NATOPS Flashcards
Primary Fuel
A fuel that the aircraft is authorized to use for continuous unrestricted operations.
Restricted Fuel
A fuel that imposes operational restrictions on the aircraft.
Emergency Fuel
A fuel which may be used for a minimim time when no other primary or restricted fuel is available in case of emergency or operational necessity.
What are our primary fuels?
JP-5, JP-8, A++ (F-24), TS-1
What are our restricted fuels?
A1, A, JP-4, B
What are our emergency fuels?
JP8+100, F-27
What engine does the MH-60R have?
T700-GE-401C engine
5 sections: inlet, compressor, combustion, turbine, and exhaust.
Compressor Section
5 stage axial, single stage centrifugal
Where is TGT sensed?
Between the gas-generator and power turbine.
Engine Airflow Distribution
30% of total airflow used for the combustion process. The rest is utilized for:
- Compressor inlet temperature (T2) air
- Compressor discharge pressure (P3) air
- Combustor and turbine cooling
- Engine oil seal pressurization
What engine governing is retained with PCLs in LOCKOUT?
Np overspeed protection
TGT limiting, Np governing, and load shading are deactivated
Functions of the Engine-Driven Fuel Boost Pump
- Provide reliable suction feed from the aircraft fuel tank to the engine
- Provide discharge pressure to satisfy the minimum inlet pressure requirement of the HMU or high-pressure fuel pump.
HMU fuel is tapped off for what purposes?
- Positioning a metering valve to ensure proper fuel flow to the engine.
- Position a servo piston that actuate the variable geometry van servo and start bleed valve.
- Amplifying various signal (T2, P3, Ng) that influence fuel flow and variable geometry servo position.
HMU responds to the PCL for what?
- Fuel shutoff
- Setting engine start fuel flow with automatic acceleration to ground idle
- Setting permissible Ng up to maximum
- Fuel priming
- EDECU override capability (LOCKOUT)
Why is power available in FLY normally more than required?
- Fail-safe to high power (loss of torque motor electric signal)
- Power available with OEI (gas generator can increase power up to its limit
Functions of the HMU
RANNAF
- Rapid engine transient response through collective compensation
- Automatics fuel scheduling for engine star
- Ng overspeed protection (110 +- 2%, centrifugal valve secures fuel flow)
- Ng governing (T2, P3, Ng governing through 3D cam)
- Acceleration limiting (Ng governor protects PCL motion from damaging engine)
- Flameout and compressor stall protection (via VGV and AI/SBV position)
Functions of the ODV
- Provide main fuel flow to the 12 fuel injectors during engine start and operation
- Purge the main fuel manifold overboard, after engine shutdown, through a shutoff and drain valve to prevent coking of the fuel injectors
- Traps fuel upstream, which keeps the fuel/oil heat exchanger full, so that system priming is not required prior to the start
- Returns fuel back to the HMU if the Np overspeed is energized or if the EDECU hot start preventer is activated
Engine limiting definitions
Protect engine components and/or reaching a max possible outcome based on ambient conditions and collective setting
TGT-limiting is defined by reaching either IRP or CRP functions within the EDECU.
HMU-fuel flow limiting is defined by engine control system limiting the max power output under the following:
- Max fuel flow limited by the physical size of the fuel lines within the HMU and ODV
- HMU protecting compressor section by limiting fuel flow as a function of Ng and ambient temp
Increasing the collective during either of these will result in a droop in rotor speed below 100%, no increase in Ng or TGT, and a slight increase in torque within the range of continuous Np limits.
Engine limit condition definition
Torque limited condition is a transmission limit defined by reaching a Chapter 4 limit.
TGT-limited is an engine limit defined by reaching a Chapter 4 limit prior to reaching EDECU IRP, MRP, or CRP functions
Ng limited is an engine limit defined by reaching a Chapter 4 Ng limit.
If an engine is Ng limited and the collective is increased further, Nr will remain at 100% and Ng will increase (along with TGT and torque).
EDECU Operations
4N CHEF TASTEE
- Np governing - Actual Np compared to reference Np to compute speed error input signal
- Np overspeed protection - When Np>120%, ODV diverts fuel to inlet of HMU, causing flameout
- Np overspeed test - A/B buttons that re-reference Np to 96%
- Ng decay relight feature - turns on igniters for 5 seconds to attempt restart if Ng deceleration rate exceeded (>63%)
- Contingency power
- Manual: C-power switch on - TGT-limiting increased to 903, no further fuel flow increase at 891 +- 10
- Auto: enabled in OEI conditions (when one eng <50% torque); reset from 861 to 891 +- 10
- Dual-engine auto: requires Np <96%, greater than 3% droop between Np actual and reference, or 5%/sec Np droop rate - Hot start prevention - stops fuel flow when TGT >900 when Ng <60% and Np <50%; restored after TGT <300 or 25 sec
- Engine load sharing - torque matches lower torque engine by increasing power without affecting higher engine
- Fault diagnostic - displayed numerically on torque indicator (4 sec on/2 sec off); verify clear with TGT >425, PCL in IDLE/FLY
- TGT limiting - prevents further fuel flow to engine when TGT 866 +-10; Np/Nr will droop <100% and Ng governing will be sacrificed to protect against overtemp
- Auto-ignition system - turns on igniters for 5 sec when Np overspeed condition reached; will continue until Np/Nr controlled
- Cockpit Signals - provides Np, TGT, and torque to VIDS
- Transient droop improvement (TDI) - initiates power turbine acceleration why using anticipatory signals from TDI Nr sensor (good for autos)
- Eng speed trim - INCR/DECR switch, adjust Np between 96-101%
- EDECU lockout - PCL manually controls Ng and Np; deactivates TGT limiting, Np governing, and load sharing; keeps Np overspeed protection
CEDECU Operational Modes
incorporates a discrete application selection (DAS) plug, which mates with the E-4 connector
Navy code: 35 +-2.9%
Army Black Hawk code: 15 +-2.9%
EDECU: 0%
Ways to anti-ice the engine
- Vent bleed air from engine swirl vanes and engine IGVs by the engine AI/SBV
- Vent bleed air into the airframe engine inlet by the engine inlet anti-ice valve
- Continuously pump engine oil through scroll vanes
What indicates a malfunctioning anti-ice/start bleed valve?
- Appeareance or disapperance of the ENG ANTI-ICE ON advisory when outside of the range specified in the chart (<80.5: Off, 80.5 - 96.5: On, >96.5: Off)
- No illumination of ENG ANTI-ICE ON advisory when switch is selected on
- No rise in TGT when ENG ANTI-ICE turned on (30-100 increase)
Max torque available is reduced by up to 18% per engine with ENG ANTI-ICE on
How to tell in engine inlet anti-ice valve is not working properly?
Appearance of INLET ANTI-ICE On when OAT >13 (should be fully closed by 13, variable between 4 and 13, fully open <4)
The resultant loss of torque could be a maximum of 49%
Advisory comes on when inlet temp reaches 93
Engine Parameter Sensing
Np and Torque - two pairs of teeth induce electrical pulses that measure torsion or twist of Np shaft, which is proportional to output torque (left provides Np signal to EDECU and VIDS, right feeds torque compensation circuit and Np overspeed protection)
Ng - alternator provides signal to VIDS
TGT - thermocouple harness of 7 thermocouples, signal has -71 bias; located between gas generator and power turbine
What is the path of flight control inputs?
Cockpit inputs –> out, upwards, and aft –> pilot assist servos –> mixing unit –> primary servos –> bridge assembly –> swashplate assembly –> pitch change rods –> pitch change horns –> spindle assembly –> blade
When are anti-flap restraints and droop stops activated?
Anti-flap restraints:>35% Nr
Droop stops: out at >70% Nr, seat at <50% Nr on shutdown
Main Rotor Blade description
Each blade has a pressurized hollow spar, honeycomb core, outer skin, abrasion strips, electrothermal deicing mats, and a removable swept-back blade tip fairing. The 20° swept tips provide both sound attenuation and increased rotor blade efficiency. An electrothermal blanket is bonded into the leading edge for de-ice capability. The abrasion strips bonded to the leading edge of the spar extend the useful life of the blades. The spar of the main rotor blade is pressurized with nitrogen.
Tail Rotor System
A bearingless, crossbeam tail rotor blade system provides antitorque action and directional control.
The tail rotor head and blades are installed on the right side of the tail pylon, canted 20° upward, and provide 1 2.5 percent of the total lifting force in a hover.
With a loss of both tail rotor control cables, a spring-tension feature of the tail rotor control system will provide positive pitch on the tail rotor equivalent to the antitorque requirements (left pedal) for a midposition collective power setting.
The tail rotor indexing system positions the tail rotor blades during pylon fold operations and prevents the tail rotor from windmilling in winds up to 60 knots with the tail in the folded position.
Tail Rotor Quadrant
Transmits tail rotor cable movements into the tail rotor servo. Two spring cylinders are connected to the quadrant. In the event a cable is broken, the quadrant will operate normally by controlling the remaining cable against spring tension.
AVCS
Consists of an AVC Computer (AVCC), ten feedback accelerometers sensors, and a Vibration Control Actuation System (VCAS).
The AVCS implements a closed-loop feedback control algorithm utilizing accelerometers as the feedback sensors and airframe-mounted force generators as actuators.
8 vertical, 2 lateral accelerometers
(4 pair) 1000 lb force generators
(1 pair) 500 lb force generator
Functions of XMSN System
The primary function of the transmission system is to take the combined power from the two engines, reduce the rpm, and transfer it to the main and tail rotors. The secondary function is to provide a drive for electrical and hydraulic power generation.
