Source: http://www.tc.gc.ca/civilaviation/RegServ/Affairs/cars/Part5/Standards/533s.htm
Timestamp: 2013-12-07 17:10:51
Document Index: 144450108

Matched Legal Cases: ['art 33', 'art 33', 'art 33', 'art 33', 'art 33', 'art 91', 'art 31', 'art 34', 'art 34', 'art 33', 'art 33', 'art 33', 'art 33', 'art 33', 'art 33', 'art 33', 'art 21', 'art 11', 'art 13', 'art 16', 'art 22', 'art 23', 'art 25', 'art 27', 'art 29', 'art 31', 'art 33', 'art 35', 'art 37', 'art 41', 'art 51', 'art 91', 'art 93']

Part V - Airworthiness Manual Chapter 533 - Aircraft Engines - Part V - Airworthiness - Canadian Aviation Regulations (CARs) - Regulations - Aviation Safety - Air Transportation - Transport Canada
Part V Standards
Part V - Airworthiness Manual Chapter 533 - Aircraft Engines
Content last revised: 2010/01/29
[In these Standards:
[(a) The passages giving the Minister power to determine, approve, establish or authorise something without stating criteria for the use of such power are to be interpreted as requiring that the power be used in consideration of two factors only: the airworthiness of the aircraft that is the subject of the determination, approval or authorisation, or on which an aeronautical product that is the subject of the determination, approval or authorisation is to be installed, and the aircraft's level of safety;
[(b) the word "approved" or "authorised", when used without an indication of a method of approval or authorisation, is to be interpreted as referring to an approval or an authorisation granted under the Aeronautics Act.]
The content of this chapter is based on the United States Code of Federal Regulations, Title 14, Chapter I, Part 33 entitled Airworthiness Standards, Aircraft Engines. These United States airworthiness standards have been used and adapted as the model for the Canadian standards supplemented by additional airworthiness requirements based on Canadian experience and required for Canadian aviation purposes.
The FAR numbering system is used. The Canadian standards bear the same number as the FAR equivalent, prefixed by the number "5", as this chapter contains the standards for Part V of the Canadian Aviation Regulations (CARs).
First Edition Effective: January 1, 1986
The first edition of this chapter is based on FAR Part 33, up to and including amendment 33-10 published in the Federal Register dated February 23, 1984. Except for administrative changes (e.g., Administrator = Minister; Part = Chapter) and the deletion of references to operating FARs, there are no Canadian variations included in this first edition.
AMA 533.90 entitled "Initial Maintenance Inspection" dated Jan. 1, 1987 is attached to this chapter.
Change 533-1
This change incorporates Amendment 33-11 to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33 published in the Federal Register dated March 25, 1986.
- Amendment 33-11, "Turboprop Engine Propeller Brake", establishes a new standards applicable to turbopropeller engines equipped with a propeller brake. This amendment is needed to establish an appropriate level of safety for certification of aircraft engines with this feature.
In 533-1 changes were identified by marginal black lines. In the future, changes will be identified by "[ ]" brackets. Editorial alterations and typographical corrections will not be identified.
Change 533-2
This change incorporates amendment 33-12 to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33 published in the Federal Register on September 2, 1988. This amendment introduces the term "One-engine-inoperative (OEI)" rating, and its application, as used in "rated continuous OEI power", "rated 30-minute OEI power", and "rated 2 1/2 minute OEI power", together with additional requirements to be met in the endurance testing of rotorcraft engines.
Change 533-3
Effective: November 1, 1991
This change introduces an amendment to section 533.1 paragraph (b) to refer to:
- the Air Regulation enabling the type approval of aeronautical products and Chapter 511; and
- Chapter 516, Second Edition; Subchapter B.
In addition, the following amendments to the United States Code of Federal Regulations, Title 14, Chapter 1, Part 33 are included in the FAR text (left column) for completeness:
- Amendment 33-13, published in the Federal Register dated August 18, 1989, is part of a larger reorganisation of the general U.S. operating and flight rules. This amendment changes a cross reference to Part 91 in Appendix A of Part 31; therefore, it does not affect Canadian standards.
- Amendment 33-14 "Fuel Venting and Exhaust Emission Requirements for Turbine Engine Powered Aircraft", published in the Federal Register dated August 10, 1990. This amendment introduces the requirement that new turbine engines shall comply with the requirements of the new FAR Part 34. Part 34 recodifies the aircraft engine fuel venting and exhaust emission standards of Special Federal Aviation Regulation (SFAR) 27-5. Transport Canada has adopted the fuel venting and engine emission standards of ICAO, Annex 16, Volume II entitled "Aircraft Engine Emission", First Edition - 1981. Accordingly, section 533.1 is amended to refer to Chapter 516, Second Edition, Subchapter B.
Change 533-4
Effective: December 30, 1993
This change incorporates amendment 33-15 to the United States Code of Federal Regulations, Title 14 Chapter I, Part 33 in the Federal Register on May 18, 1993. This amendment establishes requirements for the approval of electric and electronic engine control (EEC) systems as presented in this FAA Final Rule. Although these types of control systems have been approved under existing regulations, they do not address specific requirements related to EEC.
Change 533-5
Published with Amendment 1999-4 of the CARs on 1 December 1999
This second edition introduces a new full page format and does not feature the left-hand column containing the FARs. Only the Canadian variations from the FARs are underlined with the FAR text following in a shaded box. The amendment number and date of affected pages has been removed from the bottom of the page. Instead, affected sections will be followed by amendment numbers and dates of current changes as well as any previous changes.
This change incorporates the following amendments to the United States Code of Federal Regulations, Title 14, Chapter 1, Part 33:
Amendment 33-16, which is for the revision of the U.S. authority citation, is not applicable in Canada and is not adopted.
Amendment 33-17
- This amendment entitled: "Continued Rotation and Rotor Locking Tests, and Vibration and Vibration Tests" published in the Federal Register dated June 4, 1996 revises the continued rotation and vibration certification standards for aircraft engines. This amendment is the result of an effort to harmonize the Federal Aviation Regulations with European Joint Airworthiness Authorities requirements. Furthermore, the increased uniformity of airworthiness requirements among the respective countries will simplify international airworthiness approval. Transport Canada shares this objective of International harmonization of airworthiness standards for the certification of aircraft engines. The adoption of this amendment has been subjected to consultation with Canadian aviation industry through NPA 96-07.
Amendment 33-18
- This amendment entitled: "Aircraft Engines New One-Engine-Inoperative (OEI) Ratings, Definitions and Type Certification Standards" published in the Federal Register dated June 19, 1996 establishes definitions and type certification of standards for new rotorcraft 30-second and 2-minute one-engine-inoperative (OEI) ratings. These new OEI ratings at higher power levels will enhance rotorcraft safety after an engine failure or precautionary shutdown. In addition, this amendment improves rotorcraft take-off and landing performances and allows for the installation of higher rated engines by rotorcraft manufacturers which will enable higher payload or shorter field take-off. The adoption of this amendment has been subjected to consultation with Canadian aviation industry through NPA 96-07.
Amendment 33-19
- This amendment entitled: "Airworthiness Standards; Rain and Hail Ingestion Standards" published in the Federal Register dated March 26, 1998 establish revisions to the Federal Aviation Administration's certification standards for rain and hail ingestion for aircraft turbine engines. These amendments address engine power-loss and instability phenomena attributed to operation in extreme rain or hail that are not adequately addressed by current requirements. These amendments also generally harmonise these standards with rain and hail ingestion standards being amended by the Joint Aviation Authorities (JAA). These amendments establish nearly uniform standards for engines certified in the United States under 14 CFR Part 33 and in the JAA countries under Joint Airworthiness Requirements-Engines (JAR-E), thereby simplifying the certification of engine designs by the FAA and the JAA. Transport Canada shares this objective of International harmonisation of airworthiness standards for the certification of aircraft engines. The adoption of this amendment has been subjected to consultation with Canadian aviation industry through NPA 98-159.
Change 533-6
This change introduces a new amendment format. This new amendment format is introduced in Chapter 533 of the Airworthiness Manual in order to be more consistent with the administrative procedures followed to amend the Canadian Aviation Regulations (CARs).
The following changes to the amendment procedures are introduced in this Change 533-6:
the preamble will be the focal point regarding the sections affected by this change. The change number will no longer be provided at the end of an amended section. Rather, for the current change only, an amendment tag identifying the coming into force date of the provision will follow the amended text. (example: (amended 2003/06/01)) brackets "[ ]" will no longer be used to identify new or revised text. In the paper version, new or revised text will be highlighted. In the electronic version, new or revised text will not be highlighted but followed by an electronic link to the previous version of the modified text. (example: (amended 2003/06/01; previous version)) the preamble will include a table of change information. This table will include the Notices of Proposed Amendments (NPAs) with the corresponding amended sections. 2. FAR Amendments
This change incorporates the following amendments to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33:
FAR Amendment 33-20
This amendment revises the bird ingestion type certification standards for aircraft turbine engines to better address the actual bird threat encountered in service. This amendment also establishes nearly uniform bird ingestion standards for aircraft turbine engines certified by the United States under FAA standards and by the Joint Aviation Authorities (JAA) countries under JAA standards, thereby simplifying airworthiness approval for import and export.
There are no proposed changes to the standard recommended by the CARAC Technical Committee Part V - Certification.
Change 533-7
Change 533-8
Correction to English version
This amendment entitled “Engine Ratings & Operating Limitations” corrects the English version of Section 533.7 of Chapter 533 of the Airworthiness Manual (AWM), which is currently missing subparagraph (c)(5)(v) as compared to the French version of the same section. There was no intention to create a Canadian variation as compared to the equivalent Federal Aviation Regulations Part 33, section 33.7 by omitting the missing subparagraph. Hence, the missing subparagraph is added to Chapter 533 of the AWM in order to remain harmonized with the subject section of FAR Part 33.
This change also incorporates the following amendment to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33: Correction to FAR Amendment 33-20
This amendment entitled “Bird Ingestion” adopts by reference FAA correction to FAR Amendment No. 33-20. FAR Amdt. No. 33-20 had originally been adopted with NPA 2000‑265 and published at Change 533-6 of Airworthiness Manual Chapter 533.
As published, the adopted standards contain errors that may prove to be misleading and need to be clarified. Corrections are provided for paragraphs (c)(5), (c)(7)(ii), (c)(7)(vii), (c)(7)(viii), (c)(7)(ix), (c)(8)(v), (c)(8)(vi), Table 1 and Table 2 of section 533.76.
Change 533-9
FAR Amendment 33-22
This amendment entitled “Aircraft Engine Standards for Engine Life-Limited Parts” establishes new and uniform standards for the design and testing of life-limited parts for aircraft engines. It retains the current lifing requirements and introduces damage tolerance requirements. In addition, new standards for the design of reciprocating engine turbocharger rotors are being added.
FAR Amendment 33-23
This amendment entitled “Engine Bird Ingestion” is amends the aircraft turbine engine type certification standards to better address the threat that flocking birds present to turbine engine aircraft.
FAR Amendment 33-24
This amendment entitled “Safety Analysis” amends the safety analysis type certification standard for turbine aircraft engines.
