Airworthiness Manual Chapter 533 - Aircraft Engines - Canadian Aviation Regulations (CARs)

Interpretation Provision for Part V Standards

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.

Preamble

General

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.

The standards in this chapter are presented in a two column format with the United States FAR in the left column and the Canadian standards in the right column. Chapters, sub-chapters, sections and subsections numbering and headings are opposite to the equivalent FAR. Where the Canadian standard is identical to the FAR, the words “No Variation” appear; where a variation exists, the affected part of text is printed opposite to the FAR with all changes underlined.

AMA 533.90 entitled “Initial Maintenance Inspection” dated Jan. 1, 1987 is attached to this chapter.

Change 533-1

Effective: January 1, 1987

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.

Note:

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

Effective: January 1, 1989

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.

Second Edition

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:

Information Note:

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

Effective: April 7, 1997

- 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

Effective: April 7, 1997

- 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

Effective: October 29, 1998

- 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

Published: December 1, 2003

1. General

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

Effective: March 5, 2001

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2000-265	533.76
	533.77

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.

3. CARAC Proposed Amendment Recommendations

There are no proposed changes to the standard recommended by the CARAC Technical Committee Part V - Certification.

Change 533-7

Effective: December 1, 2004

In an effort to harmonize our regulatory guidance documents with those of other international aviation authorities and other branches within Transport Canada Civil Aviation (TCCA), the Aircraft Certification Branch has decided to replace existing Airworthiness Manual Advisories (AMA) related to certification of aeronautical products with new Advisory Circulars (AC). While the content of the new ACs will remain technically the same as the corresponding AMAs, which they will replace, the format of the ACs will be standardized to conform to other guidance documents published within the branch.

This change in guidance documentation becomes effective 1 December 2004 at which time the AMAs will be cancelled and replaced by their corresponding Advisory Circular concurrent with the next publishing of the Canadian Aviation Regulations (CAR). After this time, the CARAC Secretariat will no longer publish these AMAs and, consequently, ACs will not be published with their corresponding AWM Chapter. As of the 1 December 2004 issue of the CARs, any affected AMA references and content will have been removed. However, the AMA Index found in AMA 500/00 will, for now, continue to exist to provide a cross-reference between the old AMAs and the new ACs.

Change 533-8

Published December 30, 2006

Correction to English version

Effective: December 11, 2003

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2003-257	533.7

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

Effective: June 8, 2004

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2004-029	533.76

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

Published December 30, 2008

This change incorporates the following amendment to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33:

FAR Amendment 33-22

Effective: October 30, 2008

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2007-035	533.14 533.34 533.70

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

Effective: October 30, 2008

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2007-034	533.76

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

Effective: October 30, 2008

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2007-036	533.5 533.74 533.75 533.76

This amendment entitled “Safety Analysis” amends the safety analysis type certification standard for turbine aircraft engines.

Change 533-10

Published: December 1, 2009

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

Published: June 1, 2010

This Change incorporates the following amendments to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33:

FAR Amendment 33-26

Effective: January 29, 2010

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2009-008	533.5 533.7 533.27 533.28 533.29 533.53 533.67 533.88 533.91

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

Effective: January 29, 2010

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2009-009	533.64 533.71 533.91

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

Effective: January 29, 2010

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2009-011	533.5 533.29 533.67 533.87 533.88 533.93 Appendix A

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

Effective: January 29, 2010

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2009-014	533.19

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.

Change 533-12

Published: December 1, 2010

This change incorporates the following amendment to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33:

FAR Amendment 33-29

Effective: May 27, 2010

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2009-121	533.17

This amendment entitled “Fire Protection” amends the airworthiness standards for aircraft engines with respect to minimizing the probability and effect of a fire in and/or around the engine. This amendment harmonizes paragraph 533.17 of the Airworthiness Manual with the FAR part 33 and EASA CS-E standards.

FAR Amendment 33-30

Effective: May 27, 2010

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2009-119	533.7 533.84 533.87

This amendment entitled “Aircraft Engine Standards Overtorque Limits” establishes new engine overtorque testing requirements, amends engine ratings and operating limits and defines maximum overtorque for certain turbopropeller and turboshaft engines.

Change 533-13

Published: July 31, 2014

This Change incorporates the following amendments to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33:

FAR Amendment 33-21

Effective: January 21, 2013

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2012-017	533.71 533.90 533.201 Appendix A

This amendment entitled “Extended Operations of Multi-Engine Airplanes” amends the airworthiness standards for aeroplane engines and contains new and revised design rules pertaining to engines intended for extended operations and installed on transport category aeroplanes. This change harmonizes the Canadian design standards of AWM Chapter 533 for aeroplane engines seeking extended operations certification with those of FAR Part 33.

Note that this is an optional requirement for design approval for engines intended for extended operations. Operations criteria for the acceptance of engines and aeroplanes into an extended operations role are currently contained in the operating requirements of the Safety Criteria for Approval of Extended Range Twin-Engine Operations (ETOPS) Manual, TP 6327. Moving the design requirements into the Airworthiness Manual places them alongside other similar design standards, thereby allowing consistent management of ETOPS issues (such as establishment and maintenance of a certification basis).

Please note that a further NPA to refine terminology introduced in NPA 2012-017 has been developed. The NPA entitled “Correction of Terms” will be published shortly under “Change 533-21 Correction.”

FAR Amendment 33-26, Correction

Effective: January 28, 2013

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2012-016	533.28

The purpose of this amendment is to correct the title of section 533.28 to align with the revised section title of 14 CFR 33.28 as amended by Amendment. 33-26. The revised title of section 533.28 was inadvertently not incorporated into AWM 533 as part of the amendments introduced by NPA 2009-008.

Specifically, the title of AWM 533.28 is now corrected to read “Engine Control Systems” instead of “Electrical and Electronic Engine Control Systems.”

FAR Amendment 33-31

Effective: December 18, 2012

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2012-003	533.27

This amendment entitled “Airworthiness Standards; Rotor Overspeed Requirements” amends the airworthiness standards for aircraft engines with respect to the aircraft turbine engine rotor overspeed type certification standards. FAA Amendment 33-31 establishes uniform rotor overspeed design and test requirements for aircraft engines and turbochargers certificated by the FAA and the European Aviation Safety Agency (EASA); harmonization of Canadian standards with the now common FAA – EASA standard eliminates the need to comply with differing sets of requirements for different jurisdictions.

Subsequent to publishing Amendment 33-31, the FAA published a Correction to Amendment 33-31. This correction changed some of the language within the preamble/explanation of the final rule and was intended to ensure correct understanding of and was the FAA’s response to comments received. There were no changes to the wording of the design standard; thus, the correction did not need to be adopted into the Airworthiness Manual to maintain harmonization of the design standards.

