(1) Reference: Military Specification MIL-E-5007E(AS), Engines, Aircraft, Turbojet and Turbofan, General Specification For, Table X, Typical Fuel Contaminants," dated September 1, 1983.
h. Section 33.69, Ignition System. The engine installation instructions should define the characteristics and interface requirements for all sources of electrical power required by the engine ignition system for reliable operation.
(1) Reference: Policy No. PS-ANE100-2001-1998-33.69-R1, "Policy for Evaluating Ignitions System Requirements, 14 CFR 33.69," dated February 26, 2001.
i. Section 33.84, Engine Overtorque Test. This section applies to turboshaft and turbopropeller type engines, which incorporate a free power turbine. This requirement allows the applicant to establish a maximum overtorque operating limit as follows:
(1) The maximum overtorque limit is approved for use for up to 20 seconds maximum duration, and for all conditions except operation at the 30-second/2-minute OEI rating (see reference below for limit definition).
(2) This limit addresses a main rotor droop condition primarily affecting turboshaft engines with free power turbines; however, certain turbo propeller installations may exhibit the same phenomena. An overtorque condition may also occur following an in-flight loss of power.
For example, an engine failure on a multi-engine rotorcraft might reduce available power by half, resulting in an instantaneous decrease in scheduled main rotor speed. As the main rotor control system corrects this droop by commanding a speed increase, and with additional collective normally applied by the pilot to maintain flight path, power demand at the main rotor increases sharply. It is due to this increased power demand that an overtorque condition may occur. Torque produced while the operating engine accelerates the main rotor back to its normal operating speed can cause high stress in the engine power transmission components.
(3) The power or torque-transmitting components in a free turbine engine are typically the turbine blades, disks, shafts, and gears (if configured with a gearbox). Torque has varying effects on the stress levels of these components. The following situations generally apply when analyzing these power transmission components:
(a) Turbine blades and disks are dominated by centrifugal loads, with overall strength being affected by core gas temperature. The torque produced by core gas loads has only a minor effect with regard to the total stress on turbines and blades.
(b) The stress levels of shafts and gears are typically dominated by the amount of torque they are transmitting. Shafts and gears are minimally affected by core gas temperatures.
(4) Typically, the time spent at maximum steady-state speed and temperature limits during the endurance test required by § 33.87, results in a more severe demonstration of stress and durability than would occur during a maximum overtorque event. This data can sometimes be 7
used to support compliance with § 33.84(b)(4). Therefore, the requirement to run the overtorque test at maximum steady-state temperature may be adjusted if you obtain FAA approval.
(5) Reference: 14 CFR part 1, § 1.1, General Definitions (Maximum engine overtorque).
j. Section 33.88, Engine Overtemperature Test. This test demonstrates a minimum overtemperature capability for the engine model.
(- This rule requires a test demonstration on a complete engine. The test vehicle and facility should be set up to minimize any mismatch of internal flow capacities of components and systems.
(2) For § 33.88(a) compliance, the maximum permissible rpm and indicated turbine gas temperature refer to the highest approved steady-state values, which are specified in the engine TCDS. This includes limits associated with all ratings, including OEl ratings of 2 ½-minute duration or greater.
(3) Section 33.88(b) contains additional test requirements for engines with a 30-second OEI and 2-minute OEl rating.
(4) The test overlimit values (e.g., 75 °F / 35 °F) must be applied at the turbine inlet station (typically called T4), not the type design operating limit measurement station (e.g., EGT or ITT). It remains, however, that the overtemperature test value is a function of the maximum steady-state operating temperature limit (at the type design measurement location). To determine the proper overtemperature test value, determine the turbine inlet (T4) temperature associated with the maximum steady-state limit value (at the type design measurement location). Add the overlimit value to this turbine inlet temperature (T4) to obtain the overtemperature test value for indicated turbine gas value.
(5) Post-test acceptance criteria for § 33.88(a) testing requires that the engine turbine assembly and support structure be within serviceable limits. Applicable components include, but are not limited to, blades, discs, drums, spacers, shafts, seals, stators, nozzles, and support structure. Serviceable limits are as specified in the ICA, as required by § 33.4. Post-test acceptance criteria for § 33.88(b) testing requires the turbine assembly to maintain its integrity at the end of the test. Maintaining integrity means that no imminent failure conditions are evident, and that the engine can continue to operate for continued safe flight and landing after the event.