Main Transmission System Components
The MGB drives and supports the main rotor. It is of modular design with a built-in 3° forward tilt. The main transmission consists of five modules: two accessory modules, two input modules, and a main module.
The input module provides the first gear reduction between the engine and the main module.
The engine output shaft provides drive from the engine to the input module via the diaphragm coupling. The diaphragm coupling is designed to allow for slight angular or axial misalignment of the engine output shaft during operation.
Each accessory module provides mounting and drive for an ac electrical generator and a hydraulic pump module.
Nr Sensor Locations
An Nr sensor for the vertical instruments is mounted on the right accessory module.
The Nr sensor for the TDI and the main transmission low-oil pressure sensors are mounted on the left accessory module.
Main Transmission Lubrication System
The main transmission incorporates an integral wet sump lubrication system that provides cooled, filtered oil to all main transmission bearings and gears). The ac generators on the accessory modules also use transmission oil for cooling.
The PDI is visible above the sonobuoy launcher. An extended PDI button is an indication that maintenance is required after the last flight of the day; it is not a lubrication system malfunction indication.
Transmission Lubrication Pressure and Temperature Sensors
MAIN XMSN PRESS LOW - activated when psi <14, located on No. ! accessory module (arthest point from the pumps)
VIDS pressure reading - taken at MGB manifold inlet
MAIN XMSN HOT - activated when temp > 117°C, located at the oil
cooler input to the MGB manifold
VIDS temperature reading - sensed in the sump using a temperature
sensor that is embedded in the main module sump chip detector
Transmission Lubrication System Cautions and Advisories
- MAIN XMSN OIL HOT caution (yellow) — Generated when the oil temperature sensor is activated.
- MAIN XMSN OIL HOT advisory (white) — Generated when the main transmission oil temperature is >105 °C and the MAIN XMSN OIL HOT caution is not activated. The main transmission oil temperature vertical instrument tape may be yellow or red, depending on oil temperature value:
a. Yellow tape — If the oil temperature is greater than or equal to 105 °C and less than or equal to 120 °C.
b. Red tape — If the oil temperature is greater than 120 °C. - MAIN XMSN PRESS LOW caution (yellow) — Generated when the oil pressure sensor is activated.
- MAIN XMSN PRESS LO advisory (white) — Generated when the main transmission oil pressure is less than or equal to 30 psi, the rotor speed (Nr) is greater than or equal to 25 percent, and the MAIN XMSN
PRESS LOW caution is not activated. The main transmission oil pressure vertical instrument tape may be yellow or red, depending on the oil pressure value:
a. Yellow tape — If the oil pressure is greater than or equal to 20 psi and less than or equal to 30 psi.
b. Red tape — If the oil pressure is less than or equal to 20 psi. - MAIN XMSN PRESS HI caution (yellow) — Generated when the main transmission oil pressure is >130 psi. The main transmission oil pressure vertical instrument tape will be red.
- MAIN XMSN PRESS HI advisory (white) — Generated when the main transmission oil pressure is between 65 -130 psi and the MAIN XMSN PRESS HI caution is not activated. The main transmission oil pressure
vertical instrument tape will be yellow.
Main/IGB/TBG Chip Detection System
The main transmission chip detector system consists of five chip detectors, each with a corresponding caution. The accessory chip detectors are located in the return lines of the No. 1 and No. 2 accessory modules. The chip detectors for the input modules and main module are located in the main module.
Each chip detector has a burnoff feature, which eliminates false warnings created by fuzz and minute particles. The fuzz burnoff feature is deactivated when oil temperature is above 140 °C; however, magnetic detection will remain. The chip detector for the main module sump rests in the lowest point of the oil system, contains an embedded temperature sensor, and incorporates a 30-second time delay to further eliminate false warnings.
The IGB and TGB contain identical chip detectors that contain an embedded oil temperature switch. When the oil temperature reaches 140 °C, the INT XMSN OIL HOT or TAIL XMSN OIL HOT caution will appear. The loss of oil will preclude proper operation of the oil temperature warning system.
CHIP IBIT Test
The CHIP IBIT takes approximately 2 minutes to verify circuitry and checks the individual chip detectors. For approximately the first 40 seconds of the test, 28 Vdc is interrupted and WCAs and MASTER CAUTION lights will not correctly respond.
Fuel System Summary
The fuel system is a crashworthy, suction-type system with a self-sealing main tank. The system is capable of pressure refueling, gravity refueling, and Helicopter In-Flight Refueling (HIFR). It has provisions for priming the engines, dumping fuel, indicating fuel quantity, and warning of low fuel levels.
Main Fuel Cell
The main tank system is composed of two cells interconnected to form one tank. The system capacity is 590 usable gallons (4,012 pounds when fueled with JP-5). The lower one third of the tank is self-sealing. The main cells’ interconnect level is approximately 270 to 600 pounds per side.
Fuel Transfer Fault/Fails
When a single valve or pump fails, the PUMP/VALVE FAIL caution will illuminate and the alternate pump or valve will supply fuel transfer capabilities.
If the AUTO FUEL XFER FAULT caution illuminates, this indicates a lack of fuel transfer from the auxiliary tanks to the main tanks. A lack of fuel flow may be caused by a failed closed transfer valve, both transfer/dump pumps inoperative, FMCP control failure, or a blockage in the associated fuel lines. Fuel in the auxiliary tank may be unavailable with an AUTO FUEL XFER FAULT.
Fuel Transfer Rules
Do not initiate unmonitored manual transfer to the main tank from auxiliary tanks until main tank is below 3,200 pounds. During manual auxiliary tank transfer, the main tank high-level sensor (float valves) should prevent overflow of the main fuel tanks.
When main fuel tank capacity has decreased approximately 300 pounds, check the manual fuel transfer system to ensure proper transfer.
At AUTO, the fuel management logic is not initiated until the main fuel tank fuel level depletes to approximately 2,640 pounds.
External Tank Jettison
Emergency jettison of external fuel tanks via the ALL STORE JETT pushbutton (PB) is inhibited when less than 40 gallons (approximately 272 lbs.) remain in the tank. SEL JETT must be used to jettison when less than 40 gallons (approximately 272 lbs.) remain in the tank. The fuel gauge for the external fuel tank reads in 50 lbs. increments. If the external fuel tank requires jettison and the fuel gauge reads 300 lbs. or less, SEL JETT should be used.
Fuel Quantity System
The fuel quantity indicating system incorporates a fuel probe mounted in each fuel cell. The tank probes are capacitance-type sensors that employ fuel as a dielectric to measure the weight of the fuel in each tank.
When a main tank falls below the Low Fuel Level Sensor caution level, the MAIN FUEL tape and readout turn yellow.
Fuel Low-Level Warning System
The fuel low-level warning system has two separate and independent indications to alert the aircrew of a low-fuel state. When the fuel level in one of the fuel cells reaches 300 pounds, the #1/#2 FUEL LOW caution appears and the associated digital fuel readouts on the mission and flight displays turn yellow. When total fuel reaches 600 pounds, the total fuel display also turns yellow.
In order to minimize aircrew desensitization when fuel washes on and off the sensors, the master caution will be displayed when a fuel low condition is detected for a period > 5 seconds and will be deactivated when the condition is not detected for a period of 20 seconds.
HIFR System
The HIFR system consists of a Wiggins quick-disconnect, pressure-refueling fitting, a pressure-refueling precheck switch to allow the high-level sensors to be checked from inside the aircraft, and a five-element (fuse) GO/NO GO canister.
Flow is reduced to an extremely low level if the fuel is contaminated with water and particulate matter. Once a 20-psi pressure differential exists, fuel flow stops.
APU
The APU is a gas turbine engine consisting of a power section, a reduction gearbox, appropriate controls, and accessories. Fuel consumption at rated power is 150 pounds per hour.
The minimum accumulator pressure required for starting the APU is approximately 2,650 psi. With ac power available, the accumulator is charged by the backup hydraulic pump.
APU Start System
With the FUEL PUMP switch in the APU BOOST position, pressurized fuel is supplied from the right fuel cell by the prime/boost pump. The fuel control governs and meters fuel flow to the APU power section, permitting automatic starting under all ambient conditions and constant speed operation once the APU has accelerated to its normal speed.
Placing the APU CTRL switch to ON initiates the start sequence. The DESU sends a signal to open the APU start valve, releasing the hydraulic accumulator charge to the starter. As the accumulator pressure drops below 2,650 psi, the APU ACCUM LOW advisory and ACCUM LO caution light appears, indicating that the accumulator pressure is low.
The APU ON advisory appears when the APU is on and operating normally.
Placing the APU GENERATOR switch to ON makes electrical power available. If the backup pump is cycled ON then to the OFF or AUTO position, it will remain on for one cycle of 90 seconds (180 seconds with winterization kit installed).
AC Electrical System
Primary ac electrical power is supplied by two oil-cooled 30/45-kVA, 115-volt ac, 3-phase, 400-Hz generators driven by the transmission accessory modules. Secondary ac electrical power is supplied by an air-cooled 35-kVA, 115-volt ac, 3-phase, 400-Hz generator driven by the APU.
Generator Control Unit
A Generator Control Unit (GCU) connects each respective generator to the ac electrical bus system, regulates generator output, and protects system components and circuitry against overvoltage, undervoltage, feeder fault, and underfrequency.