Change 533-10
On December 1, 2009, Part V Subpart 21 of the Canadian Aviation Regulations (CAR 521) came into force. CAR 521 replaces the following Regulations in Part V—Airworthiness: Subpart 11 - Approval of the Type Design of an Aeronautical Product
Subpart 13 - Approval of Modification and Repair Designs Subpart 16 - Aircraft Emissions Subpart 22 - Gliders and Powered Gliders Subpart 23 - Normal, Utility, Aerobatic and Commuter Category Aeroplanes Subpart 25 - Transport Category Aeroplanes Subpart 27 - Normal Category Rotorcraft Subpart 29 - Transport Category Rotorcraft Subpart 31 - Manned Free Balloons Subpart 33 - Aircraft Engines Subpart 35 - Aircraft Propellers Subpart 37 - Aircraft Appliances and Other Aeronautical Products Subpart 41 - Airships Subpart 51 - Aircraft Equipment Subpart 91 - Service Difficulty Reporting Subpart 93 - Airworthiness Directives In addition, with publication of CAR 521, the following Chapters of the Airworthiness Manual have been withdrawn: Chapter 511 - Approval of the Type Design of an Aeronautical Product Chapter 513 - Approval of Modification and Repair Designs Standard 591 - Service Difficulty Reporting Standard 593 - Airworthiness Directives This change amends section 533.1 to reflect changes in legal drafting style, in terminology and in references required because of the introduction of CAR 521. In addition, subsection 521.31(1) of the CARs is now used to legally enable this Chapter of the AWM. Change 533-11
FAR Amendment 33-26
This amendment entitled “Engine Control System Requirements” amends the airworthiness standards for aircraft engine control systems. These changes reflect current industry practices and harmonize TCCA standards with those of the Federal Aviation Administration (FAA) and with those recently adopted by the European Aviation Safety Agency (EASA). FAR Amendment 33-27
This amendment entitled “Aircraft Engine Standards for Pressurized Engine Static Parts” amends the airworthiness standards for aircraft engines by adding and making changes to standards for pressurized engine static parts. These standards, harmonized with the FAA, are equivalent to those already adopted by the EASA.
FAR Amendment 33-25
This amendment entitled “Rotorcraft Turbine Engines One-Engine-Inoperative (OEI) Ratings, Type Certification Standards” revises the airworthiness standards by revising the ratings' standards to reflect recent analyses of the ratings' use and lessons learned from completed engine certifications and service experience. This amendment harmonizes type certification standards for these ratings with the FAA and EASA.
FAR Amendment 33-28
This amendment entitled “Airworthiness Standards; Propellers” amends the airworthiness standards for aeroplane propellers. The previous propeller requirements did not adequately address the technological advances of the past twenty years. The new standards address these advances in technology and harmonize Transport Canada Civil Aviation (TCCA), FAA, and EASA propeller certification requirements, thereby simplifying airworthiness approvals for imports and exports. AIRWORTHINESS MANUAL CHAPTER 533 - AIRCRAFT ENGINES
533.1 Applicability
(a) This Chapter sets out airworthiness standards for the issue of type certificates and changes to type certificates, for aircraft engines. (amended 2009/12/01; previous version) (b) Reserved. (amended 2009/12/01; previous version) (Change 533-3 (91-11-01))
(Change 533-5)
533.3 General
Each applicant must show that the aircraft engine concerned meets the applicable requirements of this chapter.
533.4 Instructions for Continued Airworthiness
The applicant must prepare Instructions for Continued Airworthiness in accordance with Appendix A to this Chapter that are acceptable to the Minister. The instructions may be incomplete at type certification if a program exists to ensure their completion prior to delivery of the first aircraft with the engine installed, or upon issuance of a standard certificate of airworthiness for the aircraft with the engine installed, whichever occurs later.
533.5 Instruction Manual for Installing and Operating the Engine
Each applicant shall prepare and make available to the Minister prior to the issuance of the type approval, and to the owner at the time of delivery of the engine, approved instructions for installing and operating the engine. The instructions shall include at least the following:
(amended 2008/10/30; previous version)
(a) Installation instructions.
(1) The location of engine mounting attachments, the method of attaching the engine to the aircraft, and the maximum allowable load for the mounting attachments and related structure.
(2) The location and description of engine connections to be attached to accessories, pipes, wires, cables, ducts and cowling.
(4) A definition of the physical and functional interfaces with the aircraft and aircraft equipment, including the propeller when applicable. (amended 2010/01/29)
(6) A list of the instruments necessary for control of the engine, including the overall limits of accuracy and transient response required of such instruments for control of the operation of the engine must also be stated so that the suitability of the instruments as installed may be assessed.
(1) The operating limitations established by the Minister.
(2) The power or thrust ratings and procedures for correcting for non-standard atmosphere.
(3) The recommended procedures, under normal and extreme ambient conditions for:
(5) A description of the primary and all alternate modes and any back-up system, together with any associated limitations, of the engine control system and its interface with the aircraft systems, including the propeller when applicable.
(c) Safety analysis assumptions
The assumptions of the safety analysis as described in 533.75(d) with respect to the reliability of safety devices, instrumentation, early warning devices, maintenance checks, and similar equipment or procedures that are outside the control of the engine manufacturer.
533.7 Engine Ratings & Operating Limitations
(a) Engine ratings and operating limitations are established by the Minister and included in the engine type Certification data sheet, including ratings and limitations based on the operating conditions and information specified in this section, as applicable, and any other information found necessary for safe operation of the engine.
(ii) Rated take-off power (relating to unsupercharged operation or to operation in each supercharger mode as applicable).
(4) Temperature of the:
(iii) Turbo-supercharger turbine wheel inlet gas.
(5) Pressure of:
(1) Horsepower, torque, or thrust, r.p.m., gas temperature, and time for:
(iii) Rated take-off power or thrust (augmented);
(iv) Rated take-off power or thrust (unaugmented);
(v) Rated 30 minute OEI power;
(vi) Rated 2 1/2 minute OEI power;
(vii) Rated continuous OEI power;
[(viii) Rated 2-minute OEI power;
[(ix) Rated 30-second OEI power; and
[(x)] Auxiliary power unit (APU) mode of operation.
(5) Temperature of:
(amended 2003/12/11; no previous version)
(6) Pressure of:
(7) Accessory drive torque and overhang movement.
(16) For engines to be used in supersonic aircraft, engine rotor windmilling rotational r.p.m.
(d) In determining the engine performance and operating limitations, the overall limits of accuracy of the engine control system and of the necessary instrumentation as defined in 533.5(a)(6) must be taken into account.
(Change. 533-1 (87-01-01))
(Change 533-2 (89-01-01))
533.8 Selection of Engine Power and Thrust Ratings
SUBCHAPTER B DESIGN & CONSTRUCTION: GENERAL
533.11 Applicability
This subchapter prescribes the general design and construction requirements for reciprocating and turbine aircraft engines.
533.13 (Reserved)
533.14 (Removed)
533.15 Materials
The suitability and durability of materials used in the engine must:
533.17 Fire Prevention
(a) The design and construction of the engine and the materials used must minimise the probability of the occurrence and spread of fire. In addition, the design and construction of turbine engines must minimise the probability of the occurrence of an internal fire that could result in structural failure, overheating, or other hazardous conditions.
(b) Except as provided in paragraphs (c), (d) and (e) of this section, each external line, fitting, and other component, which contains or conveys flammable fluid must be fire resistant. Components must be shielded or located to safeguard against the ignition of leaking flammable fluid.
(c) Flammable fluid tanks and supports which are part of and attached to the engine must be fireproof or be enclosed by a fireproof shield unless damage by fire to any non-fireproof part will not cause leakage or spillage of flammable fluid. For a reciprocating engine having an integral oil sump of less than 25-quart capacity, the oil sump need not be fireproof nor be enclosed by a fireproof shield.
(d) For turbine engines type certificated for use in supersonic aircraft, each external component each external component which conveys or contains flammable fluid must be fireproof.
(e) Unwanted accumulation of flammable fluid and vapour must be prevented by draining and venting.
533.19 Durability
(a) Engine design and construction must minimise the development of an unsafe condition of the engine between overhaul periods. The design of the compressor and turbine rotor cases must provide for the containment of damage from rotor blade failure. Energy levels and trajectories of fragments resulting from rotor blade failure that lie outside the compressor and turbine rotor cases must be defined.
(b) Each component of the propeller blade pitch control system which is a part of the engine type design must meet the requirements of sections 535.21, 535.23, 535.42 and 535.43 of this Manual.
(amended 2010/01/29; previous version)
533.21 Engine Cooling
Engine design and construction must provide the necessary cooling under conditions in which the aeroplane is expected to operate.
533.23 Engine Mounting Attachments and Structure
(a) The maximum allowable limit and ultimate loads for engine mounting attachments and related structure must be specified.
(b) The engine mounting attachments and related engine structure must be able to withstand:
533.25 Accessory Attachments
533.27 Turbine, Compressor, Fan & Turbosupercharger Rotors
(a) Turbine, compressor, fan, and turbosupercharger rotors must have sufficient strength to withstand the test conditions specified in paragraph (c) of this section.
(b) The design and functioning of engine systems, instruments and other methods, not covered under 533.28 must give reasonable assurance that those engine operating limitations that affect turbine, compressor fan and turbosupercharger rotor structural integrity will not be exceeded in service.
(c) The most critically stressed rotor component (except blades) of each turbine, compressor, and fan, including integral drum rotors and centrifugal compressors in an engine or turbosupercharger, as determined by analysis or other acceptable means, must be tested for a period of 5 minutes:
(1) At its maximum operating temperature, except as provided in paragraph (c)(2)(iv) of this section; and
(2) At the highest speed of the following, as applicable:
(i) 120 percent of its maximum permissible r.p.m. if tested on a rig and equipped with blades or blade weights;
(ii) 115 percent of its maximum permissible r.p.m. if tested on an engine;
(iii) 115 percent of its maximum permissible r.p.m. if tested on turbosupercharger driven by a hot gas supply from a special burner rig.
(iv) 120 percent of the r.p.m. at which, while cold spinning, it is subject to operating stresses that are equivalent to those induced at the maximum operating temperature and maximum permissible r.p.m.
(v) 105 percent of the highest speed that would result from failure of the most critical component or system in a representative installation of the engine.
(vi) The highest speed that would result from the failure of any component or system in a representative installation of the engine, in combination with any failure of a component or system that would not normally be detected during a routine pre-flight check or during normal flight operation.
Following the test, each rotor must be within approved dimensional limits for an overspeed condition and may not be cracked.
533.28 Electrical and Electronic Engine Control Systems
Each control system which relies on electrical and electronic means for normal operation must:
(a) Applicability. These requirements are applicable to any system or device that is part of engine type design that controls, limits or monitors engine operation and is necessary for the continued airworthiness of the engine.
(1) Functional aspects. The applicant must substantiate by tests, analysis or a combination thereof, that the engine control system performs the intended functions in a manner which:
(ii) Complies with the operability requirements of 533.51, 533.65 and 533.73, as appropriate, under all likely system inputs and allowable engine power or thrust demands, unless it can be demonstrated that failure of the control function results in a non-dispatchable condition in the intended application;
(2) Environmental limits. The applicant must demonstrate, when complying with 533.53 or 533.91, that the engine control system functionality will not be adversely affected by declared environmental conditions, including electromagnetic interference (EMI), High Intensity Radiated Fields (HIRF) and lightning. The limits to which the system has been qualified must be documented in the engine installation instructions.
(c) Control transitions.
(1) The applicant must demonstrate that, when fault or failure results in a change from one control mode to another, from one channel to another, or from the primary system to the back-up system, the change occurs so that: (amended 2010/01/29; previous version)
(ii) The engine does not surge, stall or experience unacceptable thrust or power changes or oscillations or other unacceptable characteristics; and (amended 2010/01/29; previous version)
(iii) There is a means to alert the flight crew if the crew is required to initiate, respond to or be aware of the control mode change. The means to alert the crew must be described in the engine installation instructions and the crew action must be described in the engine operating instructions; (amended 2010/01/29; previous version)
(d) Engine control system failures. The applicant must design and construct the engine control system so that: (amended 2010/01/29)
(1) The rate for Loss of Thrust (or Power) Control (LOTC/LOPC) events, consistent with the safety objective associated with the intended application can be achieved; (amended 2010/01/29; previous version)
(2) In the full-up configuration, the system is single fault tolerant, as determined by the Minister, for electrical or electronic failures with respect to LOTC/LOPC events; (amended 2010/01/29; previous version)
(3) Single failures of engine control system components do not result in a hazardous engine effect; and (amended 2010/01/29; previous version)
(4) Foreseeable failures or malfunctions leading to local events in the intended aircraft installation, such as fire, overheat or failures leading to damage to engine control system components, do not result in a hazardous engine effect due to engine control system failures or malfunctions. (amended 2010/01/29; previous version)
(e) System safety assessment. When complying with this section and 533.75, the applicant must complete a System Safety Assessment for the engine control system. This assessment must identify faults or failures that result in a change in thrust or power, transmission of erroneous data, or an effect on engine operability producing a surge or stall together with the predicted frequency of occurrence of these faults or failures.