FAR Amendment 33-32

Effective: January 28, 2013

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2012-004	533.87

This amendment entitled ‘Technical Amendment; Airworthiness Standards—Aircraft Engines’ revises the airworthiness standards of FAR 33.87 as a result of a number of inadvertent editing errors in the airworthiness standards for aircraft engine endurance tests. None of the corrections are substantive in nature, and none will impose any additional burden on any person.

Information Note regarding FAR Amendment 33-33:

FAR Amendment 33-33, published in the U.S. Federal Register on July 5, 2012 (FR 39623) was intended to clarify the engine vibration test requirements and revise the regulations to clarify that engine surveys require an engine test. After further consideration, the FAA published a subsequent correction to this Amendment in the U.S. Federal Register on September 20, 2012 (FR 58301) to reverse the changes to §33.83(a) and restore its previous wording. The FAR Amendment was effectively cancelled by this correction. Neither FAR Amendment 33-33 nor the correction to Amendment 33-33 was adopted into the Airworthiness Manual. AWM 533.83(a), which was last revised at Change 533-5, remains harmonized with FAR 33.83(a) at Amendment 33-17, which is its current status. As the FAR amendment is not being adopted, there will be no Change 533-33 to the AWM.

Change 533-14

Published: October 3, 2019

This Change introduces an editorial correction to section 533.201 of Chapter 533 of the Airworthiness Manual (AWM):

Correction to FAR Amendment 33-21

Effective: August 25, 2019

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2019-012	533.201

This amendment entitled Extended Operations (ETOPS) of Multi-Engine Aeroplanes amends the design standards of Chapter 533 of the Airworthiness Manual (AWM) by an editorial correction of section 533.201 corresponding to the amendments to AWM Chapter 525. Note that AWM Chapter 533 – Aircraft Engines [Change 533-13] has already incorporated the related Amendment 33-21 to 14 CFR part 33 through NPA 2012-017.

This amendment revises section 533.201 of AWM Chapter 533 to use the term “extended operations (ETOPS)” rather than “Extended Range Twin-Engine Operations (ETOPS)” in order to use the same term as in AWM Chapter 525.

Change 533-15

Published: October 16, 2019

This Change introduces correction to section 533.28 of Chapter 533 of the Airworthiness Manual (AWM):

Correction to section 533.28

Effective: September 30, 2019

Table of Change Information
Notice of Proposed Amendment	Amended Section(s)
2019-015	533.28

This amendment entitled “Engine Control Systems” amends the Airworthiness Manual (AWM) Chapter 533 to make a correction to section 533.28 “Engine Control Systems” as it was previously amended through NPA 2009-008. NPA 2009-008 contained an error in the title of section 533.28 and omitted to delete the preamble statement to AWM 533.28. Also, certain required punctuation was omitted in the English version. NPA 2012-016, effective January 28, 2013, was subsequently issued to correct the title, but did not make further corrections.

This change deletes the preamble statement to AWM 533.28, and adds the required punctuation in the English version such as to harmonize with the source code equivalent, 14 CFR 33.28.

Minor language corrections are also made in the French version.

Change 533-16

Published: July 29, 2021

This Change incorporates the following amendment to the United States Code of Federal Regulations, Title 14, Chapter I, Part 33:

FAR Amendment 33-34

Effective: April 08, 2021

Table of Change Information
Notice(s) of Proposed Amendment	Amended Section(s)
2020-017	533.68 533.77 Appendix C Appendix D

This amendment entitled “Aeroplane and Engine Certification Requirements in Supercooled Large Drop, Mixed Phase, and Ice Crystal Icing Conditions” amends the design standards of Airworthiness Manual (AWM) Chapter 533 by incorporating the requirements of Federal Aviation Regulations (FAR), Part 33, Amendment No. 33-34 as published on November 4, 2014 in the U.S. Federal Register, Volume 79, No. 213 [79 FR 65508].

This amendment revises the icing airworthiness standards applicable to certain aircraft engines. These standards would improve safety by addressing supercooled large drop (SLD), mixed phase, and ice crystal icing (ICI) conditions for all turbojet, turbofan and turboprop engines.

Subchapter A - General

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)
(b) Reserved.
(amended 2009/12/01)

(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)

(a) Installation instructions.
1. (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. (2) The location and description of engine connections to be attached to accessories, pipes, wires, cables, ducts and cowling.
3. (3) An outline drawing of the engine including overall dimensions.
4. (4) A definition of the physical and functional interfaces with the aircraft and aircraft equipment, including the propeller when applicable.
  (amended 2010/01/29)
5. (5) Where an engine system relies on components that are not part of the engine type design, the interface conditions and reliability requirements for those components upon which engine type certification is based must be specified in the engine installation instructions directly or by reference to appropriate documentation.
  (amended 2010/01/29)
6. (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.
  (amended 2010/01/29)
(b) Operating instructions.
1. (1) The operating limitations established by the Minister.
2. (2) The power or thrust ratings and procedures for correcting for non-standard atmosphere.
3. (3) The recommended procedures, under normal and extreme ambient conditions for:
  1. (i) Starting;
  2. (ii) Operating on the ground; and
  3. (iii) Operating during flight.
4. (4) For rotorcraft engines having one or more OEI ratings, applicants must provide data on engine performance characteristics and variability to enable the aircraft manufacturer to establish aircraft power assurance procedures.
  (amended 2010/01/29)
5. (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.
  (amended 2010/01/29)
(c) Safety analysis assumptions
(amended 2008/10/30; no previous version)

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.
(amended 2008/10/30; no previous version)