GCU Disconnection
- With WOW, if Nr droops below 94 percent, the GCU removes the main generator electrical input from the entire ac power distribution system.
- In flight, underfrequency protection is disabled and the generators will remain online until the Nr decreases to approximately 80 percent.
- A minimum of 97 percent Nr is required for the GCU to reconnect the generators to the ac electrical distribution system.
External Power Monitoring
External power is monitored for phase rotation, overvoltage, undervoltage, underfrequency, and overfrequency.
AC Bus Distribution
With both main generators operating normally, the No. 1 generator powers the No. 1 AC Primary, ac Essential, and AC Secondary buses; the No. 2 AC generator powers the No. 2 AC Primary and the ac Monitor bus.
If the APU generator is selected while both main generators are operating, the APU generator will not be connected to the ac bus distribution system.
Should either main generator fail, automatic bus switching limits the ac load to the available generator output.
AC Bus Load Priority
- The backup hydraulic pump (major load on the No. 1 AC Primary bus) will always be powered, if required.
- The mission avionics system is the major load on the AC Secondary bus and is the next priority. Tail rotor de-ice power is also supplied from this bus.
- The main rotor de-ice system is the only system powered from the AC Monitor bus and has the lowest priority of the major current-drawing components.
DC Electrical System
Two converters supply 28-volt dc power. The converters are powered by the No. 1 and No. 2 AC Primary buses, respectively. A 24-volt dc battery, located in the copilot seat well, is the primary source of power for APU starting and the secondary source of power for the dc essential bus and the Battery bus. With an 80 percent charge, normal battery life is 11 minutes for day operations and 9 minutes for night operations.
DC Bus Load
The No. 1 converter powers the No. 1 dc primary bus, the dc essential bus, and the battery bus. The No. 2 converter powers the No. 2 dc primary bus. The battery powers the Battery Utility bus.
If a single converter fails, the operating converter will pick up the load of the buses powered by the failed converter via automatic bus switching.
If both dc converters fail, the battery provides a source of emergency electrical power to the Battery Utility bus, the Battery bus (if the BATT switch is ON), and the dc Essential bus (if the battery is above a 35 percent charge). Power to the No. 1 and No. 2 dc Primary buses is dropped.
Battery Power Requirements
A 24-volt dc NiCad battery, located in the copilot seat well, is the primary source of power for APU starting and the secondary source of power for the dc essential bus and the Battery bus.
80% charge: 11 min / 9 min for day / night operations
40%: BATTERY LOW CHARGE caution
35%: DC Essential Bus dropped
30%: battery power may not be sufficient to activate the fire extinguisher (CAD).
Important Components of Each Bus:
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No. 1 Hydraulic System Summary
The No. 1 hydraulic system operates with rotors turning and supplies the first stage of the primary servos and the first stage of the tail rotor servo. System components include the No. 1 hydraulic pump, the No. 1 transfer module, the first stage of the primary servos, and the first stage of the tail rotor servo.
No. 1 Transfer Module
The No. 1 transfer module routes hydraulic fluid from the No. 1 hydraulic pump to the first stage of the primary servos and the first stage of the tail rotor servo. The No. 1 transfer module automatically routes hydraulic fluid from the backup pump if No. 1 hydraulic system pressure is lost.
The components of the No. 1 transfer module include a transfer/shuttle valve, pressure switch, first-stage primary servo shutoff valve, and first-stage tail rotor servo shutoff valve.
No. 2 Hydraulic System Summary
The No. 2 hydraulic system operates with rotors turning and supplies hydraulic pressure to the second stage of the primary servos and the pilot-assist servo assembly. System components include the No. 2 hydraulic pump, No. 2 transfer module, the second-stage primary servos, and the pilot-assist servo assembly.
Backup Hyd Pump
The backup hydraulic pump is identical to the No. 1 and No. 2 hydraulic pumps except that it is powered by an ac electric motor. An internal depressurizing valve reduces the output pressure of the backup hydraulic pump to aid startup of the electric motor.
When power is supplied to the pump, this valve is closed and 3,000-psi pressure is supplied to the hydraulic system. After 4 seconds on APU or external power, or 0.5 second with either main generator on, the pump will energize. The automatic low-level sensing switch, mounted on top of the pump, closes when fluid level is low, causing the BACK UP RSVR LOW caution to appear.
Automatic Backup Hyd Pump Initiation
- Loss of No. 1 hydraulic pump pressure (#1 HYD PUMP caution).
- Loss of No. 2 hydraulic pump pressure (#2 HYD PUMP caution).
- Loss of No. 1 hydraulic reservoir fluid (#1 RSVR LOW caution).
- Loss of pressure to the first stage of tail rotor servo (#1 TAIL RTR SERVO caution).
Utility Module
The utility module routes hydraulic fluid from the backup hydraulic pump to the No. 1 and No. 2 transfer modules, second stage of the tail rotor servo, rescue hoist, and APU accumulator. A pressure switch is located on the module sensing backup hydraulic pump output pressure and, if above a prescribed value, closes a circuit causing the BACKUP PUMP ON advisory to appear.
A priority valve is installed between the utility module and the rescue hoist to restrict hydraulic fluid to the rescue hoist in the event that backup hydraulic pressure decreases below a prescribed value. The utility module incorporates a velocity fuse, which secures fluid flow to the APU accumulator if flow rate exceeds a prescribed limit.
1 LDI Test
- # 1 RSVR Low Caution- Assume No.1 TR Servo; Backup Pump On; #2 TR Servo Openeda. #1 TR Servo Caution, Backup Pump On and #2 TR Servo On Advisory
2a. Leak Stops
a. Leak in #1 TR Servo: #1 Hyd supplies #1 servo, Backup supplies #2 TR servo
b. Cautions: #1 RSVR LOW, #1 TR SERVO; Advisories: BACKUP PUMP ON, #2 TR SERVO ON
2b. Leak Does not Stop
a. Complete Loss of #1 Rsvr Hyd Fluid: #1 HYD Caution -> flicker #1 PRI SERVO PRESS
b. Backup Hyd Pump supplies #1 Pri and TR Servo
Pilot Action
3a. Yes (Servo Switch 1st Off)
a. Loss of #1 PRI Servo; Backup Pump supplies #1 TR servo
b. Cautions: #1 HYD PUMP, #1 PRI SERVO PRESS, #1 RSVR LOW; Advisories: BACKUP PUMP ON
3b. No
Does Leak Stop?
4a. Yes (Leak upstream of #1 Transfer Module)
a. Backup Hyd supplies #1 Pri Servos, #1 TR Servos
c. Advisory: BACKUP PUMP ON
1 LDI Test
- # 1 RSVR Low Caution- #1 TR Servo off; Backup Pump On; #2 TR Servo Openeda. #1 TR Servo Caution, Backup Pump On and #2 TR Servo On Advisory
2a. Leak Stops
a. Leak in #1 TR Servo: #1 Hyd supplies #1 servo, Backup supplies #2 TR servo
b. Cautions: #1 RSVR LOW, #1 TR SERVO; Advisories: BACKUP PUMP ON, #2 TR SERVO ON
2b. Leak Does not Stop
a. Complete Loss of #1 Rsvr Hyd Fluid: #1 HYD Caution -> flicker #1 PRI SERVO PRESS
b. Backup Hyd Pump supplies #1 Pri and TR Servo
Pilot Action?
3a. Yes (Servo Switch 1st Off)
a. Loss of #1 PRI Servo; Backup Pump supplies #1 TR servo
b. Cautions: #1 HYD PUMP, #1 PRI SERVO PRESS, #1 RSVR LOW; Advisories: BACKUP PUMP ON
3b. No
Does Leak Stop?
4a. Yes (Leak upstream of #1 Transfer Module)
a. Backup Hyd supplies #1 Pri Servos, #1 TR Servos
c. Cautions: #1 HYD PUMP, #1 PRI SERVO PRESS, #1 RSVR LOW; Advisory: BACKUP PUMP ON
4b. No
a. Leak in Pri Servo 1st Stage; Partial Loss of Backup Hyd Rsvr fluid; Total loss of Backup Rsvr Hyd Fluid
b. Loss of #1 Pri Servos and #1 & #2 TR Servos
c. Cautions: #1 HYD PUMP, #1 PRI SERVO PRESS, #1 RSVR LOW, BACKUP RSVR LOW
2 LDI Test
- # 2 RSVR Low Caution- Pilot-Assist Servos Off;a. Boost servo Off; Cautions: SAS & AFCS DEGRADED
2a. Leak Stops
a. Leak in Pilot-Assist Servos: #2 Hyd supplies #2 servo, Pilot-Assist Servo function lost
b. Cautions: #2 RSVR LOW, BOOST SERVO OFF, SAS, AFCS DEGRADED
2b. Leak Does not Stop
a. Complete Loss of #2 Rsvr Hyd Fluid: #2 HYD Caution -> flicker #2 PRI SERVO PRESS
b. Backup Hyd Pump turned on, Pilot-Assist Servos turned on
Pilot Action
3a. Yes (Servo Switch 2nd Off)
a. Loss of #2 PRI Servo; Backup Pump supplies Pilot-Assist servos
b. Cautions: #2 HYD PUMP, #2 PRI SERVO PRESS, #2 RSVR LOW; Advisories: BACKUP PUMP ON
3b. No
Does Leak Stop?