(f) Protection systems
(1) The design and functioning of engine control devices and systems, together with engine instruments and operating and maintenance instructions, must provide reasonable assurance that those engine operating limitations that affect turbine, compressor, fan, and turbosupercharger rotor structural integrity will not be exceeded in service. (amended 2010/01/29; previous version)
(2) When electronic overspeed protection systems are provided, the design must include a means for testing, at least once per engine start/stop cycle, to establish the availability of the protection function. The means must be such that a complete test of the system can be achieved in the minimum number of cycles. If the test is not fully automatic, the requirement for a manual test must be contained in the engine instructions for operation. (amended 2010/01/29; previous version)
(g) Software. The applicant must design, implement and verify all associated software to minimize the existence of errors by using a method, approved by the Minister, consistent with the criticality of the performed functions. (amended 2010/01/29; previous version)
(h) Aircraft-supplied data. Single failures leading to loss, interruption or corruption of aircraft-supplied data (other than thrust or power command signals from the aircraft) or data shared between engines must: (amended 2010/01/29; previous version)
(1) Not result in a hazardous engine effect for any engine; and (amended 2010/01/29; previous version)
(2) Be detected and accommodated. The accommodation strategy must not result in an unacceptable change in thrust or power or an unacceptable change in engine operating and starting characteristics. The applicant must evaluate and document in the engine installation instructions the effects of these failures on engine power or thrust, engine operability and starting characteristics throughout the flight envelope. (amended 2010/01/29; previous version)
(i) Aircraft-supplied electrical power
(1) The applicant must design the engine control system so that the loss, malfunction, or interruption of electrical power supplied from the aircraft to the engine control system will not result in any of the following: (amended 2010/01/29; previous version)
(i) A hazardous engine effect; or (amended 2010/01/29; previous version)
(ii) The unacceptable transmission of erroneous data. (amended 2010/01/29; previous version)
(2) When an engine dedicated power source is required for compliance with paragraph (i)(1) of this section, its capacity should provide sufficient margin to account for engine operation below idle where the engine control system is designed and expected to recover engine operation automatically. (amended 2010/01/29; previous version)
(3) The applicant must identify and declare the need for and the characteristics of, any electrical power supplied from the aircraft to the engine control system for starting and operating the engine, including transient and steady state voltage limits, in the engine instructions for installation. (amended 2010/01/29; previous version)
(4) Low voltage transients outside the power supply voltage limitations declared in paragraph (i)(3) of this section shall meet the requirements of paragraph (i)(1) of this section. The engine control system must be capable of resuming normal operation when aircraft-supplied power returns to within the declared limits. (amended 2010/01/29; previous version)
(j) Air pressure signal. The applicant must consider the effects of blockage or leakage of the signal lines on the engine control system as part of the System Safety Assessment of (e) of this section and shall adopt the appropriate design precautions. (amended 2010/01/29; previous version)
(k) Automatic availability and control of engine power for 30-second OEI rating. Rotorcraft engines having a 30-second OEI rating shall incorporate a means, or a provision for a means, for automatic availability and automatic control of the 30-second OEI power within its operating limitations. (amended 2010/01/29; previous version)
(l) Engine shut down means. Means must be provided for shutting down the engine rapidly. (amended 2010/01/29; previous version)
(m) Programmable logic devices. The development of programmable logic devices using digital logic or other complex design technologies must provide a level of assurance for the encoded logic commensurate with the hazard associated with the failure or malfunction of the systems in which the devices are located. The applicant must provide evidence that the development of these devices has been done by using a method, approved by the Minister, that is consistent with the criticality of the performed function.
(Change 533-4 (93-12-30))
533.29 Instrument Connection
(c) Each rotorcraft turbine engine having a 30‑second OEI rating and a 2‑minute OEI rating must have a means or provision for a means to:
(amended 2010/01/29; previous version) (1) Alert the pilot when the engine is at the 30‑second OEI and the 2‑minute OEI power levels when the event begins and when the time interval expires;
(amended 2010/01/29; previous version) (3) Alert maintenance personnel in a positive manner that the engine has been operated at either or both of the 30-second and 2- minute OEI power levels and permit retrieval of the recorded data; and
(amended 2010/01/29; previous version) (4) Enable routine verification of the proper operation of the above means. (amended 2010/01/29; previous version) (d) The means or the provision for a means of paragraphs (c)(2) and (c)(3) of this section must not be capable of being reset in flight.
(amended 2010/01/29; previous version) (e) The applicant must make provision for the installation of instrumentation necessary to ensure operation in compliance with engine operating limitations. Where, in presenting the safety analysis or complying with any other requirement, dependence is placed on instrumentation that is not otherwise mandatory, in the assumed aircraft installation, then the applicant must specify this instrumentation in the engine installation instructions and declare it mandatory in the engine approval documentation.
(amended 2010/01/29; previous version) (f) As part of the System Safety Assessment of 533.28(e), the applicant must assess the possibility and subsequent effect of incorrect fit of instruments, sensors or connectors. Where necessary, the applicant must take design precautions to prevent incorrect configuration of the system.
(amended 2010/01/29; previous version) (g) The sensors, together with associated wiring and signal conditioning, must be segregated, electrically and physically, to the extent necessary to ensure that the probability of a fault propagating from instrumentation and monitoring functions to control functions, or vice versa, is consistent with the failure effect of the fault.
(amended 2010/01/29; previous version) (h) The applicant must provide instrumentation enabling the flight crew to monitor the functioning of the turbine cooling system unless appropriate inspections are published in the relevant manuals and evidence shows that:
(amended 2010/01/29; previous version) (1) Other existing instrumentation provides adequate warning of failure or impending failure;
(amended 2010/01/29; previous version) (2) Failure of the cooling system would not lead to hazardous engine effects before detection; or
(amended 2010/01/29; previous version) (3) The probability of failure of the cooling system is extremely remote.
(amended 2010/01/29; previous version) (Change 533-5)
SUBCHAPTER C DESIGN & CONSTRUCTION; RECIPROCATING AIRCRAFT ENGINES
533.31 Applicability
This subchapter prescribes additional design and construction requirements for reciprocating aircraft engines.
533.33 Vibration
The engine must be designed and constructed to function throughout its normal operating range of crank-shaft rotational speeds and engine powers without inducing excessive stress in any of the engine parts because of vibration and without imparting excessive vibration forces to the aircraft structure.
533.34 Turbocharger Rotors
Each turbocharger case shall be designed and constructed to be able to contain fragments of a compressor or turbine that fails at the highest speed that is obtainable with normal speed control devices inoperative.
533.35 Fuel & Induction System
(b) The intake passages of the engine through which air or fuel in combination with air passes for combustion purposes must be designed and constructed to minimise the danger of ice accretion in those passages. The engine must be designed and constructed to permit the use of a means for ice prevention.
533.37 Ignition System
533.39 Lubrication System
(a) The lubrication system of the engine must be designed and constructed so that it will function properly in all flight attitudes and atmospheric conditions in which the aeroplane is expected to operate. In wet sump engines, this requirement must be met when only one-half of the maximum lubricant supply is in the engine.
SUBCHAPTER D BLOCK TESTS RECIPROCATING AIRCRAFT ENGINES
533.41 Applicability
This subchapter prescribes the block tests and inspections for reciprocating aircraft engines.
(Change 533-1 (87-01-01))
533.42 General
Before each endurance test required by this subchapter, the adjustment setting and functioning characteristic of each component having an adjustment setting and a functioning characteristic that can be established independent of installation on the engine must be established and recorded.
533.43 Vibration Test
(a) Each engine must undergo a vibration survey to establish the torsional and bending vibration characteristics of the crankshaft and the propeller shaft or other output shaft, over the range of crankshaft speed and engine power, under steady state and transient conditions, from idling speed to either 110 percent of the desired maximum continuous speed rating or 103 percent of the maximum desired take-off speed rating, whichever is higher. The survey must be conducted using, for aeroplane engines, the same configuration of the propeller type which is used for the endurance test, and using, for other engines, the same configuration of the loading device type which is used for the endurance test.
(d) The vibration survey described in paragraph (a) of this section must be repeated with that cylinder not firing which has the most adverse vibration effect, in order to establish the conditions under which the engine can be operated safely in that abnormal state. However, for this vibration survey, the engine speed range need only extend from idle to the maximum desired take-off speed, and compliance with paragraph (b) of this section need not be shown.
533.45 Calibration Tests
(a) Each engine must be subjected to the calibration tests necessary to establish its power characteristics and the conditions for the endurance test specified in 533.49. The results of the power characteristics calibration tests form the basis for establishing the characteristics of the engine over its entire operating range of crankshaft rotational speeds, manifold pressures, fuel/air mixture settings, and altitudes. Power ratings are based upon standard atmospheric conditions with only those accessories installed which are essential for engine functioning.
533.47 Detonation Test
533.49 Endurance Test
(a) General. Each engine must be subjected to an endurance test that includes a total of 150 hours of operation (except as provided in paragraph (e)(1))(iii) of this section) and, depending upon the type and contemplated use of the engine, consists of one of the series of runs specified in paragraphs (b) through (e) of this section, as applicable. The runs must be made in order found appropriate by the Minister for the particular engine being tested. During the endurance test the engine power and the crankshaft rotational speed must be kept within +3 percent of the rated values. During the runs at rated take-off power and for at least 35 hours at rated maximum continuous power, one cylinder, must be operated at not less than the limiting temperature, the other cylinders must be operated at a temperature not lower than 50 degrees F below the limiting temperature, and the oil inlet temperature must be maintained within +10 degrees F of the limiting temperature. An engine that is equipped with a propeller shaft must be fitted for the endurance test with a propeller that thrust-loads the engine to the maximum thrust which the engine is designed to resist at each applicable operating condition specified in this section. Each accessory drive and mounting attachment must be loaded. During operation at rated take-off power and rated maximum continuous power, the load imposed by each accessory used only for an aircraft service must be the limit load specified by the applicant for the engine drive or attachment point.
(b) Unsupercharged engines and engines incorporating a gear-driven single-speed supercharger. For engines not incorporating a super-charger and for engines incorporating a gear-driven single-speed supercharger the applicant must conduct the following runs:
(1) Alternate periods of 5 minutes at rated take-off power with take-off speed, and 5 minutes at maximum best economy cruising power or maximum recommended cruising power.
(3) A 20-hour run consisting of alternate periods of 1 1/2 hours at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 70 percent rated maximum continuous power at 89 percent maximum continuous speed.
(5) A 20-hour run consisting of alternate periods of 1 1/2 hours at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 60 percent rated maximum continuous power and 84.5 percent maximum continuous power.
(c) Engines incorporating a gear driven two-speed supercharger. For engines incorporating a gear-driven two-speed supercharger the applicant must conduct the following runs:
(1) A 30-hour run consisting of alternate periods in the lower gear ratio of five minutes at rated take-off power with take-off speed, and five minutes at maximum best economy cruising power or at maximum recommended cruising power. If a take-off power rating is desired in the higher gear ratio, 15 hours of the 30-hour run must be made in the higher gear ratio in alternate periods of five minutes at the observed horsepower obtainable with the take-off critical altitude manifold pressure and take-off speed, and five minutes at 70 percent high ratio rated maximum continuous power and 89 percent high ratio maximum continuous speed.
(2) A 15-hour run consisting of alternate periods in the lower gear ratio of one hour at rated maximum continuous power with maximum continuous power and 91 percent maximum continuous speed.