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.
(b) For reciprocating engines, ratings and operating limitations are established relating to the following:
1. (1) Horsepower or torque, r.p.m., manifold pressure, and time at critical pressure altitude and sea level pressure altitude for:
  1. (i) Rated maximum continuous power (relating to unsupercharged operation or to operation in each supercharger mode as applicable); and
  2. (ii) Rated take-off power (relating to unsupercharged operation or to operation in each supercharger mode as applicable).
2. (2) Fuel grade or specification.
3. (3) Oil grade or specification.
4. (4) Temperature of the:
  1. (i) Cylinder;
  2. (ii) Oil at the oil inlet; and
  3. (iii) Turbo-supercharger turbine wheel inlet gas.
5. (5) Pressure of:
  1. (i) Fuel at the fuel inlet; and
  2. (ii) Oil at the main oil gallery.
6. (6) Accessory drive torque and overhang moment.
7. (7) Component life.
8. (8) Turbosupercharger turbine wheel r.p.m.
(c) For turbine engines, ratings and operating limitations are established relating to the following:
1. (1) Horsepower, torque, or thrust, r.p.m., gas temperature, and time for:
  1. (i) Rated maximum continuous power or thrust (augmented);
  2. (ii) Rated maximum continuous power or thrust (unaugmented);
  3. (iii) Rated take-off power or thrust (augmented);
  4. (iv) Rated take-off power or thrust (unaugmented);
  5. (v) Rated 30-minute OEI power;
  6. (vi) Rated 2 1/2-minute OEI power;
  7. (vii) Rated continuous OEI power;
  8. (viii) Rated 2-minute OEI power;
  9. (ix) Rated 30-second OEI power; and
  10. (x) Auxiliary power unit (APU) mode of operation.
2. (2) Fuel designation or specification.
3. (3) Oil grade or specification.
4. (4) Hydraulic fluid specification.
5. (5)Temperature of:
  1. (i) Oil at a location specified by the applicant;
  2. (ii) Induction air at the inlet face of a supersonic engine, including steady state operation and transient over-temperature and time allowed;
  3. (iii) Hydraulic fluid of a supersonic engine;
  4. (iv) Fuel at a location specified by the applicant; and
  5. (v) External surfaces of the engine, if specified by the applicant.
    (amended 2003/12/11; no previous version)
6. (6) Pressure of:
  1. (i) Fuel at the fuel inlet;
  2. (ii) Oil at a location specified by the applicant;
  3. (iii) Induction air at the inlet face of a supersonic engine, including steady state operation and transient overpressure and time allowed; and
  4. (iv) Hydraulic fluid.
7. (7) Accessory drive torque and overhang movement.
8. (8) Component life.
9. (9) Fuel filtration.
10. (10) Oil filtration.
11. (11) Bleed air.
12. (12) The number of start-stop stress cycles approved for each rotor disc and spacer.
13. (13) Inlet air distortion at the engine inlet.
14. (14) Transient rotor shaft overspeed r.p.m. and number of overspeed occurrences.
15. (15) Transient gas overtemperature and number of overtemperature occurrences.
16. (16) Transient engine overtorque and number of overtorque occurrences.
  (amended 2010/05/27)
17. (17) Maximum engine overtorque for turbopropeller and turboshaft engines incorporating free power turbines.
  (amended 2010/05/27; no previous version)
(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.
(amended 2010/01/29)
(Change. 533-1 (87-01-01))

(Change 533-2 (89-01-01))

(Change 533-5)

533.8 Selection of Engine Power and Thrust Ratings

(a) Requested engine power and thrust ratings must be selected by the applicant.

(b) Each selected rating must be for the lowest power or thrust that all engines of the same type may be expected to produce under the conditions used to determine that rating.

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)

(amended 2008/10/30)

533.15 Materials

The suitability and durability of materials used in the engine must:

(a) Be established on the basis of experience or tests; and
(b) Conform to approved specifications (such as industry or military specifications) that ensure their having the strength and other properties assumed in the design data.

533.17 Fire Protection

(amended 2010/05/27)

(a) The design and construction of the engine and the materials used must minimise the probability of the occurrence and spread of fire during normal operation and failure conditions, and must minimise the effect of such a 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 or other hazardous conditions.
(amended 2010/05/27)
(b) Except as provided in paragraph (c) of this section, each external line, fitting and other component, which contains or conveys flammable fluid during normal engine operation, must be fire resistant or fireproof, as determined by the Minister. Components must be shielded or located to safeguard against the ignition of leaking flammable fluid.
(amended 2010/05/27)
(c) A tank, which contains flammable fluid and any associated shut-off means and supports, which are part of and attached to the engine must be fireproof either by construction or by protection unless damage by fire will not cause leakage or spillage of a hazardous quantity of flammable fluid. For a reciprocating engine having an integral oil sump of less than 23.7 litres (25-quart) capacity, the oil sump need not be fireproof or enclosed by a fireproof shield.
(amended 2010/05/27)
(d) An engine component designed, constructed and installed to act as a firewall must be:
(amended 2010/05/27)
1. (1) Fireproof;
2. (2) Constructed so that no hazardous quantity of air, fluid or flame can pass around or through the firewall; and,
3. (3) Protected against corrosion;
(e) In addition to the requirements of paragraphs (a) and (b) of this section, engine control system components that are located in a designated fire zone must be fire resistant or fireproof, as determined by the Minister.
(amended 2010/05/27)
(f) Unintentional accumulation of hazardous quantities of flammable fluid within the engine must be prevented by draining and venting.
(amended 2010/05/27)
(g) Any components, modules, or equipment, which are susceptible to, or are potential sources of static discharges or electrical fault currents must be designed and constructed to be properly grounded to the engine reference, to minimise the risk of ignition in external areas where flammable fluids or vapours could be present.
(amended 2010/05/27; no previous version)

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)

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:
1. (1) The specified limit loads without permanent deformation; and
2. (2) The specified ultimate loads without failure, but may exhibit permanent deformation.

533.25 Accessory Attachments

The engine must operate properly with the accessory drive and mounting attachments loaded. Each engine accessory drive and mounting attachment must include provisions for sealing to prevent contamination of, or unacceptable leakage from, the engine interior. A drive and mounting attachment requiring lubrication for external drive splines, or coupling by engine oil, must include provisions for sealing to prevent unacceptable loss of oil and to prevent contamination from sources outside the chamber enclosing the drive connection. The design of the engine must allow for the examination, adjustment, or removal of each accessory required for engine operation.

533.27 Turbine, Compressor, Fan and Turbosupercharger Rotors Overspeed

(effective 2012/12/18)