4a. Yes (Leak upstream of #2 Transfer Module)
a. Backup Hyd supplies entire #2 Hyd system
c. Cautions: #2 RSVR LOW, #2 HYD PUMP; Advisory: BACKUP PUMP ON
4b. No
a. Leak in Pri Servo 2nd Stage; Partial Loss of Backup Hyd Rsvr fluid; Total loss of Backup Rsvr Hyd Fluid
b. Loss of #2 Pri Servos and Pilot-Assist Servos
c. Cautions: #2 HYD PUMP, #2 PRI SERVO PRESS, #2 RSVR LOW, BACKUP RSVR LOW, SAS, AFCS DEGRADED, BOOST SERVO OFF
Hydraulic Leak Test
Requirements to initiate:
1. Ac power.
2. BACKUP HYD PMP switch in the AUTO position.
3. All hydraulic reservoirs full.
4. Weight On Wheels.
5. Rotors engaged.
Satisfactory indications:
1. #1 RSVR LOW.
2. #2 RSVR LOW.
3. BACK UP RSVR LOW.
4. SAS.
5. BOOST SERVO OFF.
6. AFCS DEGRADED.
7. #1 TAIL RTR SERVO.
8. #2 TAIL RTR SERVO ON.
9. BACKUP PUMP ON.
10. MASTER CAUTION.
Flight Control - Mechanical Control System
- The cyclic, collective, and tail rotor pedal flight controls are routed aft and outboard of each pilot seat, vertically up each side of the aircraft, and are combined for each axis at the overhead torque shafts inside the hydraulics bay.
- The overhead torque shafts transfer inputs from the trim servos and flight controls through the pilot assist servos and the mixing unit.
- From the mixing unit, fore, aft, and lateral inputs are transferred to the swashplate assembly via the primary servos and the bridge assembly.
The yaw inputs to the tail rotor servo are transferred from the mixing unit aft to the tail rotor quadrant through the tail rotor cables.
Tail Rotor Control System
The tail rotor servo is mechanically actuated, but requires hydraulic pressure to operate the pitch change shaft, which moves the tail rotor pitch change beam, changing blade pitch angle through the pitch-change links.
The tail rotor servo is powered by either the No. 1 hydraulic system or the backup hydraulic system.
1. The tail rotor quadrant transmits tail rotor cable movement to the tail rotor servo.
2. Two spring cylinders connected to the quadrant allow cable tension to be maintained if either tail rotor cable becomes severed. Microswitches activate the TAIL ROTOR QUADRANT caution when either cable is broken.
3. Directional control of the tail rotor is maintained by the remaining spring. If both cables are severed, two separate centering springs will counter the tail rotor servo pilot valve positioning the tail rotor to a neutral setting to provide a fly-home capability.
Flight Control System Sections
- Mechanical control system.
- Flight control servo system.
- Automatic flight control system.
Primary Servos
There are three primary servos located in the hydraulics bay. Each primary servo has two stages that are independent and redundant with only the input linkage in common. Should one primary servo stage become inoperative due to pressure loss or a jammed input pilot valve, a bypass valve within the affected stage will automatically open, and the #1/#2 PRI SERVO PRESS caution will appear.
Pilot-Assist Servos
The pilot-assist servo assembly contains the boost servos, SAS actuators, and hydraulic (pitch and roll) trim actuators. Flight controls are operable without hydraulic pressure to the pilot-assist servos, but collective and yaw inputs will require considerable pilot effort. Hydraulic power is still required to move the primary servos.
Boost servos - collective, yaw, and pitch — located between the cockpit controls and the mixing unit that reduce cockpit control forces and SAS system feedback.
Control Mixing
Mechanical
1. Collective to yaw - Main rotor torque causes right nose yaw when collective is increased.
-Compensation: TR thrust is increased
2. Collective to lateral: Lateral lead (TR propeller effect) causes right drift when collective is increased
-Compensation: Rotor disk is tilted left
3. Collective to longitudinal: Rotor downwash on the stab causes nose pitch up and drifting aft when collective is increased
-Compensation: Rotor disk is tilted forward
4. Yaw to longitudinal: TR lift vector causes pitch down and forward drift when left pedal is applied
-Compensation: Rotor disk is tilted aft
Electronic
5. Collective/airspeed to yaw: Camber of tail rotor pylon varies side load with airspeed causes left nose yaw as airspeed increases
-Compensation: A portion of the main rotor torque compensation is provided by a trim input that is proportional to collective position and airspeed. The trim input is then progressively washed out as pylon side loads increase with airspeed.
AFCS Functions
19: 5 Hover, 5 Hold, PCSDAMBAT
- Pitch and roll attitude hold.
- Airspeed hold.
- Heading hold.
- Barometric altitude hold.
- Radar altitude hold.
- Pitch and roll hover augmentation/gust alleviation.
- Hover coupler.
- Crew hover.
- Cable angle hover.
- Automatic approach to hover.
- Pitch, roll, and yaw stability augmentation.
- Cyclic, collective, and pedal trim.
- Stabilator control.
- Diagnostics (failure advisory).
- Automatic depart.
- Maneuvering stability.
- Blade fold assist.
- Automatic preflight check.
- Turn coordination.
AFCS Functions
19: 5 Hover, 5 Hold, PCSDAMBAT
- Pitch and roll attitude hold - 5/6 °/sec <50 KIAS
- Airspeed hold - 6 KIAS/sec > 50 KIAS
- Heading hold - 3 °/sec < 50 KIAS, 1°/sec > 50 KIAS (within 2° of wings level, yaw <2°/sec)
- Barometric altitude hold
- Radar altitude hold - 0-5,000’ AGL (climb/descent 1000/200 fpm in hover) - SAS 2
- Pitch and roll hover augmentation/gust alleviation - SAS 2
- Hover coupler - within 1 KGS auto approach or 5 KGS manual; +-10 KGS 4-way. within 2 ft of radalt; max 116 torque
- Crew hover - +- 5KGS
- Cable angle hover - <5 KGS, with 10’ radalt, dome wet
- Automatic approach to hover - 2.5 kt/sec, 215 ft/sec >40 kt; 1.5 kt/s, 135 ft/sec <40; 360 ft/sec if above profile
- Pitch, roll, and yaw stability augmentation.
- Cyclic, collective, and pedal trim.
- Stabilator control.
- Diagnostics (failure advisories)
- Automatic depart - 120 KIAS/150’ AGL; 240 ft/min climb; 3 kt/sec <80 KIAS, 1 kt/sec >100 KIAS
- Maneuvering stability - >30° AOB: 1% fwd cyclic/1.5° AOB, 30-75° AOB
- Blade-fold assist.
- Automatic preflight check.
- Turn coordination - roll >1° and any of: lateral cyclic >3%, TRIM REL pressed, roll >2.5° using 4-way, HDG TRIM > 1sec
AFCC Control - Inner/Outer Loop
The AFCC employs two types of control, identified as inner-loop and outer-loop. The inner-loop (SAS) employs rate damping to improve dynamic helicopter stability. This system is fast in response, limited in authority, and operates without movement of the flight controls.
The outer-loop (autopilot) provides long-term inputs by trimming the flight controls to the position required to maintain the selected flight regime. It is capable of driving the flight controls through their full range of travel, or 100 percent authority, at a limited rate of 10 percent per second.
Stabilator Control
The stabilator is an automatic, fly-by-wire control system with a backup manual slew control. It is completely independent of the other two AFCS subsystems except for common airspeed sensors, lateral accelerometers, collective position sensor, and pitch rate gyros.
In low-speed flight, the purpose of the stabilator’s variable angle of incidence functionality is to eliminate undesirable nose-up attitudes caused by rotor downwash impinging on the stabilator. To accomplish this, the stabilator was designed to program so it aligns with rotor downwash in low-speed flight regimes.
Stabilator Inputs
- Collective position.
- Lateral acceleration.
- Airspeed.
- Pitch rate.
Stabilator Failure Travel Limits
Stabilator travel is restricted to 35° if an actuator fails in the full-down position or 30° if an actuator fails in the full-up position.
Stability Augmentation System
SAS 1 is an analog system; SAS 2 is a digital system that is part of the AFCC. SAS 1 and SAS 2 functions are identical except for the hover augmentation/gust alleviation and altitude hold/coupler features that are incorporated only into SAS 2. SAS 2 also complements the AFCC to provide turn coordination and roll attitude hold. With both SAS channels engaged, the pitch, roll, and yaw actuators have ±10 percent control authority with each channel providing ±5 percent.
Only SAS 2 commands the collective inner-loop actuator (CILA). The collective SAS operates with RAD ALT, BAR ALT, APPR/HVR, and DEPART modes engaged and is limited to ±10 percent control authority.
If either SAS channel malfunctions, the remaining operable SAS channel is limited to ±5 percent authority, but operates at twice its normal gain to partially compensate for the failed SAS channel.
Trim System
The trim system uses two high-torque electric servos for the yaw and collective axes and two hydraulic servos for the pitch and roll axes. The trim actuators command full control authority in all four control axes, but are rate-limited to 10 percent per second.
If cyclic trim is depressed and released while cyclic is in motion then trim is not engaged until cyclic motion stops.
Cable Angle Display
Outer Ring: 8.5°
Inner Ring: 4.25°
CABLE ANGLE AT LIMIT caution: 7.5°
Cable Angle Controls
Cable Angle: disengaged when TA raised through 27+-12 feet
Potentiometers: bias cable away from preset sensor calibration
-45° of rotation equates to 3° of cable angle bias to the AFCS; 90° of rotation equates to 10° of cable angle bias to the AFCS.