(3) A 15-hour run consisting of alternate periods in the lower gear ratio of one hour at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 70 percent rated maximum continuous power and 89 percent maximum continuous speed.
(5) A 5-hour run consisting of alternate periods of five minutes in each of the supercharger gear ratios. The first five minutes of the test must be made at maximum continuous speed in the higher gear ratio and observed horsepower obtainable with 90 percent of maximum continuous manifold pressure in the higher gear ratio under sea level conditions. The condition for operation for the alternate five minutes in the lower gear ratio must be that obtained by shifting to the lower gear ratio at constant speed.
(6) A 10-hour run consisting of alternate periods in the lower gear ratio of one hour at rated maximum continuous power with maximum continuous speed, and one hour at 65 percent rated maximum continuous power and 84.5 percent maximum continuous speed.
(7) A 10-hour run consisting of alternate periods in the lower gear ratio of one hour at rated maximum continuous power with maximum continuous speed, and one hour at 60 percent rated maximum continuous power and 84.5 percent maximum continuous speed.
(8) A 10-hour run consisting of alternate periods in the lower gear ratio of one hour at rated maximum continuous power with maximum continuous speed, and one hour at 50 percent rated maximum continuous power and 79.5 percent maximum continuous speed.
(9) A 20-hour run consisting of alternate periods in the lower gear ratio of 2 hours at rated maximum continuous power with maximum continuous speed, and two hours of maximum best economy cruising power and speed or at maximum recommended cruising power.
(10) A 5-hour run in the lower gear ratio at maximum best economy cruising power and speed or a maximum recommended cruising power and speed.
(d) Helicopter engines. To be eligible for use on a helicopter each engine must either comply with paragraphs (a) through (j) of 529.923 of this Manual, or must undergo the following series of runs:
(1) A 35-hour run consisting of alternate periods of 30 minutes each at rated take-off power with take-off speed, and at rated maximum continuous power with maximum continuous speed.
(4) A 25-hour run consisting of alternate periods of 2 1/2 hours each at 80 percent rated maximum continuous power with take-off speed, and at 80 percent rated maximum continuous power with 80 to 90 percent maximum continuous speed.
(5) A 25-hour run consisting of alternate periods of 2 1/2 hours each 80 percent rated maximum continuous power with take-off speed, and at either rated maximum continuous power with 110 percent maximum continuous speed or at rated take-off power with 103 percent take-off speed, whichever results in the greater speed.
(6) A 15-hour run at 105 percent rated maximum continuous power with 105 percent maximum continuous speed or at full throttle and corresponding speed at standard sea level carburettor entrance pressure, if 105 percent of the rated maximum continuous power is not exceeded.
(e) Turbosupercharged engines. For engines incorporating a turbosupercharger the following apply except that altitude testing may be simulated provided the applicant shows that the engine and supercharger are being subjected to mechanical loads and operating temperatures no less severe than if run at actual altitude conditions.
(1) For engines used in aeroplanes the applicant must conduct the runs specified in paragraph (b) of this section, except:
(i) The entire run specified in subparagraph (b)(1) of this section must be made at sea level altitude pressure;
(ii) The portions of the runs specified in subparagraph (b)(2) through (7) of this section at rated maximum continuous power must be made at critical altitude pressure, and the portions of the runs at other power must be made at 8,000 feet altitude pressure; and
(iii) The turbosupercharger used during the 150-hour endurance test must be run on the bench for an additional 50 hours at the limiting turbine wheel inlet gas temperature and rotational speed for rated maximum continuous power operation unless the limiting temperature and of the rated maximum continuous power operation.
(2) For engines used in helicopters and applicant must conduct the runs specified in paragraph (d) of this section, except:
(ii) The portions of the runs specified in paragraph (d)(2) and of this section at rated maximum continuous power must be made at critical altitude pressure and the portions of the runs at other power must be made at 8,000 feet altitude pressure;
533.51 Operation Test
The operation test must include the testing found necessary by the Minister to demonstrate backfire characteristics, starting, idling, acceleration, overspeeding, functioning of propeller and ignition, and any other operational characteristic of the engine. If the engine incorporates a multispeed super-charger driver, the design and construction must allow the supercharger to be shifted from operation at the lower speed ratio to the higher and the power appropriate to the manifold pressure and speed settings for rated maximum continuous power at the higher supercharger speed ratio must be obtainable within five seconds.
533.53 Engine System and Component Tests
(a) For those systems and components that cannot be adequately substantiated in accordance with endurance testing of 533.49, the applicant must conduct additional tests to demonstrate that systems or components are able to perform the intended functions in all declared environmental and operating conditions.
533.55 Teardown Inspection
After completing the endurance test:
(c) Each engine component must conform to the type design and be eligible for incorporation into an engine for continued operation, in accordance with information submitted in compliance with 533.4.
533.57 General Conduct of Block Tests
(b) The applicant may service and make minor repairs to the engine during the block tests in accordance with the service and maintenance instructions submitted in compliance with 533.4. If the frequency of the service is excessive, or the number of stops due to engine malfunction is excessive, or a major repair, or replacement of a part is found necessary during the block tests or as a result of findings from the teardown inspection, the engine or its parts may be subjected to any additional test the Minister finds necessary.
SUBCHAPTER E DESIGN AND CONSTRUCTION TURBINE AIRCRAFT ENGINES
533.61 Applicability
This subchapter prescribes additional design and construction requirements for turbine aircraft engines.
533.62 Stress Analysis
A stress analysis must be performed on each turbine engine showing the design safety margin of each turbine engine rotor, spacer, and rotorshaft.
533.63 Vibration
[Each engine must be designed and constructed to function throughout its declared flight envelope and operating range of rotational speeds and power/thrust, without inducing excessive stress in any engine part because of vibration and without imparting excessive vibration forces to the aircraft structure.]
533.64 Pressurized Engine Static Parts
(amended 2010/01/29; no previous version)
(1) Exhibit permanent distortion beyond serviceable limits or exhibit leakage that could create a hazardous condition when subjected to the greater of the following pressures: (amended 2010/01/29; no previous version)
(i) 1.1 times the maximum working pressure; (amended 2010/01/29; no previous version)
(ii) 1.33 times the normal working pressure; or (amended 2010/01/29; no previous version)
(iii) 35 kPa (5 p.s.i.) above the normal working pressure. (amended 2010/01/29; no previous version)
(2) Exhibit fracture or burst when subjected to the greater of the following pressures: (amended 2010/01/29; no previous version)
(i) 1.15 times the maximum possible pressure; (amended 2010/01/29; no previous version)
(ii) 1.5 times the maximum working pressure; or (amended 2010/01/29; no previous version)
(iii) 35 kPa (5 p.s.i.) above the maximum possible pressure. (amended 2010/01/29; no previous version)
(b) Compliance with this section must take into account: (amended 2010/01/29; no previous version)
(1) The operating temperature of the part; (amended 2010/01/29; no previous version)
(2) Any other significant static loads in addition to pressure loads; (amended 2010/01/29; no previous version)
(3) Minimum properties representative of both the material and the processes used in the construction of the part; and (amended 2010/01/29; no previous version)
533.65 Surge & Stall Characteristics
When the engine is operated in accordance with operating instructions required by 533.5(b), starting, a change of power or thrust, power or thrust augmentation, limiting inlet air distortion, or inlet air temperature may not cause surge or stall to the extent that flameout, structural failure, overtemperature, or failure of the engine to recover power or thrust will occur at any point in the operating envelope.
533.66 Bleed Air System
The engine must supply bleed air without adverse effect on the engine, excluding reduced thrust or power output, at all conditions set up to the discharge flow conditions established as a limitation under 533.7(c)(11). If bleed air used for engine anti-icing can be controlled, provision must be made for a means to indicate the functioning of the engine ice protection system.
533.67 Fuel System
(a) With fuel supplied to the engine at the flow and pressure specified by the applicant, the engine must function properly under each operating condition required by this Chapter. Each fuel control adjusting means that may not be manipulated while the fuel control device is mounted on the engine must be secured by a locking device and sealed, or otherwise be inaccessible. All other fuel control adjusting means must be accessible and marked to indicate the functioning of the adjustment unless the function is obvious.
(b) There must be a fuel strainer or filter between the engine fuel inlet opening and the inlet of either the fuel metering device or the engine-driven positive displacement pump whichever is nearer the engine fuel inlet. In addition, the following provisions apply to each strainer or filter required by this paragraph:
(ii) That the fuel system is capable of sustaining operation through-out its flow and pressure range with the fuel initially saturated with water at 80°F (27°C) and having 0.025 fluid ounces per gallon (0.20 millilitres per litre) of free water added and cooled to the most critical condition for icing likely to be encountered in operation. However; this requirement may be met by demonstrating the effectiveness of specified approved fuel anti-icing additives, or that the fuel system incorporates a fuel heater which maintains the fuel temperature at the fuel strainer or fuel inlet above 32°F (0°C) under the most critical conditions.
(5) The applicant must demonstrate that the filtering means has the capacity (with respect to engine operating limitations) to ensure that the engine will continue to operate within approved limits, with fuel contaminated to the maximum degree of particle size and density likely to be encountered in service. Operation under these conditions must be demonstrated for a period acceptable to the Minister, beginning when indication of impending filter blockage is first given by either:
(6) Any strainer or filter bypass must be designed and construction so that the release of collected contaminants is minimised by appropriate location of the bypass to ensure that collected contaminants are not in the bypass flow path.
(d) Rotorcraft engines having a 30-second OEI rating must incorporate a means or a provision for a means for automatic availability and automatic control of the 30-second OEI power within its operating limitations.
533.68 Induction System Icing
(a) Operate throughout its flight power range (including idling) with out the accumulation of ice on the engine components that adversely affects engine operation or that causes a serious loss of power or thrust in continuous maximum and intermittent maximum icing conditions as defined in Appendix C of Chapter 525 of this Manual; and
(b) Idle for 30 minutes on the ground, with the available air bleed for icing protection at its critical condition, without adverse effect, in an atmosphere that is at a temperature between 15° and 30°F (between -9° and -1°C) and has a liquid water content not less than 0.3 grams per cubic metre in the form of drops having a mean effective diameter not less than 20 microns, followed by a momentary operation at take-off power or thrust. During the 30 minutes of idle operation the engine may be run up periodically to a moderate power or thrust setting in a manner acceptable to the Minister.
533.69 Ignitions System
Each engine must be equipped with an ignition system for starting the engine on the ground and in flight. An electric ignition system must have at least two igniter and two separate secondary electric circuits, except that only one igniter is required for fuel burning augmentation systems.
533.70 Engine Life-limited Parts
By a procedure approved by the Minister, operating limitations shall be established which specify the maximum allowable number of flight cycles for each engine life-limited part. Engine life-limited parts are rotor and major static structural parts whose primary failure is likely to result in a hazardous engine effect. Typically, engine life- limited parts include, but are not limited to disks, spacers, hubs, shafts, high-pressure casings, and non- redundant mount components. For the purposes of this section, a hazardous engine effect is any of the conditions listed in 533.75. The applicant will establish the integrity of each engine life-limited part by:
(a) An engineering plan that contains the steps required to ensure each engine life-limited part is withdrawn from service at an approved life before hazardous engine effects can occur. These steps include validated analysis, test, or service experience which ensures that the combination of loads, material properties, environmental influences and operating conditions, including the effects of other engine parts influencing these parameters, are sufficiently well known and predictable so that the operating limitations can be established and maintained for each engine life-limited part. Applicants shall perform appropriate damage tolerance assessments to address the potential for failure from material, manufacturing, and service induced anomalies within the approved life of the part. Applicants shall publish a list of the life-limited engine parts and the approved life for each part in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness as required by 533.4.
533.71 Lubrication System
(1) Each strainer or filter required by this paragraph that has a bypass must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer of filter element completely blocked.