(a) For each turbine, fan, compressor, and turbosupercharger rotor, the applicant must establish by test, analysis, or a combination of both, that each rotor will not burst when operated in the engine for 5 minutes at whichever of the conditions defined in paragraph (b) of this section is the most critical with respect to the integrity of such a rotor.
1. (1) Test rotors used to demonstrate compliance with this section that do not have the most adverse combination of material properties and dimensional tolerances must be tested at conditions which have been adjusted to ensure the minimum specification rotor possesses the required overspeed capability. This can be accomplished by increasing test speed, temperature, and/or loads.
2. (2) When an engine test is being used to demonstrate compliance with the overspeed conditions listed in paragraph (b)(3) or (b)(4) of this section and the failure of a component or system is sudden and transient, it may not be possible to operate the engine for 5 minutes after the failure. Under these circumstances, the actual overspeed duration is acceptable if the required maximum overspeed is achieved.
(b) When determining the maximum overspeed condition applicable to each rotor in order to comply with paragraphs (a) and (c) of this section, the applicant must evaluate the following rotor speeds taking into consideration the part’s operating temperatures and temperature gradients throughout the engine’s operating envelope:
1. (1) 120 percent of the maximum permissible rotor speed associated with any of the engine ratings except one engine- inoperative (OEI) ratings of less than 2½ minutes;
2. (2) 115 percent of the maximum permissible rotor speed associated with any OEI ratings of less than 2½ minutes;
3. (3) 105 percent of the highest rotor speed that would result from either:
  1. (i) the failure of the component or system which, in a representative installation of the engine, is the most critical with respect to overspeed when operating at any rating condition except OEI ratings of less than 2½ minutes, or
  2. (ii) the failure of any component or system which in a representative installation of the engine, in combination with any other failure of a component or system that would not normally be detected during a routine pre-flight check or during normal flight operation, that is the most critical with respect to overspeed, except as provided by paragraph (c) of this section, when operating at any rating condition except OEI ratings of less than 2½ minutes; and
4. (4) 100 percent of the highest rotor speed that would result from the failure of the component or system which, in a representative installation of the engine, is the most critical with respect to overspeed when operating at any OEI rating of less than 2½ minutes.
(c) The highest overspeed that results from a complete loss of load on a turbine rotor, except as provided by paragraph (f) of this section, must be included in the overspeed conditions considered by paragraphs (b)(3)(i), (b)(3)(ii), and (b)(4) of this section, regardless of whether that overspeed results from a failure within the engine or external to the engine. The overspeed resulting from any other single failure must be considered when selecting the most limiting overspeed conditions applicable to each rotor. Overspeeds resulting from combinations of failures must also be considered unless the applicant can show that the probability of occurrence is not greater than extremely remote (probability range of 10^-7 to 10^-9 per engine flight hour).
(d) In addition, the applicant must demonstrate that each fan, compressor, turbine, and turbosupercharger rotor complies with paragraphs (d)(1) and (d)(2) of this section for the maximum overspeed achieved when subjected to the conditions specified in paragraphs (b)(3) and (b)(4) of this section. The applicant must use the approach in paragraph (a) of this section which specifies the required test conditions.
1. (1) Rotor growth must not cause the engine to:
  1. (i) catch fire,
  2. (ii) release high-energy debris through the engine casing or result in a hazardous failure of the engine casing,
  3. (iii) generate loads greater than those ultimate loads specified in 533.23(a), or
  4. (iv) lose the capability of being shut down.
2. (2) Following an overspeed event and after continued operation, the rotor may not exhibit conditions such as cracking or distortion which preclude continued safe operation.
(e) The design and functioning of engine control systems, instruments, and other methods not covered under 533.28 must ensure that the engine operating limitations that affect turbine, compressor, fan and turbosupercharger rotor structural integrity will not be exceeded in service.
(f) Failure of a shaft section may be excluded from consideration in determining the highest overspeed that would result from a complete loss of load on a turbine rotor if the applicant:
1. (1) identifies the shaft as an engine life-limited-part and complies with 533.70;
2. (2) uses material and design features that are well understood and that can be analyzed by well-established and validated stress analysis techniques;
3. (3) determines, based on an assessment of the environment surrounding the shaft section, that environmental influences are unlikely to cause a shaft failure. This assessment must include complexity of design, corrosion, wear, vibration, fire, contact with adjacent components or structure, overheating and secondary effects from other failures or combination of failures;
4. (4) identifies and declares, in accordance with 533.5, any assumptions regarding the engine installation in making the assessment described in paragraph (f)(3) of this section;
5. (5) assesses and considers as appropriate, experience with shaft sections of similar design;
6. (6) does not exclude the entire shaft.
(g) If analysis is used to meet the overspeed requirements, then the analytical tool must be validated to prior overspeed test results of a similar rotor. The tool must be validated for each material. The rotor being certified must not exceed the boundaries of the rotors being used to validate the analytical tool in terms of geometric shape, operating stress, and temperature. Validation includes the ability to accurately predict rotor dimensional growth and the burst speed. The predictions must also show that the rotor being certified does not have lower burst and growth margins than rotors used to validate the tool.

533.28 Engine Control Systems

(effective 2013/01/28)

(Preamble statement deleted 2019/09/30)

(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.
(effective 2019/09/30)
(b) Validation.
(amended 2010/01/29)
1. (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:
  (amended 2010/01/29)
  1. (i) Enables selected values of relevant control parameters to be maintained and the engine kept within the approved operating limits over changing atmospheric conditions in the declared flight envelope;
    (amended 2010/01/29)
  2. (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;
    (amended 2010/01/29)
  3. (iii) Allows modulation of engine power or thrust with adequate sensitivity over the declared range of engine operating conditions; and
    (amended 2010/01/29)
  4. (iv) Does not create unacceptable power or thrust oscillations.
    (amended 2010/01/29)
2. (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.
  (amended 2010/01/29)
(c) Control transitions.
(amended 2010/01/29)
1. (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)
  1. (i) The engine does not exceed any of its operating limitations;
    (amended 2010/01/29)
  2. (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)
  3. (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)
2. (2) The magnitude of any change in thrust or power and the associated transition time must be identified and described in the engine installation instructions and the engine operating instructions.
  (amended 2010/01/29)
(d) Engine control system failures. The applicant must design and construct the engine control system so that:
(amended 2010/01/29)
1. (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)
2. (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)
3. (3) Single failures of engine control system components do not result in a hazardous engine effect; and
  (amended 2010/01/29)
4. (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)
(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.
(amended 2010/01/29)
(f) Protection systems
(amended 2010/01/29)
1. (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)
2. (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)
3. (3) When overspeed protection is provided through hydromechanical or mechanical means, the applicant must demonstrate by test or other acceptable means that the overspeed function remains available between inspection and maintenance periods.
  (amended 2010/01/29)
(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)
(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)
1. (1) Not result in a hazardous engine effect for any engine; and
  (amended 2010/01/29)
2. (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)
(i) Aircraft-supplied electrical power
(amended 2010/01/29)
1. (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)
  1. (i) A hazardous engine effect; or
    (amended 2010/01/29)
  2. (ii) The unacceptable transmission of erroneous data.
    (amended 2010/01/29)
2. (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)
3. (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)
4. (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)
(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)
(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)
(l) Engine shut down means. Means must be provided for shutting down the engine rapidly.
(amended 2010/01/29)
(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.
(amended 2010/01/29)
(Change 533-4 (93-12-30))

533.29 Instrument Connection

(a) Unless it is constructed to prevent its connection to an incorrect instrument, each connection provided for powerplant instruments required by aircraft airworthiness regulations or necessary to insure operation of the engine in compliance with any engine limitation must be marked to identify it with its corresponding instrument.
(b) A connection must be provided on each turbojet engine for an indicator system to indicate rotor system unbalance.
(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)
1. (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;
2. (2) Automatically record each usage and duration of power at the 30-second OEI and 2-minute OEI levels;
  (amended 2010/01/29)
3. (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)
4. (4) Enable routine verification of the proper operation of the above means.
  (amended 2010/01/29)
(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)
(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)
(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)
(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)
(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)
1. (1) Other existing instrumentation provides adequate warning of failure or impending failure;
  (amended 2010/01/29)
2. (2) Failure of the cooling system would not lead to hazardous engine effects before detection; or
  (amended 2010/01/29)
3. (3) The probability of failure of the cooling system is extremely remote.
  (amended 2010/01/29)
  (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