Requirements for Cable Angle to engage
Grounds speed +_ 5 kts, altitude hold within 10 feet of selected altitude, and a dome wet indication.
AFCS Preflight Check Requirements
- Weight on wheels signal.
- Rotor brake on.
- Engine torques below 10%.
- Both EGI attitudes valid.
- SAS 1 pushbutton engaged (after AFCC on for at least 20 seconds).
Microswitches required for a PYLON FLIGHT light
- Pylon lock pin switch
- 5 degree switch
- Tail rotor blade indexer switch
4/5. Stabilator lock pin switches (2)
Lights on the Master Warning Panel
1 ENG OUT
#2 ENG OUT
FIRE
LOW ROTOR RPM
MASTER CAUTION
Fire Detection System
5 detectors:
3 firewall-mounted (#1/#2 ENG, APU), 2 deck mounted (#1/#2)
When #1 ENG and APU T-handles have been pulled, the fire-extinguishing agent will be discharged into the compartment that whose T-handle was last actuated.
Also has a multiple-axis impact sensor (10g) that discharges one fire bottle into each engine compartment when a crash is detected.
APU T-Handle Activation Responses
- Removes electrical power from the APU airframe fuel shutoff valve, allowing it to close.
- Removes power from the prime boost pump.
- Sends a stop signal to the ESU/DESU.
- Arms the fire extinguisher system.
- Positions the extinguisher directional control valve to the APU compartment.
ECS Operation Torque / Fuel Reductions
NORM: max torque available is reduced by 4% per engine, and fuel flow to each engine will increase by 8 lb/hour/engine
HOT and OAT <15: max torque reduced by 5% per engine
HIGH: max torque available is reduced by 7% per engine and fuel flow increases by 45 lb/hr/engine
Using APU to power ECS: increase fuel consumption by 150 lb/hr
De-Ice Master Switch
When switched ON, will automatically turn on ENG ANTI-ICE, WINDSHIELD ANTI-ICE, and BLADE DEICE POWER on when ice is detected.
If turned off, or any of the switches are turned on by themselves, the automatic function is disable and the systems will operate continuously.
Blade De-ice System
Ice detector senses ice accumulation on a vibrating probe by measuring the change in probe frequency.
Sends ICE DETECTED caution and heats the probe to shed the ice and reset it for another cycle.
OAT governs heating EOT (element on time).
Rescue Hoist System
Speed:
-0-215 fpm (Breeze-Eastern), -0-250 fpm (Lucas Western)
-Cockpit Control: 100 fpm
-Backup: 50 fpm
200’ of cable, first and last 20’ painted bright orange
When raised/lowered at >50 fpm, it will automatically decelerate to 50 fpm approximately 10’ from full-up position or 5’ from full-down (not available in backup)
First sentence in NATOPS
The Naval Air Training and Operating Procedure Standardization (NATOPS) Program is a positive approach towards improving combat readiness and achieving a substantial reduction in the aircraft mishap rate.
Aircraft Overall Length
64’ 10”
Main Rotor Diameter
53’ 8”
Tail Rotor Diameter
11’ 0”
Seating Configuration and Restrictions
SUW: 3 seats, 2 pax
ASW: 2 seats, 1 pax
Log/VERTREP: <5 seats, <3 pax
SAR/MEDEVAC: <5 seats, <3 pax
Due to confined cabin space, dedicated SAR in an ASW configured aircraft equipped with ALFS is not recommended.
Up to 5 cabin seats may be installed for Log/VERTREP or SAR/MEDEVAC configuration. A max of 4 seats may be occupied in any configuration. Litter instillation prevents use of SO console seat and instructor seat.
Aircraft Height
16’ 10”
Tail Rotor clearance from ground
6’ 6”
MH-60R Primary Missions
-Surface Warfare (SUW)
-Anti-Submarine Warfare (ASW)
-Electronic Warfare (EW)
-Command and Control (CC)
Non-Combat Operations (NCO)
MH-60R Secondary Missions
-Amphibious Warfare (AMW)
-Air Warfare (AW)
-Health Services (HS)
-Fleet Support-Operations (FS)
-Intelligence Operations (INT)
-Logistics (LOG)
-Naval Special Warfare (NSW)
Exceeding Operating Limitations
Anytime an operational limit is exceeded, an appropriate entry shall be made on a MAF. The entry shall state which limits where exceeded and include range, duration, and any additional data that would aid maintenance personnel.
The flight display digital readout will display “XXX” in red when a TGT or torque limit is exceeded.
If a limit is exceeded to the extent that a red “XXX” is indicated, the crew shall land as soon as practicable.
Things to check in each location when preflighting
Corrosion, FOD, condition, and security
Preflighting the Hydraulic Handpump Service Reservoir
a. Fluid level
b. Hydraulic service valve selector pointed towards reservoir
c. Cap secure
d. Utility pump hydraulic service valve in vertical position
What to check when Preflighting the Main Rotor Head
Accumulator level/pressure, BIM indicators, blade, centering sockets, dampers, droop stops, elastomeric bearings, flap restraints, main rotor system, Pitch Control Rods (PCRs), scissor bearings
What to check in Engine when Preflighting
- Engine oil caps - Secure
- No.1 / No. 2 engines - oil level
- Engine oil/fuel filter PDIs - Flush
- Deswirl duct clamps - Secure
- Engine compartment
How many latches are on the top of the helicopter?
12
(2 - hyd bay, 4 - utility hyd pump, 2x - 1/engine, 2 - APU/ECS, 2 - fire bottles)
Tailwheel Preflight
a. Manual Unlock lever - Up position
b. Slip mark - aligned
Cabin Preflight Main Hits
- APU accumulator pressure - 2650 psi minimum
- Transmission PDI - flush
- Sono launcher - 1125 +-25 psi
Head Check
a. Blade lock pins engaged
b. Pitch lock pins retracted
c. Gust lock disengaged
Cable Angle Pilot Action Technique
- Use cyclic trim to command the aircraft in the corrective direction.
- Fly against lateral trim.
- Using trim release button, reposition aircraft.
- Fly against longitudinal trim.
With cable angle hover established, a pilot making a manual longitudinal correction against trim of 3% (0.3 inches) for 2 or more seconds will cause the AFCS to remove the cable angel hold solution in the longitudinal channel and requires up to 20 seconds to fully reincorporate the solution.
Hover Precautions w/ TGT
While operating in a salt spray environment for any period of time, a TGT rise of 20 degrees or more for a constant torque is an indication of engine performance degradation and possible salt encrustation.
A TGT rise of greater than 40 degrees for a constant torque is an indication of engine performance degradation that may result in compressor stalls.
Head Check
a. Blade lock pins engaged
b. Pitch lock pins retracted
c. Gust lock disengaged
Overwater SAR Precautions
- Static electricity from the helo must be grounded by the hoist prior to commencing pickup.
- Maintain a minimum of 1 rotor diameter from a parachute or canopy.
- Keep the hoist clear of all parts of the aircraft while operating. If contact or snagging occurs, suspend hoist operations and inspect cable.
- Simmer shall not be required to enter the water to recover an inanimate object.
- There shall be a hoist operator at the cabin if a swimmer is deployed.
- Personnel hoist shall not be attempted with a damaged hoist cable.
- Hoist operator shall wear a heavy-duty glove.
- Any time the cabin door is open during flight, all cabin occupants shall wear crewman safety harnesses or remained strapped in a seat.
- The Mk 25/58 produces smoke which is highly caustic.
- Mk 25 shall not be launched in a hove because of valve plug potentially striking aircraft or personnel.
- TGT concerns in a hover for salt encrustation/compressor stalls.
- Removal of Mk 58 pull ring shall not be removed until launching is to be accomplished.
- Recommend altitude for hovering is 70’. Prolonged low overwater hover with little or no headwind shall be avoided due to engine salt ingestion.
- If lost ICS occurs during SAR evolution, the co-pilot shall be notified and advisory hand signals/CREW HVR used.
- Use care to avoid cutting hand on edges of can after removing tear strip of Mk 58.
- Do not fly at low altitude of Mk 58. Second candle ignition can reach 50’.
- Hoist devices may contact forward part of starboard external fuel tank if installed.
- RAD ALT HOLD disengagement may be caused by swimmer/survivor oscillations.
- Keep survivor on right side of aircraft to allow air crewman to complete pickup.
SAR Rescue/Recovery Methods
- Landing to effect a rescue.
- Rescue via one or two wheels.
- Rescue via hoist.
- Rappelling.
- Direct deployment.
One or Two Wheel Landings
Shall not be conducted except for reasons of operational necessity.
VERTREP Power Margin Requirement
Power available is defined as a continuous transmission torque or TGT limiting with contingency power selected.
A minimum power margin of 6% shall be required prior to commencing each VERTREP evolution.
Ground Effect
A helicopter is said to be in “ground effect” when the rotor disk is within one rotor diameter of the ground.
The MH-60R is considered to be hovering in ground effect at radar altimeter altitudes at or below 45 feet.
Tail Rotor Characteristics
Tractor Tail Rotor
The tail rotor is designed to provide 2.5% of the total lift in hovering flight. To provide this lift, the tail rotor is canted 20 degrees from the vertical plane.
Loss of Tail Rotor Authority vs Loss of Tail Rotor Effectiveness
Loss of tail rotor authority is an issue of power. In extreme cases, the main rotor speed will droop. As Nr droops, torque increases while power available to the main and tail rotor decreases rapidly. Eventually, the tail rotor can no longer produce enough thrust to react against the high torque and the helicopter will spin to the right.