(3) Each strainer or filter required by this paragraph must have the capacity (with respect to operating limitations established for the engine) to ensure that engine oil system functioning is not impaired with the oil contaminated to a degree (with respect to particle size and density) that is greater than that established for the engine in subparagraph (2) of this paragraph.
(5) Any filter bypass must be designed and constructed so that the release of collected contaminants is minimised by appropriate location of the bypass to ensure that the collected contaminants are not in the bypass flow path.
(6) Each strainer or filter required by this paragraph that has no bypass, except the strainer or filter at an oil tank outlet or for a scavenge pump, must have provisions for connection with a warning means to warn the pilot of the occurrence of contamination of the screen before it reaches the capacity established in accordance with subparagraph (3) of this paragraph.
(c) Oil tanks.
(1) Each oil tank must have an expansion space for not less than 10 percent of the tank capacity.
(4) Each oil tank cap must provide an oil-tight seal.
(5) Each oil tank filler must be marked with the word "oil".
(6) Each oil tank must be vented from the top part of the expansion space, with the vent so arranged that condensed water vapour that might freeze and obstruct the line cannot accumulate at any point.
(8) There must be a shut-off valve at the outlet of each oil tank, unless the external portion of the oil system (including oil tank supports) is fireproof.
(9) Each unpressurised oil tank must not leak when subjected to a maximum operating temperature and an internal pressure of 5 p.s.i., and each pressured oil tank must meet the requirements of 533.64.
(11) Each oil tank must have an oil quantity indicator or provision for one.
(12) If the propeller feathering system depends on engine oil:
(i) There must be means to trap an amount of oil in the tank if the supply becomes depleted due to failure of any part of the lubricating system other than the tank itself.
(ii) The amount of trapped oil must be enough to accomplish the feathering operation and must be available only to the feathering pump; and
(d) Oil Drains. A drain (or drains) must be provided to allow safe drainage of the oil system.
Each drain must:
533.72 Hydraulic Actuating Systems
Each hydraulic actuating system must function properly under all conditions in which the engine is expected to operate. Each filter or screen must be accessible for servicing and each tank must meet the design criteria of 533.71.
533.73 Power or Thrust Response
The design and construction of the engine must enable an increase:
(a) From minimum to rated take-off power or thrust with the maximum bleed air and power extraction to be permitted in an aircraft, without overtemperature, surge, stall, or other detrimental factors occurring to the engine whenever the power control lever is moved from the minimum to the maximum position in not more than 1 second, except that the Minister may allow additional time increments for different regimes of control operation requiring control scheduling; and
(b) From the fixed minimum flight idle power lever position when provided, or if not provided, from not more than 15% of the rated take-off power or thrust available to 95% rated take-off power or thrust is not over 5 seconds. The 5-second power or thrust response must occur from a stabilised static condition using only the bleed air and accessories loads necessary to run the engine. This take-off rating is specified by the applicant and need not include thrust augmentation.
[533.74 Continued Rotation
If any of the engine main rotating systems continue to rotate after the engine is shutdown for any reason while in flight, and if means to prevent that continued rotation are not provided, then any continued rotation during the maximum period of flight, and in the flight conditions expected to occur with that engine inoperative, shall not result in any condition described in 533.75 (g)(2)(i) through (vi) of this chapter.
533.75 Safety Analysis
It must be shown by analysis that any probable malfunction or any probable single or multiple failure, or any probable improper operation of the engine will not cause the engine to:
(1) The applicant shall analyze the engine, including the control system, to assess the likely consequences of all failures that can reasonably be expected to occur. This analysis will take into account, if applicable:
(i) Aircraft-level devices and procedures assumed to be associated with a typical installation. Such assumptions shall be stated in the analysis. (amended 2008/10/30; no previous version)
(2) The applicant shall summarize those failures that could result in major engine effects or hazardous engine effects, as defined in paragraph (g) of this section, and estimate the probability of occurrence of those effects. Any engine part the failure of which could reasonably result in a hazardous engine effect shall be clearly identified in this summary.
(3) The applicant shall show that hazardous engine effects are predicted to occur at a rate not in excess of that defined as extremely remote (probability range of 10-7 to 10-9 per engine flight hour). Since the estimated probability for individual failures may be insufficiently precise to enable the applicant to assess the total rate for hazardous engine effects, compliance may be shown by demonstrating that the probability of a hazardous engine effect arising from an individual failure can be predicted to be no greater than 10‑8 per engine flight hour. In dealing with probabilities of this low order of magnitude, absolute proof is not possible, and compliance may be shown by reliance on engineering judgment and previous experience combined with sound design and test philosophies.
(4) The applicant shall show that major engine effects are predicted to occur at a rate not in excess of that defined as remote (probability range of 10-5 to 10-7 per engine flight hour).
(b) The Minister may require that any assumption as to the effects of failures and likely combination of failures be verified by test. (amended 2008/10/30; previous version)
(c) The primary failure of certain single elements cannot be sensibly estimated in numerical terms. If the failure of such elements is likely to result in hazardous engine effects, then compliance may be shown by reliance on the prescribed integrity requirements of 533.15, 533.27, and 533.70 as applicable. These instances shall be stated in the safety analysis.
(d) If reliance is placed on a safety system to prevent a failure from progressing to hazardous engine effects, the possibility of a safety system failure in combination with a basic engine failure shall be included in the analysis. Such a safety system may include safety devices, instrumentation, early warning devices, maintenance checks, and other similar equipment or procedures. If items of a safety system are outside the control of the engine manufacturer, the assumptions of the safety analysis with respect to the reliability of these parts shall be clearly stated in the analysis and identified in the installation instructions under 533.5 of this chapter.
(e) If the safety analysis depends on one or more of the following items, those items shall be identified in the analysis and appropriately substantiated. (amended 2008/10/30; no previous version)
(1) Maintenance actions being carried out at stated intervals. This includes the verification of the serviceability of items that could fail in a latent manner. When necessary to prevent hazardous engine effects, these maintenance actions and intervals shall be published in the instructions for continued airworthiness required under 533.4 of this chapter. Additionally, if errors in maintenance of the engine, including the control system, could lead to hazardous engine effects, the appropriate procedures shall be included in the relevant engine manuals.
(2) Verification of the satisfactory functioning of safety or other devices at pre-flight or other stated periods. The details of this satisfactory functioning shall be published in the appropriate manual.
(4) Flight crew actions to be specified in the operating instructions established under 533.5. (amended 2008/10/30; no previous version)
(f) If applicable, the safety analysis shall also include, but not be limited to, investigation of the following:
(g) Unless otherwise approved by the Minister and stated in the safety analysis, for compliance with chapter 533, the following failure definitions apply to the engine: (amended 2008/10/30; no previous version)
533.76 Bird Ingestion
(amended 2001/03/05; no previous version)
Compliance with (b), (c) and (d) of this section shall be in accordance with the following:
(1) except as specified in paragraph (d) of this section, all ingestion tests shall be conducted with the engine stabilized at no less than 100‑percent take‑off power or thrust, for test day ambient conditions prior to the ingestion. In addition, the demonstration of compliance shall account for engine operation at sea level take‑off conditions on the hottest day that a minimum engine can achieve maximum rated take‑off thrust or power;
(2) the engine inlet throat area as used in this section to determine the bird quantity and weights shall be established by the applicant and identified as a limitation in the installation instructions required under section 533.5;
(3) the impact to the front of the engine from the large single bird, the single largest medium bird which can enter the inlet and the large flocking bird shall be evaluated. Applicants shall demonstrate that the associated components when struck under the conditions prescribed in paragraphs (b), (c) or (d) of this section, as applicable, will not affect the engine to the extent that the engine cannot comply with the requirements of (b)(3), (c)(6) and (d)(4) of this section;
(4) for an engine that incorporates an inlet protection device, compliance with this section shall be established with the device functioning. The engine approval shall be endorsed to demonstrate that compliance with the requirements has been established with the device functioning;
(5) objects that are accepted by the Minister may be substituted for birds when conducting the bird ingestion tests required by (b), (c) and (d) of this section; and
(6) if compliance with the requirements of this section is not established, the engine type certification documentation shall demonstrate that the engine shall be limited to aircraft installations in which it is demonstrated that a bird cannot strike the engine, or be ingested into the engine, or adversely restrict airflow into the engine.
(b) Large single birds
Compliance with the large bird ingestion requirements shall be in accordance with the following:
(1) the large bird ingestion test shall be conducted using one bird of a weight determined from Table 1 aimed at the most critical exposed location on the first stage rotor blades and ingested at a bird speed of 200-knots for engines to be installed on aeroplanes, or the maximum airspeed for normal rotorcraft flight operations for engines to be installed on rotorcraft;
(2) power lever movement shall not be permitted within 15 seconds following ingestion of the large bird;
(3) ingestion of a single large bird tested under the conditions prescribed in this section shall not result in any condition described in 533.75(g)(2) of this chapter.
(4) Compliance with the large bird ingestion requirements of this paragraph may be established by demonstrating that the requirements of 533.94(a) constitute a more severe demonstration of blade containment and rotor unbalance than the requirements of this paragraph.
Table 1 to 533.76 -- Large Bird Weight Requirements
Engine Inlet Throat Area (A)-Square/meters
Bird weight kg (lb)
1.35 (2,092)> A.....................
1.85 (4.07) minimum, unless a smaller bird is determined to be a more severe demonstration.
1.35 (2,092) <A <3.90 (6,045)
(amended 2004/06/08; previous version)
2.75 (6.05)
3.90 (6,045) <A ......................