(amended 2008/10/30; no previous version)

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.
(amended 2008/10/30; no previous version)

533.35 Fuel &Induction System

(a) The fuel system of the engine must be designed and constructed to supply an appropriate mixture of fuel to the cylinders throughout the complete operating range of the engine under all flight and atmospheric conditions.
(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.
(c) The type and degree of fuel filtering necessary for protection of the engine fuel system against foreign particles in the fuel must be specified. The applicant must show that foreign particles passing through the prescribed filtering means will not critically impair engine fuel system functioning.
(d) Each passage in the induction system that conducts a mixture of fuel and air must be self-draining, to prevent a liquid lock in the cylinders, in all attitudes that the applicant establishes as those the engine can have when the aircraft in which it is installed is in the static ground attitude.
(e) If provided as part of the engine, the applicant must show for each fluid injection (other than fuel) system and its controls that the flow of the injected fluid is adequately controlled.

533.37 Ignition System

Each spark ignition engine must have a dual ignition system with at least two spark plugs for each cylinder and two separate electric circuits with separate sources of electrical energy, or have an ignition system of equivalent in-flight reliability.

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.
(b) The lubrication system of the engine must be designed and constructed to allow installing a means of cooling the lubricant.
(c) The crankcase must be vented to the atmosphere to preclude leakage of oil from excessive pressure in the crankcase.

Subchapter D - Block Tests Reciprocating Aircaft 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.

(Change 533-1 (87-01-01))

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.
(b) The torsional and bending vibration stresses of the crankshaft and the propeller shaft or other output shaft may not exceed the endurance limit stress of the material from which the shaft is made. If the maximum stress in the shaft cannot be shown to be below the endurance limit by measurement, the vibration frequency and amplitude must be measured. The peak amplitude must be shown to produce a stress below the endurance limit; if not, the engine must be run at the condition producing the peak amplitude until, for steel shafts, 10 million stress reversals have been sustained without fatigue failure and, for other shafts, until it is shown that fatigue will not occur within the endurance limit stress of the material.
(c) Each accessory drive and mounting attachment must be loaded, with the loads imposed by each accessory used only for an aircraft service being the limit load specified by the applicant for the drive or attachment point.
(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.
(b) A power check at sea level conditions must be accomplished on the endurance test engine after the endurance test. Any change in power characteristics which occurs during the endurance test must be determined. Measurements taken during the final portion of the endurance test may be used in showing compliance with the requirements of this paragraph.

533.47 Detonation Test

Each engine must be tested to establish that the engine can function without detonation throughout its range of intended conditions of operation.

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. (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.
2. (2) 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 75 percent rated maximum continuous power and 91 percent maximum continuous speed.
3. (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.
4. (4) 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 65 percent rated maximum continuous power and 87 percent maximum continuous speed.
5. (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.
6. (6) 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 50 percent rated maximum continuous power and 79.5 percent maximum continuous speed.
7. (7) A 20-hour run consisting of alternate periods of 2 1/2 hours at rated maximum continuous power with maximum continuous speed, and 2 1/2 hours at maximum best economy cruising power or at maximum recommended cruising 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. (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. (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. (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.
4. (4) A 30-hour run in the higher gear ratio at rated maximum continuous power with maximum continuous speed.
5. (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. (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. (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. (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. (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. (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.
  Where simulated altitude test equipment is not available when operating in the higher gear ratio, the runs may be made at the observed horsepower obtained with the critical altitude manifold pressure or specified percentages thereof, and the fuel-air mixtures may be adjusted to be rich enough to suppress detonation.
(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. (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.
2. (2) A 25-hour run consisting of alternate periods of 2 1/2-hours each at rated maximum continuous power with maximum continuous speed, and at 70 percent rated maximum continuous power with maximum continuous speed.
3. (3) A 25-hour run consisting of alternate periods of 2 1/2-hours each at rated maximum continuous power with maximum continuous speed, and at 70 percent rated maximum continuous power with 80 to 90 percent maximum continuous speed.
4. (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. (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. (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. (1) For engines used in aeroplanes the applicant must conduct the runs specified in paragraph (b) of this section, except:
  1. (i) The entire run specified in subparagraph (b)(1) of this section must be made at sea level altitude pressure;
  2. (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
  3. (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. (2) For engines used in helicopters and applicant must conduct the runs specified in paragraph (d) of this section, except:
  1. (i) The entire run specified in paragraph (d)(1) of this section must be made at critical altitude pressure;
  2. (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;
  3. (iii) The entire run specified in paragraph (d)(4) of this section must be made at 8,000 feet altitude pressure;
  4. (iv) The portion of the runs specified in paragraph (d)(5) of this section at 80 percent of rated maximum continuous power must be made at 8,000 feet altitude pressure and the portions of the runs at other power must be made at critical altitude pressure;
  5. (v) The entire run specified in paragraph (d)(6) of this section must be made at critical altitude pressure; and
  6. (vi) The turbosupercharger used during the endurance test must be run on the bench for 50 hours at the limiting turbine wheel inlet gas temperature and rotational speed for rated maximum continuous power operation unless the limiting temperature and speed are maintained during 50 hours of the rated maximum continuous power operation.
    (Change 533-1 (87-01-01))

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

(amended 2010/01/29)

(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.
(amended 2010/01/29)
(b) Temperature limits must be established for each component that requires temperature controlling provisions in the aircraft installation to assure satisfactory functioning, reliability, and durability.

533.55 Teardown Inspection

After completing the endurance test:

(a) Each engine must be completely disassembled;
(b) 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
(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.
(Change 533-1 (87-01-01))

533.57 General Conduct of Block Tests

(a) The applicant may, in conducting the block tests, use separate engines of identical design and construction in the vibration, calibration, detonation, endurance, and operation tests, except that, if a separate engine is used for the endurance test it must be subjected to a calibration check before starting the endurance test.
(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.
(c) Each applicant must furnish all testing facilities, including equipment and competent personnel, to conduct the block tests.

Subchapter E - Design and Construction Turbine Aircraft Engines

533.61 Applicability

This subchapter prescribes additional design and construction requirements for turbine aircraft engines.

(Change 533-1 (87-01-01))

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.