LTE is defined as the inability of the tail rotor to provide sufficient force to maintain yaw controllability. LTE occurs when full pedal input is insufficient to provide directional control.
LTE Regions
-Loss of Translational Lift: all azimuths
-AOA Reduction: 060 - 120
-Weather Vaning: 120 - 240
-Tail Rotor Vortex Ring State: 210 - 310
-Main Rotor Disk Vortex Interaction: 280 - 330
Recovery from LTE
- Altitude permitting, lower collective.
- Use forward cyclic to increase airspeed
- At very low speeds or in a hover, apply full left pedal.
Vortex Ring State
An aerodynamic condition where a helo may be in a vertical descent with maximum power applied and little or no cyclic authority. (settling into it’s own downwash)
Measurable at descent rates greater than 700 fpm and airspeeds between 0 - 20 KIAS, and is worst at descent rates of ~1500 fpm with airspeeds of 5 - 10 KIAS.
Vortex Ring State Recovery
- Decrease collective pitch.
- Increase forward airspeed.
- Enter autorotation if altitude permits.
Autorotation rates of descent
Min rate of descent: 75 KIAS
Max rate of descent: 95 KIAS
Practice autorotations are typically shot at 5 kts faster than optimum to allow the aircraft speed to slow towards an ideal condition instead of away from it.
Heavier aircrafts descend slower in a steady-state autorotation than lighter ones due to the greater rate of exchange of potential energy (altitude and weight) into kinetic energy (rpm) of the rotor system.
Static Rollover Angle
28 degrees
Critical Rollover
Critical rollover angle is the maximum lateral slope that can be negotiated in takeoff or landing. The this angle, full lateral cyclic input is required to trim the wheels level with the slope without sliding. With left wheel uphill and brakes on, the angle is ~12 degrees.
Turbulence Penetration Airspeeds
Light turbulence: blade stall speed minus 10 kts
Moderate turbulence: blade stall speed minus 15 kts
Instrument Takeoff Procedure
- Select hover mode on FD.
- Smoothly increase collelctive to takeoff power and maintain a hover attitude by referencing the AI. Allow AFCS to maintain heading (feet off pedal trim microswitches once airborne)
- Smoothly increase collective to climbout power. As the helicopter passes through 20’ on the RADALT, position cyclic forward to a 5 degree nosedown attitude and accelerate into forward climbing flight.
- As the helicopter accelerates, cross-check RADALT and VSI for positive rates of climb. Rate of climb should be 500 fpm or greater.
- Maintain a smooth acceleration up to 90 KIAS, referencing the AI and airspeed indicator.
External Power CB
The external power system include a 4-am circuit breaker that places a 2-amp draw on the power cable when connected to the aircraft. Some shipboard power systems require a load to be sensed in order to apply power to the cable.
Single Engine Condition Definition
The term “single-engine condition” is defined as a flight regime that permits sustained flight with One Engine Inoperative (OEI).
Stabilator Failure Airspeeds
0 degrees: 150 KIAS
10 degrees: 100 KIAS
20 degrees: 80 KIAS
30 degrees: 65 KIAS
40 degrees: 45 KIAS
Procedures not recommended in Stabilator Manual Mode
NFASS
-Night takeoffs, approaches, and landings (except one-time landing following failure)
-Flight into known IMC
-Automatic approaches to a hover
-Swimmer deployments lower than 15’ AGL
-Simulated emergencies, including practice autorotations
ADHEELS Activation
- Fresh or saltwater immersion.
- Impact force of 11 to 13 g’s or greater
- Attitude changes of 100 +- 5 degrees or greater in either pitch or roll.
When activated, remains on for a minimum of 45 minutes.
IHEELS Activation
- Fresh or saltwater immersion.
- Attitude changes of 100 +- 5 degrees or greater in either pitch or roll.
Types of Loss of Tail Rotor Control
- Tail rotor control cable failures. (25 - 145 KIAS, 19,500 lbs)
- Tail rotor servo failures. (40 - 120 KIAS)
- Restricted flight controls.
Reason why malfunctioning Anti-Ice Valve is important
The temporary hang-up of the engine Variable Geometry (VG) system at the anti-ice/start bleed valve may cause engine flameouts at low collective settings.
the HMU will restrict fuel flow to the engine to prevent hangup in the VG system, which may result in flameout during autorotations or quick stops.
Primary Servo Check Procedure
- SAS 1 and 2 - Off
- Check for servo interlock by swapping between 1st off at one seat and 2nd off at another, then repeating the process for the other switch and seat.
- turn 1st Off, then check pitch lock status by holding the collective fixed in the full up position and the tail rotor pedals neutral, the move cyclic slowly in a square pattern. Any restriction to the full range of cyclic motion may be indicative of an extended pitch lock.
- Move collective full up to full down in ~ 2 seconds and return to full down in ~2 seconds. (Possible ratcheting in the forward left corner - normal)
- Check for binds and restrictions while slowly moving flight controls through full range.
- Move collective from full down to full up in approximately 2 seconds and from full up to full down in approximately 2 seconds. Ensure no longitudinal or lateral cyclic control feedback is felt and the #2 PRI SERVO PRESS caution does not appear while the collective is in motion.
- Repeat steps for 2nd Off
AFCS SAS Ground Checks
SAS 1 and 2 - No movement should occur in main rotor blades. Minimal movement in flight controls is acceptable.
Stabilator Auto Mode Ground Check
- Indicator should read 34-42 degrees down.
- Press TEST button. Stab should move 5-12 degrees up. STAB caution and aural tone should appears. Press STAB AUTO CTRL PB.
- Hold STAB MAN SLEW up until Stabilator stops. Stabilator position should indicate 5-10 degrees up and travel should take 4-8 seconds.
- Slew back to 0 degrees.
- Reengage STAB AUTO MODE PB.
Stabilator Ground Check Limitations
Helicopter shall not be flown if the stabilator fails any check involving manual operation, positon indications, or warnings.
Night/IMC Descent Over Water Procedure
Required for all night/IMC descents over water at 1,000’ AGL and below.
Descent:
1. The PAC reports, “ON INSTRUMENTS” and states the leaving altitude, intended altitude, and variable RAWS/LAWS index position.
2. The PNAC acknowledges descent commencement, intended altitude, and RAWS/LAWS variable index position.
3. The aircrewman acknowledges the intended altitude. (During descent, the aircrewman should monitor the altitude via the NAV PARAMETERS table or the altitude display to the maximum extent permitted by the tactical situation.)
Level-off:
1. As the helicopter nears the intended altitude, the PNAC reports 200’ and 100’ prior.
2. When level, the PAC reports “LEVEL” and “ALTITDUE HOLD ENGAGED”.
Dual Concurrence Items
PCLs, T-Handles, Fuel Selectors, Generators, Computers, and EGIs
Comm Systems Backup and Failure Modes
- Single Mission Computer failure (AMC operational, ac power available) - No impact on ICS or radio communications.
- 1553 data bus failure (AMC operational, ac power available) - No impact to ICS. AMC continues normal operation with the exception that all control is maintained through the OCP and RCU. The radios function normally, but can only be selected via the OCP/RCU and tuned only via the RCU. CLEAR/SECURE comms are only available via RCU. On the OCP, CLEAR/SECURE, DF, and RELAY are not available.
- AMC or OCP failure (mission computer operational, ac power available) - RAD 1 is hardwired to the pilot station, and RAD 2 is hardwired to the copilot station (both controlled by RCU, PTT only). ICS CALL is available to all stations.
- Battery mode (electrical problem or prior to APU start, dc battery power available) - Only RAD 1 is available (hardwired to the pilot station and controlled by the RCU, PTT only). ICS PTT (fixed volume) is available to the pilots. ICS is available to the aft stations.
Checklist Methods
- Challenge-reply-reply
- Challenge-reply
- Starting-completing
CRM Skills
DAMCLAS
Decision Making
Assertiveness
Mission Analysis
Communication
Leadership
Adaptability/Flexibility
Situational Awareness
Which CRM skill is the most important?
Situational Awareness
HSI Wind Vector Colors
> 50 KIAS: Cyan
<50 KIAS: Amber and remains in last known calculated course and speed (“remembered wind”)
Navigation subsystem Components
- 2 EGIs
- 2 ADC
- Air data transducer
- Air speed transducer
- TACAN
- VOR/ILS
- DFG
- RAD ALT
- AFCS with AFCC
- DTIU
- GAS-1
- GPS/SATCOM
Navigation subsystem Components
- 2 EGIs
- 2 ADC
- Air data transducer
- Air speed transducer
- TACAN
- VOR/ILS
- DFG
- RAD ALT
- AFCS with AFCC
- DTIU
- GAS-1
- GPS/SATCOM
GAS-1 system
BLUF: Filters out erroneous signals and passes “good” signals to the EGIs. EGI 2 automatically connects to SATCOM/GPS in the event of a GAS-1 failure.
The GPS antenna system (GAS-1) is a seven-element anti-jam controlled reception pattern antenna (CRPA) and an antenna electronics (AE) unit . The AE unit nulls out suspected erroneous signals and passes “good” signals to the EGI navigation system. In case of an AE failure, a backup fixed reception pattern antenna (FRPA) will provide a GPS signal to EGI 2, utilizing a relay box called the data transfer interface unit (DTIU).