3.65 (8.03)
(c) Small and medium flocking birds
Compliance with the small and medium bird ingestion requirements shall be in accordance with the following:
(1) analysis or component test, or both, acceptable to the Minister, shall be conducted to determine the critical ingestion parameters affecting power loss and damage. Critical ingestion parameters shall include, but are not limited to, the effects of bird speed, critical target location, and first stage rotor speed. The critical bird ingestion speed should reflect the most critical condition within the range of airspeeds used for normal flight operations up to 1,500 feet above ground level, but not less than V1 minimum for aeroplanes;
(amended 2004/06/08; previous version) (2) medium bird engine tests shall be conducted so as to simulate a flock encounter, and will use the bird weights and quantities specified in Table 2. When only one bird is specified, that bird will be aimed at the engine core primary flow path; the other critical locations on the engine face area must be addressed, as necessary, by appropriate tests or analysis, or both. When two or more birds are specified in Table 2, the largest of those birds shall be aimed at the engine core primary flow path, and a second bird shall be aimed at the most critical exposed location on the first stage rotor blades. Any remaining birds shall be evenly distributed over the engine face area;
(3) in addition, except for rotorcraft engines, it shall also be substantiated by appropriate tests or analysis or both, that when the full fan assembly is subjected to the ingestion of the quantity and weights of bird from Table 3, aimed at the fan assembly’s most critical location outboard of the primary core flowpath, and in accordance with the applicable test conditions of this paragraph, that the engine can comply with the acceptance criteria of this paragraph;
(4) a small bird ingestion test is not required if the prescribed number of medium birds pass into the engine rotor blades during the medium bird test;
(5) small bird ingestion tests shall be conducted so as to simulate a flock encounter using one 85 gram (0.187 lb.) bird for each 0.032 square-meter (49.6 square-inches) of inlet area, or fraction thereof, up to a maximum of 16 birds. The birds shall be aimed so as to account for any critical exposed locations on the first stage rotor blades, with any remaining birds evenly distributed over the engine face area;
(6) ingestion of small and medium birds tested under the conditions prescribed in this paragraph shall not cause any of the following:
(i) more than a sustained 25-percent power or thrust loss,
(ii) the engine to be shut down during the required run-on demonstration prescribed in paragraphs (c)(7) or (c)(8) of this section,
(iii) the conditions defined in paragraph (b)(3) of this section, and
(iv) unacceptable deterioration of engine handling characteristics;
(i) ingestion so as to simulate a flock encounter, with approximately 1 second elapsed time from the moment of the first bird ingestion to the last,
(ii) followed by 2 minutes without power lever movement after the ingestion,
(iii) followed by 3 minutes at 75-percent of the test condition,
(iv) followed by 6 minutes at 60-percent of the test condition,
(v) followed by 6 minutes at 40-percent of the test condition,
(vi) followed by 1 minute at approach idle,
(vii) followed by 2 minutes at 75‑percent of the test condition,
(viii) followed by stabilising at idle and engine shut down, and
(ix) the durations specified are times at the defined conditions with the power being changed between each condition in less than 10 seconds; and
(i) ingestion so as to simulate a flock encounter within approximately 1 second elapsed time between the first ingestion and the last,
(ii) followed by 3 minutes at 75-percent of the test condition,
(iii) followed by 90 seconds at descent flight idle,
(iv) followed by 30 seconds at 75-percent of the test condition, and
(v) followed by stabilizing at idle and engine shut down, and
(vi) the durations specified are times at the defined conditions with the power being changed between each condition in less than 10 seconds, and
(amended 2004/06/08; previous version) (9) engines intended for use in multi-engine rotorcraft are not required to comply with the medium bird ingestion portion of this section, providing that the appropriate type certificate documentation is so endorsed; and
Table 2 to 533.76 -- Medium Flocking Bird Weight and Quantity Requirements
Engine Inlet Throat Area (A) -- Square-meters (square-inches)
0.05 (77.5) >A
0.05 (77.5) <A <0.10((155)
0.10 (155) <A <0.20 (310)
0.20 (310) <A <0.40 (620)
0.40 (620) <A <0.60 (930)
0.60 (930) <A <1.00 (1,550)
1.00 (1,550) <A <1.35 (2,092)
1.35 (2,092) <A <1.70 (2,635)
1.70 (2,635) <A <2.10 (3,255)
2.10 (3,255) <A <2.50 (3,875)
2.50 (3,875) <A <3.90 (6045)
3.90 (6045) <A <4.50 (6975)
4.50 (6975) <A
Table 3 to 533.76 -- Additional Integrity Assessment
1.35 (2,092)>A
1.35 (2,092) <A <2.90 (4,495)
1.15(2.53)
2.90 (4,495) <A <3.90 (6,045)
3.90 (6,045) <A
(d) Large flocking bird
An engine test will be performed as follows:
(1) Large flocking bird engine tests must be performed using the bird mass and weights in Table 4, and ingested at a bird speed of 200 knots. (amended 2008/10/30; no previous version)
(2) Prior to the ingestion, the engine shall be stabilized at no less than the mechanical rotor speed of the first exposed stage or stages that, on a standard day, would produce 90 percent of the sea level static maximum rated take-off power or thrust.
(3) The bird shall be targeted on the first exposed rotating stage or stages at a blade airfoil height of no less than 50 percent measured at the leading edge.
(4) Ingestion of a large flocking bird under the conditions prescribed in (d)(1), (d)(2) and (d)(3) shall not cause any of the following:
(i) A sustained reduction of power or thrust to less than 50 percent of maximum rated take-off power or thrust during the run-on segment specified under (d)(5)(i) of this section.
(ii) Engine shutdown during the required run-on demonstration specified in (d)(5) of this section.
(iii) The conditions specified in (b)(3) of this section.
(5) The following test schedule shall be used:
(ii) Followed by 13 minutes at no less than 50 percent of maximum rated take-off power or thrust.
(iii) Followed by 2 minutes between 30 and 35 percent of maximum rated take-off power or thrust.
(iv) Followed by 1 minute with power or thrust increased from that set in (d)(5)(iii) of this section, by between 5 and 10 percent of maximum rated take-off power or thrust.
(v) Followed by 2 minutes with power or thrust reduced from that set in (d)(5)(iv) of this section, by between 5 and 10 percent of maximum rated take-off power or thrust.
(vi) Followed by a minimum of 1 minute at ground idle then engine shutdown. The durations specified are times at the defined conditions. Power lever movement between each condition will be 10 seconds or less, except that power lever movements allowed within (d)(5)(ii) of this section are not limited, and for setting power under (d)(5)(iii) of this section will be 30 seconds or less.
(6) Compliance with the large flocking bird ingestion requirements of (d) may also be demonstrated by:
(i) Incorporating the requirements of (d)(4) and (d)(5) of this section, into the large single bird test demonstration specified in (b)(1) of this section; or
(ii) Use of an engine subassembly test at the ingestion conditions specified in (b)(1) of this section if:
(A) All components critical to complying with the requirements of (d) of this section are included in the subassembly test;
(B) The components of (d)(6)(ii)(A) of this section are installed in a representative engine for a run-on demonstration in accordance with (d)(4) and (d)(5) of this section; except that (d)(5)(i) is deleted and (d)(5)(ii) shall be 14 minutes in duration after the engine is started and stabilized; and
(C) The dynamic effects that would have been experienced during a full engine ingestion test can be demonstrated to be negligible with respect to meeting the requirements of (d)(4) and (d)(5) of this section.
(7) Applicants shall demonstrate that an unsafe condition will not result if any engine operating limit is exceeded during the run-on period.
Table 4 to 533.76 -- Large Flocking Bird Mass and Weight
(amended 2008/10/30; no previous version) Engine Inlet Throat Area (A) -- square-meters (square-inches)
A < 2.50 (3875)
2.50 (3875) £ A < 3.50 (5425)
1.85 (4.08)
3.50 (5425) £ A < 3.90 (6045)
2.10 (4.63)
3.90 (6045) £ A
2.50 (5.51)
533.77 Foreign Object Ingestion - Ice
(amended 2001/03/05; previous version)
(c) Ingestion of ice under the conditions of paragraph (e) of this section may not
(2) require the engine to be shut down.
(d) For an engine that incorporates protection device, compliance with a this section need not be demonstrate with respect to foreign objects to be ingested under the conditions prescribed in paragraph (e) of this section if it is shown that:
(3) the foreign object, or objects, stopped by the protective device shall not obstruct the flow of induction air into the engine with a resultant sustained reduction in power or thrust greater than those values required by paragraph (c) of this section.
(e) Compliance with paragraph (c) of this section shall be demonstrated by engine test under the following ingestion conditions:
(1) ice quantity shall be the maximum accumulation on a typical inlet cowl and engine face resulting from a 2-minute delay in actuating the anti-icing system; or a slab of ice which is comparable in weight or thickness for that size engine;
(2) the ingestion velocity shall simulate ice being sucked into the engine inlet;
(3) engine operation shall be maximum cruise power or thrust; and
(4) the ingestion shall simulate a continuous maximum icing encounter at 25 degress Fahrenheit.
[533.78 Rain and Hail Ingestion
[(a) All engines.
[(1) The ingestion of large hailstones (0.8 to 0.9 specific gravity) at the maximum true air speed, up to 15,000 feet (4,500 meters), associated with a representative aircraft operating in rough air, with the engine at maximum continuous power, may not cause unacceptable mechanical damage or unacceptable power or thrust loss after the ingestion, or require the engine to be shut down. One-half the number of hailstones shall be aimed randomly over the inlet face area and the other half aimed at the critical inlet face area. The hailstones shall be ingested in a rapid sequence to simulate a hailstone encounter and the number and size of the hailstones shall be determined as follows:
[(i) One 1-inch (25 millimeters) diameter hailstone for engines with inlet areas of not more than 100 square inches (0.0645 square meters). [(ii) One 1-inch (25 millimeters) diameter and one 2-inch (50 millimeters) diameter hailstone for each 150 square inches (0.0968 square meters) of inlet area, or fraction thereof, for engines with inlet areas of more than 100 square inches (0.0645 square meters).
[(2) In addition to complying with paragraph (a)(1) of this section and except as provided in paragraph (b) of this section, it must be shown that each engine is capable of acceptable operation throughout its specified operating envelope when subjected to sudden encounters with the certification standard concentrations of rain and hail, as defined in appendix B to this part. Acceptable engine operation precludes flameout, run down, continued or non-recoverable surge or stall, or loss of acceleration and deceleration capability, during any three minute continuous period in rain and during any 30 second continuous period in hail. It must also be shown after the ingestion that there is no unacceptable mechanical damage, unacceptable power or thrust loss, or other adverse engine anomalies.
[(b) Engines for rotorcraft. As an alternative to the requirements specified in paragraph (a)(2) of this section, for rotorcraft turbine engines only, it must be shown that each engine is capable of acceptable operation during and after the ingestion of rain with an overall ratio of water droplet flow to airflow, by weight, with a uniform distribution at the inlet plane, of at least four percent. Acceptable engine operation precludes flameout, run down, continued or non-recoverable surge or stall, or loss of acceleration and deceleration capability. It must also be shown after the ingestion that there is no unacceptable mechanical damage, unacceptable power loss, or other adverse engine anomalies. The rain ingestion must occur under the following static ground level conditions:
[(1) A normal stabilization period at take-off power without rain ingestion, followed immediately by the suddenly commencing ingestion of rain for three minutes at take-off power, then
[(2) Continuation of the rain ingestion during subsequent rapid deceleration to minimum idle, then
[(3) Continuation of the rain ingestion during three minutes at minimum idle power to be certified for flight operation, then
[(4) Continuation of the rain ingestion during subsequent rapid acceleration to take-off power.
[(c) Engines for supersonic aeroplanes. In addition to complying with paragraphs (a)(1) and (a)(2) of this section, a separate test for supersonic aeroplane engines only, shall be conducted with three hailstones ingested at supersonic cruise velocity. These hailstones shall be aimed at the engine's critical face area, and their ingestion must not cause unacceptable mechanical damage or unacceptable power or thrust loss after the ingestion or require the engine to be shut down. The size of these hailstones shall be determined from the linear variation in diameter from 1-inch (25 millimeters) at 35,000 feet (10,500 meters) to 1/4 - inch (6 millimeters) at 60,000 feet (18,000 meters) using the diameter corresponding to the lowest expected supersonic cruise altitude. Alternatively, three larger hailstones may be ingested at subsonic velocities such that the kinetic energy of these larger hailstones is equivalent to the applicable supersonic ingestion conditions.
[(d) For an engine that incorporates or requires the use of a protection device, demonstration of the rain and hail ingestion capabilities of the engine, as required in paragraphs (a), (b), and (c) of this section, may be waived wholly or in part by the Minister if the applicant shows that:
[(1) The subject rain and hail constituents are of a size that will not pass through the protection device;
[(2) The protection device will withstand the impact of the subject rain and hail constituents; and
[(3) The subject of rain and hail constituents, stopped by the protection device, will not obstruct the flow of induction air into the engine, resulting in damage, power or thrust loss, or other adverse engine anomalies in excess of what would be accepted in paragraphs (a), (b), and (c) of this section.]
533.79 Fuel Burning Thrust Augmentor
Each fuel burning thrust augmentor, including the nozzle, must:
(a) Provide cut-off of the fuel burning thrust augmentor;
SUBCHAPTER F BLOCK TESTS TURBINE AIRCRAFT ENGINES
533.81 Applicability
This subchapter prescribes the block tests and inspections for turbine engines.
533.82 General
533.83 Vibration Test
(a) [Each engine must undergo vibration surveys to establish that the vibration characteristics of those components that may be subject to mechanically or aerodynamically induced vibratory excitations are acceptable throughout the declared flight envelope. The engine surveys shall be based upon an appropriate combination of experience, analysis, and component test and shall address, as a minimum, blades, vanes, rotor discs, spacers, and rotor shafts.
(b) [The surveys shall cover the ranges of power or thrust, and both the physical and corrected rotational speeds for each rotor system, corresponding to operations throughout the range of ambient conditions in the declared flight envelope, from the minimum rotational speed up to 103 percent of the maximum physical and corrected rotational speed permitted for rating periods of two minutes or longer, and up to 100 percent of all other permitted physical and corrected rotational speeds, including those that are overspeeds. If there is any indication of a stress peak arising at the highest of those required physical or corrected rotational speeds, the surveys shall be extended sufficiently to reveal the maximum stress values present, except that the extension need not cover more than a further 2 percentage points increase beyond those speeds.