(Change 533-5)

533.64 Pressurized Engine Static Parts

(amended 2010/01/29; no previous version)

(a) Strength. The applicant must establish by test, validated analysis, or a combination of both, that all static parts subject to significant gas or liquid pressure loads for a stabilized period of one minute will not:
(amended 2010/01/29; no previous version)
1. (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)
  1. (i) 1.1 times the maximum working pressure;
    (amended 2010/01/29; no previous version)
  2. (ii) 1.33 times the normal working pressure; or
    (amended 2010/01/29; no previous version)
  3. (iii) 35 kPa (5 p.s.i.) above the normal working pressure.
    (amended 2010/01/29; no previous version)
2. (2) Exhibit fracture or burst when subjected to the greater of the following pressures:
  (amended 2010/01/29; no previous version)
3. (i) 1.15 times the maximum possible pressure;
  (amended 2010/01/29; no previous version)
4. (ii) 1.5 times the maximum working pressure; or
  (amended 2010/01/29; no previous version)
5. (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. (1) The operating temperature of the part;
  (amended 2010/01/29; no previous version)
2. (2) Any other significant static loads in addition to pressure loads;
  (amended 2010/01/29; no previous version)
3. (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)
4. (4) Any adverse geometry conditions allowed by the type design.
  (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:
1. (1) It must be accessible for draining and cleaning and must incorporate a screen or element that is easily removable.
2. (2) It must have a sediment trap and drain except that it need not have a drain if the strainer or filter is easily removable for drain purposes.
3. (3) It must be mounted so that its weight is not supported by the connecting lines or by the inlet or outlet connections of the strainer or filter, unless adequate strength margins under all loading conditions are provided in the lines and connections.
4. (4) It must have the type and degree of fuel filtering specified as necessary for protection of the engine fuel system against foreign particles in the fuel. The applicant must show:
  1. (i) That foreign particles passing through the specified filtering means do not impair the engine fuel system functioning; and
  2. (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. (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:
  1. (i) Existing engine instrumentation; or
  2. (ii) Additional means incorporated into the engine fuel system.
6. (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.
(c) If provided as part of the engine, the applicant must show for each fluid injection (other than fuel) system and its controls that the flow of the injected fluid is adequately controlled.
(d) [Repealed]
(effective 2010/01/29)
(Change 533-5)

533.68 Induction System Icing

Each engine, with all icing protection systems operating, must:

(a) Operate throughout its flight power range including the minimum descent idle rotor speeds achievable in flight, in the icing conditions defined for turbojet, turbofan, and turboprop engines in Appendices C and O of Chapter 525 of this manual, and Appendix D of this chapter, and for turboshaft engines in Appendix C of Chapter 529 of this manual, without the accumulation of ice on the engine components that:
1. (1) Adversely affects engine operation or that causes an unacceptable permanent loss of power or thrust or unacceptable increase in engine operating temperature; or
2. (2) Results in unacceptable temporary power loss or engine damage; or
3. (3) Causes a stall, surge, or flameout or loss of engine controllability. The applicant must account for in-flight ram effects in any critical point analysis or test demonstration of these flight conditions. (effective 2021/04/08)
(b) Operate throughout its flight power range, including minimum descent idle rotor speeds achievable in flight, in the icing conditions defined for turbojet, turbofan, and turboprop engines in Appendices C and O of Chapter 525 of this manual, and for turboshaft engines in Appendix C of Chapter 529 of this manual. In addition:
1. (1) It must be shown through Critical Point Analysis (CPA) that the complete ice envelope has been analyzed, and that the most critical points shall be demonstrated by engine test, analysis, or a combination of the two to operate acceptably. Extended flight in critical flight conditions such as hold, descent, approach, climb, and cruise, shall be addressed, for the ice conditions defined in these appendices.
  (effective 2021/04/08)
2. (2) It must be shown by engine test, analysis, or a combination of the two that the engine can operate acceptably for the following durations:
  (effective 2021/04/08)
  1. (i) At engine powers that can sustain level flight: A duration that achieves repetitive, stabilized operation for turbojet, turbofan, and turboprop engines in the icing conditions defined in Appendices C and O of Chapter 525 of this manual, and for turboshaft engines in the icing conditions defined in Appendix C of Chapter 529 of this chapter.
  2. (ii) At engine power below that which can sustain level flight:
    1. (A) Demonstration in altitude flight simulation test facility: A duration of 10 minutes consistent with a simulated flight descent of 10,000 ft (3 km) in altitude while operating in Continuous Maximum icing conditions defined in Appendix C of Chapter 525 of this manual for turbojet, turbofan, and turboprop engines, and for turboshaft engines in the icing conditions defined in Appendix C of Chapter 529 of this manual, plus 40 percent liquid water content margin, at the critical level of airspeed and air temperature; or
    2. (B) Demonstration in ground test facility: A duration of 3 cycles of alternating icing exposure corresponding to the liquid water content levels and standard cloud lengths starting in Intermittent Maximum and then in Continuous Maximum icing conditions defined in Appendix C of Chapter 525 of this manual for turbojet, turbofan, and turboprop engines, and for turboshaft engines in the icing conditions defined in Appendix C of Chapter 529 of this manual, at the critical level of air temperature.
      (effective 2021/04/08)
(c) In addition to complying with paragraph (b) of this section, the following conditions shown in Table 1 of this section unless replaced by similar CPA test conditions that are more critical or produce an equivalent level of severity, must be demonstrated by an engine test:
(effective 2021/04/08)

**Table 1—Conditions That Must be Demonstrated by an Engine Test**
(effective 2021/04/08)
Condition	Total air temperature	Supercooled water concentrations (minimum)	Median Volume diameter	Duration
1. Glaze ice conditions	21 to 25 °F (-6 to -4 °C)	2 g/m³	25 to 35 microns	(a) 10-minutes for power below sustainable level flight (idle descent). (b) Must show repetitive, stabilized operation for higher powers (50%, 75%, 100% MC).
2. Rime ice conditions	-10 to 0 ^oF (-23 to -18 °C)	1 g/m³.	15 to 25 microns	(a) 10-minutes for power below sustainable level flight (idle descent). (b) Must show repetitive, stabilized operation for higher powers (50%, 75%, 100%MC).
3. Glaze ice holding conditions. (Turbojet, turbofan, and turboprop only)	Turbojet and Turbofan, only: 10 to 18 °F (-12 to -8 °C)	Alternating cycle: First 1.7 g/m³ (1 minute), Then 0.3 g/m3 (6 minute)	20 to 30 microns	Must show repetitive, stabilize operation (or 45 minutes max).
	Turboprop, only: 2 to 10 °F (-17 to -12 ^oC)		20 to 30 microns
4. Rime ice holding conditions. (Turbojet, turbofan, and turboprop only)	Turbojet and Turbofan, only: -10 to 0 °F (-23 to -18 ^oC)	0.25g/m³	20 to 30 microns	Must show repetitive, stabilized operation (or 45 minutes max).
	Turboprop, only: 2 to 10 °F (-17 to -12 °C)	0.25g/m³	20 to 30 microns