For those aircraft with GAS-1 installed, both EGIs are connected to GAS-1; EGI 2 automatically connects to SATCOM/GPS (FRPA) in the event of a GAS-1 failure. EGI 1 has no backup. SAASM EGIs can directly track the encrypted military P(Y) code without first acquiring the civilian C/A code (Direct-Y acquisition), significantly increasing resistance to jamming.
EGI 1 vs EGI 2
EGI 1: pitch/roll syncro data, attitude validity signal
EGI 2: Pitch/roll/heading syncro data, attitude/heading/velocity validity signals
LDS Malfunction Indications
PCLS in IDLE: Ng of malfunctioning engine 3-4% higher than other engine (Start Checks: IDLE Ng’s: Above 63, matched within 3)
When is a Head Check required?
Any time the BLADEFOLD switch is moved from the OFF position.
NATOPS is not what?
NATOPS is not a substitute for sound judgement.
Collective movement allowed to mitigate Droop Stop pounding
Startup: Raise collective not to exceed 0.5”
Shutdown: Raised a maximum of 1.5”
AFCS Auto Pilot Checks
(1) SAS 1, SAS 2, and TRIM pushbuttons — On.
(2) AUTO PLT pushbutton — On, then off. No movement should occur in flight controls.
(3) AUTO PLT pushbutton — On.
(4) Move flight controls slowly through full range of travel. Check for restrictions, control feedback, and rotor blade chatter.
Note: If any restricting control feedback or rotor blade chatter is detected, repeat this step with SAS/trim/autopilot channels individually disengaged to determine the channel and axis where the discrepancy exists.
3 required steps for all Sonar emergencies
The following should be performed for all sonar malfunctions, caution lights, or error codes.
- Alert crew.
- Execute Reeling Machine Malfunction emergency procedure.
- Complete the Sonar Troubleshooting checklist.
Soft Init vs Hard Init
Soft INIT: initializing a device by selecting the DIAG bezel key, scrolling to highlight the specified device and selecting INIT from the pop-up menu.
Hard INIT: initializing a device by pulling the circuit breaker for the specified device, waiting 20 seconds, resetting the circuit breaker, then performing a soft INIT.
Ground Proximity Warning System (GPWS) aka “Bitching Betty”
GPWS provides CFIT warning during all phases of flight. The type of warning provided by GPWS is dependent upon phase of flight, as determined by aircraft speed and altitude.
GPWS uses an authoritative female voice to issue the following warnings:
1. POWER, POWER
2. PULL UP, PULL UP
3. ROLL LEFT, ROLL LEFT
3. ROLL RIGHT, ROLL RIGHT
Np Overspeed Circuit Breakers
A popped NO. 1 ENG OVSP or NO. 2 ENG OVSP circuit breaker shall not be reset in flight. Resetting a popped NO. 1 or NO. 2 ENG OVSP circuit breaker may initiate an engine overspeed signal and result in engine failure.
C Power Types
- Manual - turning C PWR switch on. Allows TGT to reach 903°C and CRP limiter to re-reference to 891 °C ±10 °C.
- Auto - When the torque from one engine is below 50% (180 ft-lb), the opposite engine EDECU will automatically reset the TGT limiter to 891 °C ±10 °C. The #1 and #2 ENG CONTGCY PWR ON advisories will not illuminate to indicate that contingency power has been activated.
- Dual engine auto-contingency power - allows the EDECU to bypass the 10-min TGT limit and limit the aircraft at the contingency power TGT limit of 891 °C ±10 °C. For this feature to activate, one or more of the following conditions must exist:
a. Np drops below 96%.
b. Greater than 3% droop between reference power turbine speed (Np) and actual Np reference set point.
c. Greater than 5% per second Np droop rate exists with Np less than or equal to Np reference set point.
Wind Calculation Methods
- Steady State Estimate — Calculates the difference between ground track and air speed vectors during flight above 49 knots, with minimum acceleration. Expected average performance is within 10 knots and 20° from actual wind.
- Turn Estimate — Turns performed above 49 knots and 180° or above with relatively steady bank angle and air speed. Four 180° or greater turns refine wind calculation to a near Cycloid quality measurement.
- Cycloid — A 360° to 390° turn at a steady bank angle and airspeed. Dependent upon the helo start and end positions when flying a 360° turn. Expected average performance is within 2 knots and under 5° from actual wind.
Dipping Sonar Day/VMC Pattern Requirements for Pilot
- Fly directly to the FTP. Initiate or review Automatic Approach Checklist, as appropriate.
- If heading nearly into the wind, adjust heading at 0.6 nm (1,200 yards) toward gate position. For all other headings, offset heading a minimum of 30° toward downwind at 0.6 nm (1,200 yards). PAC may begin to
decelerate prior to reaching 1,200 yards as necessary. - At the initiation of manual approach report (“Commencing Manual Approach”), maneuver as required to arrive at the gate position.
- Maneuver as required to arrive at the gate position (aircraft into wind, wings level, FTP on nose, ~.3 nm [600 yards]) and fly the helicopter to a 70-foot hover.
- Signify intention to commence a manual approach to a sonar hover. - “AUTOMATIC APPROACH CHECKLIST”
- Announce initiation of the manual approach - “COMMENCING MANUAL APPROACH”
- Fly the helicopter to a 70-foot hover.
The PAC scan should be primarily outside. Do not decelerate below 50 KIAS until within 90° of the windline and do not descend below 90 ft until into the wind. Do not uncouple the rotorhead with large reductions in collective during the maneuver. - When the helicopter reaches the hover altitude/attitude with less than 5 KGS (RAD ALT hold ENGAGED), engage hover coupler. - “ENGAGE HOVER MODE”
- Establish a steady coupled hover (70-ft hover, less than 4 KGS), ensure cable angle is displayed in the cockpit, and initiate deployment of transducer. - “STEADY HOVER, DOWN DOME”
Cold Weather Preflight
- Check fuel drains for ice.
- Check engine inlets for ice or snow, specifically at the lowest point up to the swirl vanes.
- Check main rotor head and blades, tail rotor, and flight controls for ice and snow.
- Check the following vents/ports for ice blockage: Fuel tank vents, engine oil tank vents, transmission vents, battery vent, pitot-static tubes and ports.
- Check that tires are not frozen to the ground.
- Check landing gear struts and hydraulic accumulator for proper servicing.
- Apply preheat if available.
Cold Weather Engine Start considerations
When starting an engine that has been exposed to low temperature overnight, watch for a rise in TGT within 40 seconds. If no TGT rise is evident, prime the engine and attempt another engine start. If there is no overboard fuel flow during prime, inspect for ice in the sumps and filters. During cold-weather operation, allow a longer warmup period to bring transmission oil pressure to desired operating range. Monitor oil pressure and temperature closely.
PNAC Steps for every emergency
The PNAC shall:
1. Assist in ensuring the continued safe flight of the aircraft.
2. Perform the critical memory items that do not involve the flight controls.
3. Use the pocket checklist to complete non-critical memory items.
4. Troubleshoot as required.
What precipitates ground resonance?
Ground resonance may occur when landing with a malfunctioning damper system, or blade out of track condition.
Additional contributors to the development of ground resonance include landing with lateral drift, improperly serviced main landing gear, unique gross weight/CG and pilot-induced oscillations.
Ground resonance can also occur when a hard one-wheel landing cause an out-of-phase main rotor blades to be aggravated to the point where maximum lead and lag blade displacement is realized.
Retreating Blade Stall indications
-4x/rev vibrations
-Left roll
-Pitch up
-Loss of control authority
JP-4/B Fuel Restrictions
- All takeoffs shall stabilize in a hover with no fuel pressure cautions for a minimum of 10 seconds before commencing transition to forward flight.
- Single-engine training is prohibited.
- Operating characteristics may change. Lower operating temperatures, slower acceleration, and shorter range may be experienced.
- Due to the vapor qualities of mixed JP-4/JET B, the next two refuelings with a primary fuel shall be treated as if JP-4/JET B is in the tanks.
DC vs AC failure on Stab Indicator
DC failure: no flag will illuminate
AC: flag will illuminate
If you can’t slew, hold MAN SLEW up for 8 seconds to get to 0
Tail Gearbox Preflight
An over-serviced tail gearbox and/or a red tint to the tail gearbox oil are possible signs of contamination with hydraulic fluid. Failure of the tail gearbox is possible.
Night/IMC Decent Over Water N/W/C
Warning:
Failure to follow night/IMC descent procedure over water may lead to a loss of situational awareness and result in water impact.
Notes:
* Prior to commencing night/IMC descents over water, barometric altimeters should be synced to the radar altimeter.
* Descents should be commenced and conducted in a wings-level attitude when circumstances allow.
* Altitude hold shall be used in level flight at 500 feet AGL and below.
* RAWS/LAWS tones shall be verbally acknowledged by pilot and copilot.
Things that require operational necessity (NATOPS)
- Night VERTREP
- Night HIFR
- Rescue via One or Two Wheels
- Using Emergency Fuel
Oil Servicing Times
*Engines: 20 min
*APU: 1 hour
*XMSN: 30 min - 2 hours: HOT; >2 hours: COLD
How many blade lock pins are there?
8 (2x blade; only 1 visible from the top)
Blade Damper Accumulators Servicing
The blade damper accumulator is serviced with a combination of hydraulic fluid and nitrogen.
*The fluid-level gauge should be in the green or yellow (full) area.