(c) [Evaluations shall be made of the following:
[(1) The effects on vibration characteristics of operating with scheduled changes (including tolerances) to variable vane angles, compressor bleeds, accessory loading, the most adverse inlet air flow distortion pattern declared by the manufacturer, and the most adverse conditions in the exhaust duct(s); and
[(2) The aerodynamic and aeromechanical factors which might induce or influence flutter in those systems susceptible to that form of vibration.
[(d) Except as provided by paragraph (e) of this section, the vibration stresses associated with the vibration characteristics determined under this section, when combined with the appropriate steady stresses, must be less than the endurance limits of the materials concerned, after making due allowances for operating conditions for the permitted variations in properties of the materials. The suitability of these stress margins must be justified for each part evaluated. If it is determined that certain operating conditions, or ranges, need to be limited, operating and installation limitations shall be established.
[(e) The effects on vibration characteristics of excitation forces caused by fault conditions (such as, but not limited to, out-of balance, local blockage or enlargement of stator vane passages, fuel nozzle blockage, incorrectly schedule compressor variables, etc.) shall be evaluated by test or analysis, or by reference to previous experience and shall be shown not to create a hazardous condition.
[(f) Compliance with this section shall be substantiated for each specific installation configuration that can affect the vibration characteristics of the engine. If these vibration effects cannot be fully investigated during engine certification, the methods by which they can be evaluated and methods by which compliance can be shown shall be substantiated and defined in the installation instructions required by 533.5.]
533.85 Calibration Tests
(a) Each engine must be subjected to those calibration tests necessary to establish its power characteristics and the conditions for the endurance test specified in 533.87. The results of the power characteristics of the engine over its entire operating range of speeds, pressures, temperatures, and altitudes. Power ratings are based upon standard atmospheric conditions with no airbleed for aircraft services and with only those accessories installed which are essential for engine functioning.
[(c) In showing compliance with this section, each condition must stabilize before measurements are taken, except as permitted by paragraph (d) of this section.
[(d) In the case of engines having 30-second OEI, and 2-minute OEI ratings, measurements taken during the applicable endurance test prescribed in 533.87(f) (1) through (8) may be used in showing compliance with the requirements of this section for these OEI ratings.]
533.87 Endurance Test
(a) [General. Each engine must be subjected to an endurance test that includes a total of at least 150 hours of operation and, depending upon the type and contemplated use of the engine, consists of one of the series of runs specified in paragraphs (b) through (g) of this section, as applicable. For engines tested under paragraphs (b), (c), (d), (e) or (g) of this section, the prescribed 6-hour test sequence must be conducted 25 times to complete the required 150 hours of operation. Engines for which the 30-second OEI and 2-minute OEI ratings are desired must be further tested under paragraph (f) of this section. The following test requirements apply:]
(1) The runs must be made in the order found appropriate by the Minister for the particular engine being tested.
(4) The runs must be made using fuel, lubricants and hydraulic fluid which conform to the specifications specified in complying with 533.7(c).
(5) Maximum air bleed for engine and aircraft services must be used during at least one‑fifth of the runs, except for the test required under paragraph (f) of this section, provided the validity of the test is not compromised. However, for these runs, the power or thrust or the rotor shaft rotational speed may be less than 100 percent of the value associated with the particular operation being tested if the Minister finds that the validity of the endurance test is not compromised.
(i) The load imposed by each accessory used only for aircraft service must be the limit load specified by the applicant for the engine drive and attachment point during rated maximum continuous power or thrust and higher output. (amended 2010/01/29; previous version)
(7) During the runs at any rated power or thrust the gas temperature and the oil inlet temperature must be maintained at the limiting temperature except where the test periods are not longer than 5 minutes and do not allow stabilisation. At least one run must be made with fuel, oil, and hydraulic fluid at the minimum pressure limit and at least one run must be made with fuel, oil, and hydraulic fluid at the maximum pressure limit with fluid temperature reduced as necessary to allow maximum pressure to be attained.
(8) [If the number of occurrences of either transient rotor shaft overspeed or transient gas overtemperature is limited, that number of the accelerations required by paragraphs (b) through (g) of this section must be made at the limiting overspeed or overtemperature. If the number of occurrences is not limited, half the required accelerations must be made at the limiting overspeed or overtemperature.]
(i) To change the thrust setting, the power control level must be moved from the initial position to the final position in not more than one second except for movements into the fuel burning thrust augmentor augmentation position if additional time to confirm ignition is necessary.
(ii) During the runs at any rated augmented thrust the hydraulic fluid temperature must be maintained at the limiting temperature except where the test periods are not long enough to allow stabilisation.
(iv) The endurance test must be conducted with the fuel burning thrust augmentor installed, with the primary and secondary exhaust nozzles installed, and with the variable area exhaust nozzles operated during each run according to the methods specified in complying with 533.5(b).
(b) Engines other than certain rotorcraft engines. For each engine, except a rotorcraft engine for which a rating is desired under paragraph (c), (d), or (e) of this section, the applicant must conduct the following runs:
(1) Take-off and idling. One hour of alternate 5-minute periods at rated take-off power and thrust and at idling power and thrust. The developed powers and thrusts at take-off and idling conditions and their corresponding rotor speed and gas temperature conditions must be as established by the power control in accordance with the schedule established by the manufacturer. The applicant may, during any one period, manually control the rotor speed, power, and thrust while taking data to check performance. For engines with augmented take-off power ratings that involve increase in turbine inlet temperature, rotor speed, or shaft power, this period of running at take-off must be at the augmented rating. For engines that do not materially increase operating severity, the amount of running conducted at the augmented rating is determined by the Minister. In changing the power setting after each period, the power-control lever must be moved in the manner prescribed in sub-paragraph (5) of this paragraph.
(2) Rated maximum continuous and take-off power and thrust. Thirty minutes at:
(ii) Rated take-off power and thrust during ten of the twenty-five 6-hour endurance test cycles.
(5) Acceleration and deceleration runs. Thirty minutes of accelerations and decelerations, consisting of 6 cycles from idling power and thrust to rated take-off power and thrust and maintained at the take-off power lever position for 30 seconds and at the idling power lever position for approximately 4 1/2 minutes. In complying with this subparagraph, the power-control lever must be moved from one extreme position to the other in not more than 1 second, except that, if different regimes of control operations are incorporated necessitating scheduling of the power-control lever motion in going from one extreme position to the other, a longer period of time is acceptable, but not more than 2 seconds.
(6) Starts. One hundred starts must be made, of which 25 starts must be preceded by at least a 2-hour engine shutdown. There must be at least 10 false engine starts, pausing for the applicant's specified minimum fuel drainage time, before attempting a normal start. There must be at least 10 normal restarts with not longer than 15 minutes since engine shutdown. The remaining starts may be made after completing the 150 hours of endurance testing.
(c) Rotorcraft engines for which a 30‑minute OEI power rating is desired. For each rotorcraft engine for which a 30‑minute OEI power rating is desired, the applicant must con duct the following series of tests:
(1) ) Take‑off and idling. One hour of alternate 5‑minute periods at rated take‑off power and at idling power. The developed powers at take‑off and idling conditions and their corresponding rotor speed and gas temperature conditions must be as established by the power control in accordance with the schedule established by the manufacturer. During any one period, the rotor speed and power may be controlled manually while taking data to check performance. For engines with augmented take‑off power ratings that involve increases in turbine inlet temperature, rotor speed, or shaft power, this period of running at rated take‑off power must be at the augmented power rating. In changing the power setting after each period, the power control lever must be moved in the manner prescribed in paragraph (c) (5) of this section.
(2) Rated maximum continuous and takeoff power. Thirty minutes at —
(i) Rated maximum continuous power during fifteen of the twenty-five 6-hour endurance test cycles; and (amended 2010/01/29; previous version)
(ii) Rated take-off power during ten of the twenty-five 6-hour endurance test cycles.
(6) Acceleration and deceleration runs. Thirty minutes of accelerations and decelerations, consisting of six cycles from idling power to rated take‑off power and maintained at the take‑off power lever position for 30 seconds and at the idling power lever position for approximately 4 1/2 minutes. In complying with this paragraph, the power control lever must be moved from one extreme position to the other in not more than 1 second, however, different regimes of control operations are incorporated that necessitate scheduling of the power control lever motion in going from one extreme position to the other, then a longer period of time is acceptable, but not more than 2 seconds.
(7) Starts. On hundred starts, of which 25 starts must be preceded by at least a 2‑hour engine shut‑down. There must be at least 10 false engine starts, pausing for the applicant’s specified minimum fuel drainage time, before attempting a normal start. There must be at least 10 normal restarts not more than 15 minutes after engine shutdown. The remaining starts may be made after completing the 150 hours of endurance testing.
(1) Take-off and idling. One hour of alternate 5-minute periods at rated take-off power and at idling power. The developed powers at take-off and idling conditions and their corresponding rotor speed and gas temperature conditions must be as established by the power control in accordance with the schedule established by the manufacturer. During any one period the rotor speed and power may be controlled manually while taking data to check performance. For engines with augmented take-off power ratings that involve increases in turbine inlet temperature, rotor speed, or shaft power, this period of running at rated take-off power must be at the augmented power rating. In changing the power setting after each period, the power control lever must be moved in the manner prescribed in paragraph (c)(5) of this section.
(2) Rated maximum continuous and take-off power. Thirty minutes at:
(6) Acceleration and deceleration runs. Thirty minutes of accelerations and decelerations, consisting of six cycles from idling power to rated take-off power and maintained at the take-off power lever position for 30 seconds and at the idling power lever position for approximately 4 1/2 minutes. In complying with this paragraph, the power control lever must be moved from one extreme position to the other in not more than 1 second, except that if different regimes of control operations are incorporated necessitating scheduling of the power control lever motion in going from one extreme position to the other, a longer period of time is acceptable, but not more than 2 seconds.
(e) Rotorcraft engines for which a 2 1/2 minute OEI power rating is desired. For each rotorcraft engine for which a 2 1/2 minute OEI power rating is desired, the applicant must conduct the following series of tests:
(1) Take-off, 2 1/2-minute OEI, and idling. One hour of alternate 5-minute periods at rated take-off power and at idling power except that, during the third and sixth take-off power periods, only 2 1/2 minutes need be conducted at rated take-off power, and the remaining 2 1/2 minutes must be conducted at rated 2 1/2-minute OEI power. The developed powers at take-off, 2 1/2-minute OEI, and idling conditions and their corresponding rotor speed and gas temperature conditions must be as established by the power control in accordance with the schedule established by the manufacturer. The applicant may, during any one period, control manually the rotor speed and power while taking data to check performance. For engines with augmented take-off power ratings that involve increases in turbine inlet temperature, rotor speed, or shaft power, this period of running at rated take-off power must be at the augmented rating. In changing the power setting after or during each period, the power control lever must be moved in the manner prescribed in paragraph (d)(6) of this section.
(2) The tests required in paragraphs (b)(2) through (b)(6), or (c)(2) through (c)(6), or (d)(2) through (d)(7) of this section, as applicable, except that in one of the 6-hour test sequences, the last 5 minutes of the 30 minutes at take-off power test period of paragraph (b)(2) of this section, or of the 30 minutes at 30-minute OEI power test period of paragraph (c)(2) of this section, or of the 1 hour at continuous OEI power test period of paragraph (d)(3) of this section, must be run at 2 1/2-minute OEI power.