(d) Operate at ground idle speed for a minimum of 30 minutes at each of the following icing conditions shown in Table 2 of this section with the available air bleed for icing protection at its critical condition, without adverse effect, followed by acceleration to takeoff power or thrust. During the idle operation, the engine may be run up periodically to a moderate power or thrust setting in a manner acceptable to the Minister. Analysis may be used to show ambient temperatures below the tested temperature are less critical. The applicant must document any demonstrated run ups and minimum ambient temperature capability in the engine operating manual as mandatory in icing conditions. The applicant must demonstrate, with consideration of expected airport elevations, the following:
(effective 2021/04/08)

**Table 2—Demonstration Methods for Specific Icing Conditions**
(effective 2021/04/08)
Condition	Total air temperature	Supercooled water concentrations (minimum)	Median effective particle diameter	Demonstration
1. Rime ice condition	0 to 15 °F (-18 to -9 °C)	Liquid—0.3 g/m³	15–25 microns	By engine test.
2. Glaze ice condition	20 to 30 °F (-7 to -1 °C)	Liquid—0.3 g/m³	15–25 microns	By engine test.
3. Snow ice condition	26 to 32 °F (‑3 to 0 °C)	Ice—0.9 g/m³	100 microns (minimum)	By test, analysis or combination of the two.
4. Large drop glaze ice condition (Turbojet, turbofan and turboprop only).	15 to 30 °F (-9 to -1 °C)	Liquid—0.3 g/m³	100 microns (minimum)	By test, analysis or combination of the two.

(e) Demonstrate by test, analysis, or combination of the two, acceptable operation for turbojet, turbofan, and turboprop engines in mixed phase and ice crystal icing conditions throughout Appendix D of this chapter, icing envelope throughout its flight power range, including minimum descent idling speeds.
(effective 2021/04/08)

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

(amended 2008/10/30; no previous version)

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:
(amended 2008/10/30; no previous version)

(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.
(amended 2008/10/30; no previous version)
(b) A manufacturing plan that identifies the specific manufacturing constraints necessary to consistently produce each engine life-limited part with the attributes required by the engineering plan.
(amended 2008/10/30; no previous version)
(c) A service management plan that defines in-service processes for maintenance and the limitations to repair for each engine life-limited part that will maintain attributes consistent with those required by the engineering plan. These processes and limitations will become part of the Instructions for Continued Airworthiness.
(amended 2008/10/30; no previous version)

533.71 Lubrication System

(a) General. Each lubrication system must function properly in the flight attitudes and atmospheric conditions in which an aircraft is expected to operate.
(b) Oil strainer or filter. There must be an oil strainer or filter through which all of the engine oil flows. In addition:
1. (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.
2. (2) The type and degree of filtering necessary for protection of the engine oil system against foreign particles in the oil must be specified. The applicant must demonstrate that foreign particles passing through the specified filtering means do not impair engine oil system functioning.
3. (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.
4. (4) For each strainer or filter required by this paragraph, except the strainer or filter at the oil tank outlet, there must be means to indicate contamination before it reaches the capacity established in accordance with paragraph (b)(3) of this section.
5. (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. (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.
7. (7) Each strainer or filter required by this paragraph must be accessible for draining and cleaning.
(c) Oil tanks.
1. (1) Each oil tank must have an expansion space for not less than 10 percent of the tank capacity.
2. (2) It must be impossible to inadvertently fill the oil tank expansion space.
3. (3) Each recessed oil tank filler connection that can retain any appreciable quantity of oil must have provision for fitting a drain.
4. (4) Each oil tank cap must provide an oil-tight seal. For an applicant seeking eligibility for an engine to be installed on an aeroplane approved for Extended Range Twin-Engine Operations (ETOPS), the oil tank must be designed to prevent a hazardous loss of oil due to an incorrectly installed oil tank cap.
  (effective 2013/01/21)
5. (5) Each oil tank filler must be marked with the word “oil”.
6. (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.
7. (7) There must be means to prevent entrance into the oil tank or into any oil tank outlet, of any object that might obstruct the flow of oil through the system.
8. (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. (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.
  (amended 2010/01/29)
10. (10) Leaked or spilled oil may not accumulate between the tank and the remainder of the engine.
11. (11) Each oil tank must have an oil quantity indicator or provision for one.
12. (12) If the propeller feathering system depends on engine oil:
  1. (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.
  2. (ii) The amount of trapped oil must be enough to accomplish the feathering operation and must be available only to the feathering pump; and
  3. (iii) Provision must be made to prevent sludge or other foreign matter from affecting the safe operation of the propeller feathering system.
(d) Oil Drains. A drain (or drains) must be provided to allow safe drainage of the oil system.

Each drain must:

1. (1) Be accessible; and
2. (2) Have manual or automatic means for positive locking in the closed position.
(e) Oil radiators. Each oil radiator must withstand, without failure, any vibration, inertia, and oil pressure load to which it is subjected during the block tests.

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.
(amended 2008/10/30)

(Change 533-5)

533.75 Safety Analysis

(a)
(amended 2008/10/30)
1. (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:
  (amended 2008/10/30; no previous version)
  1. (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. (ii) Consequential secondary failures and latent failures.
    (amended 2008/10/30; no previous version)
  3. (iii) Multiple failures referred to in paragraph (d) of this section or that result in the hazardous engine effects defined in paragraph (g)(2) of this section.
    (amended 2008/10/30; no previous version)
2. (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.
  (amended 2008/10/30; no previous version)
3. (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.
  (amended 2008/10/30; no previous version)
4. (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).
  (amended 2008/10/30; no previous version)
(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)
(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.
(amended 2008/10/30)
(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.
(amended 2008/10/30)
(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. (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.
  (amended 2008/10/30; no previous version)
2. (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.
  (amended 2008/12/30; no previous version)
3. (3) The provisions of specific instrumentation not otherwise required.
  (amended 2008/10/30; no previous version)
4. (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:
(amended 2008/10/30; no previous version)
1. (1) Indicating equipment;
  (amended 2008/10/30; no previous version)
2. (2) Manual and automatic controls;
  (amended 2008/10/30; no previous version)
3. (3) Compressor bleed systems;
  (amended 2008/10/30; no previous version)
4. (4) Refrigerant injection systems;
  (amended 2008/10/30; no previous version)
5. (5) Gas temperature control systems;
  (amended 2008/10/30; no previous version)
6. (6) Engine speed, power, or thrust governors and fuel control systems;
  (amended 2008/10/30; no previous version)
7. (7) Engine overspeed, overtemperature, or topping limiters;
  (amended 2008/10/30; no previous version)
8. (8) Propeller control systems; and
  (amended 2008/10/30; no previous version)
9. (9) Engine or propeller thrust reversal systems.
  (amended 2008/10/30; no previous version)
(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)
1. (1) An engine failure in which the only consequence is partial or complete loss of thrust or power (and associated engine services) from the engine will be regarded as a minor engine effect.
  (amended 2008/10/30; no previous version)
2. (2) The following effects will be regarded as hazardous engine effects:
  (amended 2008/10/30; no previous version)
  1. (i) Non-containment of high-energy debris;
    (amended 2008/10/30; no previous version)
  2. (ii) Concentration of toxic products in the engine bleed air intended for the cabin sufficient to incapacitate crew or passengers;
    (amended 2008/10/30; no previous version)
  3. (iii) Significant thrust in the opposite direction to that commanded by the pilot;
    (amended 2008/10/30; no previous version)
  4. (iv) Uncontrolled fire;
    (amended 2008/10/30; no previous version)
  5. (v) Failure of the engine mount system leading to inadvertent engine separation;
    (amended 2008/10/30; no previous version)
  6. (vi) Release of the propeller by the engine, if applicable; and
    (amended 2008/10/30; no previous version)
  7. (vii) Complete inability to shut the engine down.
    (amended 2008/10/30; no previous version)
3. (3) An effect whose severity falls between those effects covered in paragraphs (g)(1) and (g)(2) of this section will be regarded as a major engine effect.
  (amended 2008/10/30; no previous version)