*The nitrogen charge gauge should indicate the pressure for the ambient temperature (°F) listed on the accompanying chart. (min 1200 psi)
Dipping Sonar Day/VMC Manual Approach Profile Safety Requirements
The PAC scan should be primarily outside. Do not decelerate below 50 KIAS until within 90° of the windline and do not descend below 90 ft until into the wind. Do not uncouple the rotorhead with large reductions in collective during the maneuver
Backup Hyd Pump Switchology WOW/Airborne
With weight on wheels and BACKUP HYD PMP switch position:
*OFF — The backup pump remains off.
*ON — With the No. 1 and No. 2 hydraulic pumps operating normally, the backup pump recirculates hydraulic fluid to and from each transfer module and maintains pressure in the APU accumulator and rescue hoist systems, as required. The backup pump remains on until it is secured.
*AUTO — The backup pump automatically maintains hydraulic pressure to the No. 1 and/or No. 2 hydraulic systems, including the second stage of the tail rotor servo (as required for hydraulic system pressure and/or fluid losses detected by the LDI system).
With weight off wheels and BACKUP HYD PMP switch position:
*OFF or AUTO — The backup pump automatically maintains hydraulic pressure to the No. 1 and/or No. 2 hydraulic systems, including the second stage of the tail rotor servo (as required for hydraulic system pressure and/or fluid losses detected by the LDI system). The backup pump also maintains hydraulic pressure in the second stage of the tail rotor servo when the TAIL SERVO switch is placed to BKUP.
*ON — With the No. 1 and No. 2 hydraulic pumps operating normally, the backup pump recirculates hydraulic fluid to and from each transfer module and maintains pressure in the APU accumulator and rescue hoist systems, as required. The backup pump remains on until it is secured.
APU Accumulator Recharge
When the APU accumulator pressure is low and the backup pump is activated, the backup pump will run for at least 90 seconds (180 seconds with winterization kit installed), regardless of BACKUP HYD PMP switch position.
Ditching Preferences Over Water vs Over Land
For minimum loads on impact and to minimize the possibility of immediate rollover on touchdown, a ditching should be made into the prevailing winds and into or just past the crest on the backside of a wave. It is recommended that the aircraft ditch parallel to and near the crest of the swell, if there is a crosswind of 25 knots or less.
When does Ng begin to display?
6%
How long should you stay on the rotor break during Amber Deck?
Recommend no more than 5 mins (minimize Np shaft rub and asymetric heating)
Waterwash: 6 mins on rotor brake
What signals does the Alternator provide?
- Ng signal to VIDS
- EDECU
- AC power to ignitors
How does the Stabilator move?
Two electric jackscrew actuators, acting in series, position the stabilator. Each actuator provides one-half the input to position the stabilator and is controlled by a separate, redundant stabilator input.
Rolling Pull Outs
Power needs to be applied via transient power input (aft cyclic) and sustained power input (collective), which may result in a situation where power required for recovery greatly exceeds total power available in the rotor system and “mushing” can occur. During “mushing”, the aircraft will continue to descend rapidly even though the maximum power may be applied; longitudinal cyclic control will feel sluggish, a noticeable increase in main rotor vertical 4/rev vibrations and retreating blade stall may occur.
Figure of Merit (FOM)
FOM values range from 1-9. A normally functioning, aligned EGI will normally display FOMs of 2 or less and EHE of less than 25 feet.
Improperly aligned EGIs will show a large EHE (usually 32765) and a large FOM (usually 9).
What is the name of our radio?
VHF/UHF ARC-210
How does the aircraft perform when ditched into the water?
After a water landing, the aircraft tends to sink nose down and roll unpredictably to either side within 10 seconds.
The aircraft may maintain some degree of buoyancy in the fuel cell transition section (approximately 2 to 5 minutes) after water landing.
Nr decay during autorotations
10%/second
NATOPS Requirements to be a HAC
in addition to the requirements set forth for an MH-60R H2P, an MH-60R HAC shall:
1. Meet the requirements for HAC specified in CNAF M-3710.7.
2. Satisfactorily complete an approved MH-60R HAC PQS syllabus.
Type of ship approaches
- Visual
- Instrument
- ELVA
- Hung ordnance
Backup pump initiation on automatic approach / coupled hover
SAR: Backup pump turns ON to power the Rescue Hoist
Dipping: Switchology doesn’t matter
How many hours can you fly with a GEN BRG caution?
10 hours
How to disengage the Starter
If the starter fails to drop out automatically (52-65% Ng), it may be disengaged by pulling down on the PCL, pulling the circuit breaker, or removing the air source.
A malfunctioning starter may be overridden by manually holding in the starter button until Ng reaches 52-65% Ng.
Dynamic Rollover definition
When the angular velocity about the wheel is greater than can be countered with full opposite cyclic.
Autorotative regions
Prop region: 30% (outside)
Autot region: 45% (middle)
Stall region: 25% (inboard)
Tail Rotor Spar Loading
CCW turning single MR helos exhibit transient torque increases in forward flight with roll rates to the left. Left roll rates increases retreating blade AOA, driving torque up, and MR precession load contribute further to this effect. Left roll rates (Above ~30 degrees/sec in forward flight above 75 KIAS) can combine with induced tail rotor gyroscopic and flapping loads to cause excessive tail rotor spar loading.
When exceeding high roll rate maneuvers to the left, collective should be lowered concurrently with maneuver initiation to control transient torque increases and reduce high tail rotor spar loads.
Caution: Left roll rates in excess of 30 degrees/sec in forward flight above 75 KIAS may cause damage to the tail rotor spar.
What is lost without the DC Essential Bus?
Stabilator system Power, Cockpit floodlight, utility lights, and glareshield lights will be unavailable.
Following a loss of the MDs, the best indication of the loss of the DC Essential bus will be the MFR RCU display going blank.
Why is a clear area needed when turning on/off generators?
Warnings: Power transfer from the APU generator or external power to the No. 1 generator may cause disengagement of SAS1, SAS 2, TRIM, AUTO PLT, and stabilator. An unguarded cyclic may allow the rotor arc to dip as low as four feet above the ground.
Causes of an Engine Hi-Side Failure
High side failures may result from failures within the HMU or EDECU.
With HMU high-side failures, engine signals remain valid, making it easier to identify the malfunctioning engine.
EDECU high-side failures are often associated with erroneous engine signals and tend to follow one of two distinct failure modes characterized by either erratic torque or signal drop off (Np, torque, or TGT bias), making it more difficult to identify the malfunction.
Np Overspeed Test Fail
If the engine flames out/fails to relight:
1. Restart the engine
2. PCL (engine being checked) - FLY
3. PCL (engine being checked) - IDLE
4. Perform engine overspeed test again
If the engine relights/Ng returns to normal speed on the second attempt, the engine is acceptable. If the engine flames out/fails to relight again, MX action is required.
DLQ Currency
Aviation Ship/Amphibious Assault Aviation Ship
-Day: 2/365
-Night: 3/90
Air-Capable Ship (Clear Deck)
-Day: 4/180
Night: 6/90
Air-Capable Ship (Free Deck)
-Day: 4/180
-Night 6/90
Air-Capable Ship (RA)
-Day: 1/180
-Night: 2/180
Types of Transmission failures
Chip and lubrication malfunctions
Chip: decrease load on that module (PCL -> IDLE, secure GEN)
Lubrication: first sight may be low oil pressure
How long can you fly with a popped Main XMSN PDI?
18-20 hours
How to tell if an engine is still stuck in EDECU Lockout?
4 things to do as part of a general preflight
- In each location, inspect for corrosion, FOD, condition, and security.
- Check that all securing hardwire is safety-wired, cotter-keyed, and/or slipmarked.
- Plastic caps on avionics present a FOD hazard; check for security.
- If weapons are loaded, refer to technical manual for appropriate weapons preflight inspection.
How does the tailwheel lock
An electrical actuator positions the locking pin in the fork assembly. The pin is extended when in the LOCK position.
What are the two most likely reasons to shoot an actual autorotation?
- Dual engine failure (fuel related).
- Loss of tail rotor drive.
What is defined as an unusual attitude?
Unusual attitudes are considered to be pitch attitudes in excess of 30 degrees and/or roll attitudes in excess of 60 degrees.
Unusual attitudes are most likely to be encountered following an AFCS malfunction, AI degradation, vertigo, or poor instrument scan.
Nose attitudes to watch for in an autorotation
- Do not exceed 35 degrees nose up at any time during the flare or power-on recovery.
- Nose attitudes in excess of 13 degrees nose up at altitudes less than or equal to 15 feet will cause the tail bumper/stabilator to impact the ground.
When does the BATTERY LOW caution disappear?
- Battery is charged to >40%.
- Cautions/advisories disappear (DC essential bus dropped).
Torque considerations in Icing conditions
Ice accumulation resulting in a 20% torque increase indicates that normal autorotational rpm may not be attainable should dual-engine failure occur.
LDS Malfunction indications in an autorotation
Rapid Np/Nr rise. Engine with failed LDS may show a residual torque of ~12% with collective full down.
Engine Anti-Ice considerations in an autorotation
A malfunctioning anti-ice/start bleed valve may cause engine flameout during flight when the collective is full down, such as during quick stops and autorotative flight. No illumination of the ENG ANTI-ICE advisory light below ~88-92% Ng with ENG ANTI-ICE selected ON or OFF is an indication of a malfunctioning anti-ice/start bleed valve.
Warning: Engine flameouts during practice autorotations my result in loss of helicopter and personnel.
Oil System Servicing
Engine Oil: 7.3 quarts
APU: 2 quarts
Main XMSN: 7.5 gallons
IGB / TGB: 2.75 pints
Hyds: 1 quart