(6) References:
(a) AC 33-3,Turbine and Compressor Rotors Type Certification Substantiation Procedures."
(b) Policy No. PS-ANE100-1988-00007, "Clarification Guidance on Implementation of FAR Section 33.88 Overtemperature Test," dated December 22, 1988.
(c) Policy No. PS-ANE100-1997-00004, "Engine Overtemperature Test; Governing 8
Temperature Location," dated October 7, 1997.
k. Section 33.89, Operation Test. The operations test must include all testing found necessary by the FAA to demonstrate safe operating characteristics throughout the specified operating envelope. Engine and component testing required for other part 33 sections may be used to show compliance to this section. The operations test should include tests to:
(1) Demonstrate start characteristics throughout the starting envelope.
(2) Demonstrate engine operation throughout the temperature and altitude envelopes.
(3) Demonstrate the ability of the engine to operate without damaging surge or stall (see § 33.65).
(4) Show adequate operational margins for variations in fuel schedule, variable geometry schedules, and installation/environmental effects.
(5) Demonstrate the ignition system design duty cycle.
(6) Demonstrate operation of the propeller parking brake (if equipped).
(7) References: None.
1. Section 33.95, Engine-Propeller Systems Tests. This test exercises the complete engine relative to propeller pitch and speed control, feathering, negative torque sensing, and decouplers.
(1) Compliance with this section is achieved by successfully completing the applicable tests and test cycles listed in § 33.95, paragraphs (a), (b), (c), and (d). Tests may be performed on an engine test stand, a test rig (such as a modified dynamometer rig or a propeller gearbox component test rig), or during flight test.
(2) Paragraphs (a) and (d) of this section address the feathering and reversing cycle tests.
They are typically conducted with a propeller on an engine test stand in conjunction with other tests, such as the engine endurance test. Alternately they may be conducted on an aircraft in conjunction with flight tests.
(3) Paragraphs (b) and (c) of this section address the negative torque and thrust tests. The decoupling tests required by paragraph (c) are typically run on an aircraft during flight or on a test rig. The reason for this is the engine system requires a negative torque or thrust-load input from the propeller. Negative torque is achieved in-flight with a windmilling propeller. On a test rig, a device drives the propeller shaft to achieve negative torque.
(4) When testing on a rig, the propeller should be represented in a manner that directly matches its polar moment of inertia. A flywheel or similar device is appropriate to this purpose.
Failure to properly match the design of the "representative propeller" to the propeller may cause inadvertent decoupling or torque sensing to occur during undesired operating conditions, such as during engine shutdown or other transients.
(5) When conducting any of these tests, specialized instrumentation will generally be needed to detect and record the operation of the systems being tested.
(6) Successful completion of the tests means that the engine-actuated propeller control functions as intended, without detrimental effect on either the engine or the propeller.
(7) References: None.
m. Section 33.96, Engine Tests In Auxiliary Power Unit (APU) Mode. This test demonstrates engine operability and durability when operating in APU mode.
(1) These tests assess steady-state and dynamic operations including:
(a) Hot section tolerance for locked rotor operations, including effects on combustor stability/coking, turbine nozzle integrity, turbine blade integrity, and vibration.
(b) Engine drive system stability/response, including the power turbine rotors, shafting, bearings and supports, the reduction and accessory gear trains, and rotor brake components.
(2) These propeller brake tests may be conducted in conjunction with the required tests of § 33.87, where applicable, and as approved by the FAA. The rotor brake system features, including operating instructions and limitations, should be described in the installation and operating manual as required under § 33.5.
(3) References: None.
n. Section 33.97, Thrust Reversers. This section requires demonstration of thrust reverser operation and durability. Section 33.97(a) states that the thrust reverser must be installed for the endurance, calibration, operation, and vibration tests required under subpart F of part 33.
(1) Reference: AC 20-18A, "Qualification Testing of Turbo-jet Engine Thrust Reversers."
Collen M D'Alessandee Colleen D'Alessandro Assistant Manager, Engine & Propeller Directorate Aircraft Certification Service 10