 RPDB (RIGHT MAIN DC) COPILOT 7 RPDB (RIGHT MAIN DC) COPILOT S RPDB (RIGHT MAIN DC) COPILOT S RPDB (RIGHT MAIN DC) COPILOT 10 RPDB (RIGHT MAIN DC) COPILOT 11 RPDB (RIGHT MAIN DC) COPILOT 12 RPDB (RIGHT MAIN DC) GND SVC BUS RPDB (RIGHT MAIN DC) Left AUX DC Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) CABIN 2 LPDB (LEFT AUX DC) FWD LAV/GALLEY LPDB (LEFT AUX DC) Right AUX DC Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) CABIN 3 RPDB (RIGHT AUX DC) AFT LAV/GALLEY RPDB (RIGHT AUX DC) Left Essential DC Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) PILOT 1 LPDB (LEFT ESSENTIAL DC) PILOT 2 LPDB (LEFT ESSENTIAL DC) PILOT 3 LPDB (LEFT ESSENTIAL DC) PILOT 4 LPDB (LEFT ESSENTIAL DC) PILOT 5 LPDB (LEFT ESSENTIAL DC) PILOT 6 LPDB (LEFT ESSENTIAL DC) PILOT 7 LPDB (LEFT ESSENTIAL DC) PILOT S LPDB (LEFT ESSENTIAL DC) PILOT 9 LPDB (LEFT ESSENTIAL DC) PILOT 10 LPDB (LEFT ESSENTIAL DC) PILOT 11 LPDB (LEFT ESSENTIAL DC) LMAIN FUEL PUMP LPDB (LEFT ESSENTIAL DC) Right Essential DC Bus CBs Circuit Breaker Name CB Panel Location (PDB Section) COPILOT 1 RPDB (RIGHT ESSENTIAL DC) COPILOT 2 RPDB (RIGHT ESSENTIAL DC) COPILOT 3 RPDB (RIGHT ESSENTIAL DC) COPILOT 4 RPDB (RIGHT ESSENTIAL DC) COPILOT 5 RPDB (RIGHT ESSENTIAL DC) COPILOT S RPDB (RIGHT ESSENTIAL DC) COPILOT 7 RPDB (RIGHT ESSENTIAL DC) COPILOT S RPDB (RIGHT ESSENTIAL DC) COPILOT 9 RPDB (RIGHT ESSENTIAL DC) COPILOT 10 RPDB (RIGHT ESSENTIAL DC) R MAIN FUEL PUMP RPDB (RIGHT ESSENTIAL DC) (7) Battery Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) L BATT CTRL POP G-7 BATT BUS CONT CPOP F-7 L ESS DC CTRL 1 POP G-5 L ESS DC CTRL 2 CPOP G-5 R ESS DC CTRL 1 CPOP G-6 R ESS DC CTRL 2 POP G-6 R BATT CTRL CPOP G-7 (8) Ground Service Bus (External DC) CBs: Circuit Breaker Name CB Panel Location (PDB Section) GND SVC BUS - EXT DC RPDB {EXTERNAL DC) GND SVC BUS CTRL- EXT DC RPDB {EXTERNAL DC) (9) External Power CBs: Circuit Breaker Name CB Panel Location (PDB Section) EXT DC-L BPCU RPDB {EXTERNAL DC) EXT DC-R BPCU RPDB {EXTERNAL DC) EXT DC INTERLOCK SENSE RPDB {EXTERNAL DC) (10) EBHA DC Bus CBs: Circuit Breaker Name CB Panel Location (PDB Section) LAIL EBHA PWR FCSPDB {EBHA DC) RAILEBHA PWR FCSPDB {EBHA DC) L OB SPLR EBHA PWR FCSPDB {EBHA DC) R OB SPLR EBHA PWR FCSPDB {EBHA DC) L ELEV EBHA PWR FCSPDB {EBHA DC) R ELEV EBHA PWR FCSPDB {EBHA DC) RUD EBHA PWR FCSPDB {EBHA DC) b. Crew Alerting System (CAS) Messages: CAS messages associated with the DC power system are: Area Monitored CAS Message L-R Bus Power Control Unit @ifii;Jiipm (caution) Auxiliary TRU @h ii;Jii:M\ (caution) L-R Bus Power Control Unit Aux TRU Overloa (caution) Area Monitored CAS Message External Service Panels External Battery Switch 0 (caution) Right Bus Power Control Unit External DC Fai {caution) Right Bus Power Control Unit External DC Power Faul (caution) Right Bus Power Control Unit External DC Power Overloa (caution) Battery Contactor, L-R L-R Battery Contacto (caution) L-R Bus Power Control Unit l ijl•Sij§:J§l (caution) L-R Bus Power Control Unit L-R Essential DC-Batterie (caution) L-R Bus Power Control Unit L-R Essential DC-Contacto (caution) L-R Bus Power Control Unit L-R Essential DC-Vol (caution) L-R Bus Power Control Unit L-R Essential TRU Faul {caution) L-R Essential TRU L-R Essential TRU Ho (caution) L-R Bus Power Control Unit L-R Essential TRU Overloa (caution) L-R Bus Power Control Unit L-R Main DC Vol (caution) L-R Bus Power Control Unit L-R Main TRU Faul (caution) L-R Main TRU L-R Main TRU Ho (caution) L-R Bus Power Control Unit L-R Main TRU Over1oa (caution) Limitations: There are no limitations to the operation of the DC eledrical system as of this writing. Figure 7. Ground Service Bus Switch Locations FWD SECURITY I GND SERVICE S\\'ITCH PANEL SYSTEM MONITOR TEST PANEL GND SVC CONTROL PANEL REFUEL PANEL TIL-0069458 Figura 8. Electrical Power Control Panel: MAIN BATTERIES Sedion SEE DETAIL A VOLTS/AMPS Top rCN.J of digits displays battery voltage in amber from 0.0 to 50.0 Bottom rCN.J of digits displays battery amperage in amber from -400 to +400. MAIN BATTERIES LEFT: OFF: Left Main Battery is off. ON (Amber): Left Mein Battery is available for use as determined by system logic. RIGHT: OFF: Right Main Battery is off. ON (Amber): Right Main Battery is available for use as determined by system logic. DETAILA TIL-002978 Figure 9. Electrical Power Control Panel: EMERGENCY POVVER Sedion SEE DETAILA ARM: Arms the emergency lighting system. All legends in all switch capsules will be extinguished. OFF: ON: llluminalss the overwing and underwing egress lights. The "LIGHTS" LEGEND in Iha switch capsule will illuminate. Shuts off the emergency lighting system. The "LIGHTS" LEGEND in this switch capsule and the ARM switch capsule will illuminate. DETAIL A TIL-000127 2A-24-40: Secondary Power Distribution System (SPDS) General Description: (See Figure 10. Modular Power Tile (MPD Locations and Figure 11.Remote Interface Unit (RIU) Locations.) The Secondary Power Distribution System (SPDS) takes 28V DC and 115V AC power from the aircraft primary power system and distributes it to the aircraft loads. Each power output from the SPDS is routed using a Solid State Power Controller (SSPC). The SSPC performs a control and protection function. It allows the output power to be tumed on and off in response to an appropriate system command. In addition, the SSPC also protects against faults in the supplied load and its associated wiring, replacing multiple traditional electro-mechanical circuit breakers used on eartier model Gulfstream aircraft. The SPDS is composed of two dual-redundant Remote Interface Units (RIUs) and eight (four DC, four AC) Modular Power Tiles (MPTs) in a split-lane architecture.Each lane consists of one dual-redundant RIU and four MPTs. Control Function: Each SSPC output can be controlled 'ON' or 'OFF' by a control function using analog, discrete or communications bus inputs. This function may range in complexity from being permanently enabled (i.e. the SSPC is functioning only as a circuit breaker) to a closed-loop control.In the event of a failure of the control function, the SSPC can assume a pre-defined default state. The default state of the control function does not affect the state of the protection function. Protection Function: The SSPC output feeds to the aircraft loads incorporate protections against overcurrent, parallel arc fault (DC and 400 Hz AC loads only), series arc fault (DC and 400 Hz AC loads only) and Ground Fault Interrupt (GFI) (60 Hz AC loads only). Simli ar to a traditional circuit breaker, the protection function opens the SSPC and prevents its closure by the control function, essentially latching the SSPC in the open position. The state of the protection function is retained over power cycles. Reset of the SSPC protection function is through the MCDU. SPDS Power Architecture: The SPDS provides left and right system separation throughthe split-lane architecture.The left lane distributes