(f) Rotorcraft engines for which 30‑second OEI and 2‑minute OEI ratings are desired. For each rotorcraft engine for which 30‑second OEI and 2‑minute OEI power ratings are desired, and following completion of the tests under paragraphs (b), (c), (d), or (e) of this section, the applicant may disassemble the tested engine to the extent necessary to show compliance with the requirements of 533.93(a). The tested engine must then be reassembled using the same parts used during the test runs of paragraphs (b), (c), (d), or (e) of this section, except those parts described as consumables in the Instructions for Continued Airworthiness. Additionally, the tests required in paragraphs (f)(1) through (f)(8) of this section must be run continuously. If a stop occurs during these tests, the interrupted sequence shall be repeated unless the applicant shows that the severity of the test would not be reduced if it were continued. The applicant must then conduct the following test sequence four times, for a total time of not less than 120 minutes:
(amended 2010/01/29; previous version) (1) Take‑off power. Three minutes at rated take‑off power.
(2) 30‑second OEI power. Thirty seconds at rated 30‑second OEI power.
(3) 2‑minute OEI power. Two minutes at rated 2‑minute OEI power.
(4) 30‑minute OEI power, continuous OEI power, or maximum continuous power. Five minutes at whichever is the greatest of rated 30‑minute OEI power, rated continuous OEI power, or rated maximum continuous power, whichever is greatest, except that, during the first test sequence, this period shall be 65 minutes. However, where the greatest rated power is 30-minute OEI power, that sixty-five minute period must consist of 30 minutes at 30-minute OEI power followed by 35 minutes at whichever is the greater of continuous OEI power or maximum continuous power.
(amended 2010/01/29; previous version) (5) 50 percent take‑off power. One minute at 50 percent take‑off power.
(6) 30‑second OEI power. Thirty seconds at rated 30‑second OEI power.
(7) 2‑minute OEI power. Two minutes at rated 2‑minute OEI power.
(amended 2010/01/29; previous version) [(g)] Supersonic aircraft engines. For each engine type certificated for use on supersonic aircraft the applicant must conduct the following:
(1) Subsonic test under sea level ambient atmospheric conditions. Thirty runs of one hour each must be made, consisting of:
(i) Two periods of 5 minutes at rated take-off augmented thrust each followed by 5 minutes at idle thrust;
(ii) One period of 5 minutes at rated take-off thrust followed by 5 minutes at not more than 15 percent of rated take-off thrust;
(iii) One period of 10 minutes at rated take-off augmented thrust followed by 2 minutes at idle thrust, except that if rated maximum continuous augmented thrust is lower than rated take-off augmented thrust, 5 of the 10-minute periods must be at rated maximum continuous augmented thrust; and
(iv) Six periods of 1 minute at rated take-off augmented thrust each followed by 2 minutes, including acceleration and deceleration time, at idle thrust.
(2) Simulated supersonic test. Each run of the simulated supersonic test must be preceded by changing the inlet air temperature and pressure from that attained at subsonic condition to the temperature and pressure attained at supersonic velocity, and must be followed by a return to the temperature attained at subsonic condition. Thirty runs of 4 hours each must be made, consisting of:
(i) One period of 30 minutes at the thrust obtained with the power control lever set at the position for rated maximum continuous augmented thrust followed by 10 minutes at the thrust obtained with the power control lever set at the position for 90 percent of rated maximum continuous augmented thrust. The end of this period in the first five runs must be made with the induction air temperature at the limiting condition of transient overtemperature, but need not be repeated during the periods specified in paragraphs (g)(2) (ii) through (iv) of this section;
(iii) One period repeating the run specified in paragraph (g)(2)(i) of this section, except that it must be followed by 10 minutes at the thrust obtained with the power control lever set at the position for 60 percent of rated maximum continuous augmented thrust and then 10 minutes at not more than 15 percent of rated take-off thrust;
(iv) One period repeating the runs specified in paragraph (g)(2)(i) and (ii) of this section; and
(v) One period of 30 minutes with 25 of the runs made at the thrust obtained with the power control lever set at the position rated maximum continuous augmented thrust, each followed by idle thrust and with the remaining 5 runs at the thrust obtained with the power control lever set at the position for rated maximum continuous augmented thrust for 25 minutes each, followed by subsonic operation at not more than 15 percent of rated take-off thrust and accelerated to rated take-off thrust for 5 minutes using hot fuel.
533.88 Engine overtemperature test
(a) Each engine must run for 5 minutes at maximum permissible r.p.m. with the gas temperature at least 75 (F (42 (C) higher than the maximum rating’s steady‑state operating limit, excluding maximum values of rpm and gas temperature associated with the 30‑second OEI and 2‑ minute OEI ratings. Following this run, the turbine assembly must be within serviceable limits.
(b) In addition to the test requirements in paragraph (a) of this section, each engine for which 30‑second OEI and 2‑minute OEI ratings are desired, that incorporates a means for automatic temperature control within its operating limitations in accordance with 533.67(d), must run for a period of 4 minutes at the maximum power‑on rpm with the gas temperature at least 35°F (19°C) higher than the maximum operating limit at 30-second OEI rating. Following this run, the turbine assembly may exhibit distress beyond the limits for an overtemperature condition provided the engine is shown by analysis or test, as found necessary by the Minister, to maintain the integrity of the turbine assembly.
(amended 2010/01/29; previous version) (c) A separate test vehicle may be used for each test condition.
533.89 Operation Test
(a) The operation test must include testing found necessary by the Minister to demonstrate:
(2) Compliance with the engine response requirements of 533.73; and
(3) The minimum power or thrust response time to 95% rated take-off power or thrust, from power lever positions representative of minimum idle and of minimum flight idle, starting from stabilised idle operation, under the following engine load conditions:
(4) If testing facilities are not available, the determination of power extraction required in paragraphs (a)(3)(ii) and (iii) of this section may be accomplished through appropriate analytical means.
(b) The operation test must include all testing found necessary by the Minister to demonstrate that the engine has safe operating characteristics throughout its specified operating envelope.
533.90 Initial Maintenance Inspection
Each engine, except engines being type certificated through amendment of an existing type approval or through supplemental type certification procedures, must undergo an approved test run that simulates the conditions in which the engine is expected to operate in service, including start-stop cycles, to establish when the initial maintenance inspection is required. The test run must be accomplished on an engine which substantially conforms to the final type design.
533.91 Engine System and Component Tests
(amended 2010/01/29; previous version) (a) For those systems or components that cannot be adequately substantiated in accordance with endurance testing of 533.87, the applicant must conduct additional tests to demonstrate that the systems or components are able to perform the intended functions in all declared environmental and operating conditions.
(c) Each unpressurised hydraulic fluid tank must not fail or leak when subjected to maximum operating temperature and an internal pressure of 5 p.s.i., and each pressurised hydraulic fluid tank must meet the requirements of 533.64.
(d) For an engine type certificated for use in supersonic aircraft, the systems, safety devices, and external components that may fail because of operation at maximum and minimum operating temperatures must be identified and tested at maximum and minimum operating temperatures and while temperature and other operation conditions are cycled between maximum and minimum operating values.
533.92 [Rotor Locking Tests
[If continued rotation is prevented by a means to lock the rotor(s), the engine must be subjected to a test that includes 25 operations of this means under the following conditions:
(a) [The engine must be shut down from rated maximum continuous thrust or power; and
(b) [The means for stopping and locking the rotor(s) must be operated as specified in the engine operating instructions while being subjected to the maximum torque that could result from continued flight in this condition; and
[(c) Following rotor locking, the rotor(s) must be held stationary under these conditions for five minutes for each of the 25 operations.]
533.93 Teardown Inspection
(a) [After completing the endurance testing of 533.87 (b), (c), (d), (e), or (g) of this Chapter, each engine must be completely disassembled, and
[(1) Each component having an adjustment setting and a functioning characteristic that can be established independent of installation on the engine must retain each setting and functioning characteristic within the limits that were established and recorded at the beginning of the test; and
[(2) Each engine part must conform to the type design and be eligible for incorporation into an engine for continued operation, in accordance with information submitted in compliance with 533.4.
(b) [After completing the endurance testing of 533.87(f), each engine must be completely disassembled, and
(1) [Each component having an adjustment setting and a functioning characteristic that can be established independent of installation on the engine must retain each setting and functioning characteristic within the limits that were established and recorded at the beginning of the test; and
(2) Each engine may exhibit deterioration in excess of that permitted in paragraph (a)(2) of this section including some engine parts or components that may be unsuitable for further use. The applicant must show by inspection, analysis, test, or by any combination thereof as found necessary by the Minister, that structural integrity of the engine is maintained; or
[(c) In lieu of compliance with paragraph (b) of this section, each engine for which the 30-second OEI and 2-minute OEI ratings are desired, may be subjected to the endurance testing of 533.87 (b), (c), (d), or (e) of this Chapter, and followed by the testing of 533.87(f) without intervening disassembly and inspection. However, the engine must comply with paragraph (a) of this section after completing the endurance testing of 533.87(f).]
533.94 Blade containment and rotor unbalance tests
(2) Failure of the most critical turbine blade while operating at maximum permissible r.p.m. The blade failure must occur at the outermost retention groove or, for integrally-bladed rotor discs, at least 80 percent of the blade must fail. The most critical turbine blade must be determined by considering turbine blade weight and strength of the adjacent turbine case at case temperatures and pressures associated with operation at maximum permissible r.p.m.
(b) Analysis based on rig testing component testing, or service experience may be substituted for one of the engine tests prescribed in paragraphs (a)(1) and (a)(2) of this section if:
533.95 Engine-Propeller Systems Tests
If the engine is designed to operate with a propeller, the following tests must be made with a representative propeller installed by either including the tests in the endurance run or otherwise performing them in a manner acceptable to the Minister:
533.96 Engine tests in auxiliary power unit (APU) mode
If the engine is designed with a propeller brake which will allow the propeller to be brought to a stop while the gas generator portion of the engine remains in operation, and remain stopped during operation of the engine as an auxiliary power unit ("APU mode"), in addition to the requirements of 533.87, the applicant must conduct the following tests:
(a) Ground locking: A total of 45 hours with propeller brake engaged in a manner which clearly demonstrates its ability to function without adverse effects on the complete engine while the engine is operating in the APU mode under the maximum conditions of engine speed, torque, temperature, air bleed, and power extraction as specified by the applicant.
(b) Dynamic braking: A total of 400 application-release cycles of brake engagements must be made in a manner which clearly demonstrates its ability to function without adverse effects on the complete engine under the maximum conditions of engine acceleration/deceleration rate, speed, torque and temperature as specified by the applicant. The propeller must be stopped prior to brake release.
(d) The tests required by paragraphs (a), (b) and (c) of this section must be performed on the same engine, but this engine need not be the same engine used for the tests required by 533.87.
(e) The tests required by paragraphs (a), (b) and (c) of this section must be followed by engine disassembly to the extent necessary to show compliance with the requirements of 533.93(a) and 533.93(b).
533.97 Thrust Reversers
(a) If the engine incorporates a reverser, the endurance, calibration, operation, and vibration tests prescribed in this subchapter must be run with the reverser installed. In complying with this section, the power control lever must be moved from one extreme position to the other in not more than 1 second except, if regimes of control operations are incorporated necessitating scheduling of the power control lever motion in going from one extreme position to the other, a longer period of time is acceptable but not more than 3 seconds. In addition, the test prescribed in paragraph (b) must be made. This test may be scheduled as part of the endurance run.
(b) One hundred seventy-five reversals must be made from flight-idle forward and 25 reversals must be made from rated take-off thrust to maximum reverse thrust. After each reversal the reverser must be operated at full reverse thrust for a period of 1 minute, except that, in the case of a reverser intended for use only as a braking means on the ground, the reverser need only be operated at full reverse thrust for 30-seconds.
533.99 General Conduct of Block Tests
(b) Each applicant may service and make minor repairs to the engine during the block tests in accordance with the service and maintenance instructions submitted in compliance with 533.4. If the frequency of the service is excessive, or the number of stops due to engine malfunction is excessive, or a major repair, or replacement of a part is found necessary during the block tests or as the result of findings from the teardown inspection, the engine or its parts must be subjected to any additional tests the Minister finds necessary.
Appendix A - Instructions for Continued Airworthiness Appendix B - Certification Standard Atmospheric Concentrations of Rain and Hail Date Modified: Top of Page