533.76 Bird Ingestion

(amended 2001/03/05; no previous version)

(a) General
Compliance with (b), (c) and (d) of this section shall be in accordance with the following:
(amended 2008/10/30)
1. (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;
  (amended 2008/10/30)
2. (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. (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;
  (amended 2008/10/30)
4. (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. (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
  (amended 2008/12/30)
6. (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
(amended 2008/10/30)

Compliance with the large bird ingestion requirements shall be in accordance with the following:
(amended 2008/10/30)

(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.
(amended 2008/10/30)

(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 (square-inches)	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) (amended 2004/06/08)	2.75 (6.05)
3.90 (6,045)	3.65 (8.03)

Compliance with the small and medium bird ingestion requirements shall be in accordance with the following:
(amended 2008/12/30)

(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)
(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:
1. (i) more than a sustained 25-percent power or thrust loss,
2. (ii) the engine to be shut down during the required run-on demonstration prescribed in paragraphs (c)(7) or (c)(8) of this section,
3. (iii) the conditions defined in paragraph (b)(3) of this section, and
4. (iv) unacceptable deterioration of engine handling characteristics;
(7) except for rotorcraft engines, the following test schedule shall be used:
1. (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,
2. (ii) followed by 2 minutes without power lever movement after the ingestion,
  (amended 2004/06/08)
3. (iii) followed by 3 minutes at 75-percent of the test condition,
4. (iv) followed by 6 minutes at 60-percent of the test condition,
5. (v) followed by 6 minutes at 40-percent of the test condition,
6. (vi) followed by 1 minute at approach idle,
7. (vii) followed by 2 minutes at 75-percent of the test condition,
8. (viii) followed by stabilising at idle and engine shut down, and
9. (ix) 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)
(8) for rotorcraft engines, the following test schedule shall be used:
1. (i) ingestion so as to simulate a flock encounter within approximately 1 second elapsed time between the first ingestion and the last,
2. (ii) followed by 3 minutes at 75-percent of the test condition,
3. (iii) followed by 90 seconds at descent flight idle,
4. (iv) followed by 30 seconds at 75-percent of the test condition, and
5. (v) followed by stabilizing at idle and engine shut down, and
  (amended 2004/06/08)
6. (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)
(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

(10) if any engine operating limit(s) is exceeded during the initial 2 minutes without power lever movement, as provided by paragraph (c)(7)(ii) of this section, then it shall be established that the limit exceedence will not result in an unsafe condition.

**Table 2 to 533.76 -- Medium Flocking Bird Weight and Quantity Requirements**
Engine Inlet Throat Area (A) -- Square-meters (square-inches)	Bird quantity	Bird weight kg (lb)
0.05 (77.5) >A	none
0.05 (77.5)	1	0.35 (0.77)
0.10 (155)	1	0.45 (0.99)
0.20 (310)	2	0.45 (0.99)
0.40 (620)	2	0.70 (1.54)
0.60 (930)	3	0.70 (1.54)
1.00 (1,550)	4	0.70 (1.54)
1.35 (2,092)	1	1.15 (2.53)
	plus 3	0.70 (1.54)
1.70 (2,635)	1	1.15 (2.53)
	plus 4	0.70 (1.54)
2.10 (3,255)	1	1.15 (2.53)
	plus 5	0.70 (1.54)
2.50 (3,875)	1	1.15 (2.53)
	plus 6	0.70 (1.54)
3.90 (6045)	3	1.15 (2.53)
4.50 (6975)	4	1.15 (2.53)

**Table 3 to 533.76 -- Additional Integrity Assessment**
Engine Inlet Throat Area (A) -- Square-meters (square-inches)	Bird quantity	Bird weight kg (lb)
1.35 (2,092) >A	none
1.35 (2,092)	1	1.15(2.53)
2.90 (4,495)	2	1.15 (2.53)
3.90 (6,045)	1	1.15 (2.53)
	plus 6	0.70 (1.54)

(d) Large flocking bird
(amended 2008/10/30; no previous version)

An engine test will be performed as follows:
(amended 2008/10/30; no previous version)

(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.
(amended 2008/10/30; no previous version)
(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.
(amended 2008/10/30; no previous version)
(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:
(amended 2008/10/30; no previous version)
1. (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.
  (amended 2008/10/30; no previous version)
2. (ii) Engine shutdown during the required run-on demonstration specified in (d)(5) of this section.
  (amended 2008/10/30; no previous version)
3. (iii) The conditions specified in (b)(3) of this section.
  (amended 2008/10/30; no previous version)
(5) The following test schedule shall be used:
(amended 2008/10/30; no previous version)
1. (i) Ingestion followed by 1 minute without power lever movement.
  (amended 2008/10/30; no previous version)
2. (ii) Followed by 13 minutes at no less than 50 percent of maximum rated take-off power or thrust.
  (amended 2008/10/30; no previous version)
3. (iii) Followed by 2 minutes between 30 and 35 percent of maximum rated take-off power or thrust.
  (amended 2008/10/30; no previous version)
4. (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.
  (amended 2008/10/30; no previous version)
5. (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.
  (amended 2008/10/30; no previous version)
6. (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.
  (amended 2008/10/30; no previous version)
(6) Compliance with the large flocking bird ingestion requirements of (d) may also be demonstrated by:
(amended 2008/10/30; no previous version)
1. (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
  (amended 2008/10/30; no previous version)
2. (ii) Use of an engine subassembly test at the ingestion conditions specified in (b)(1) of this section if:
  (amended 2008/10/30; no previous version)
  1. (A) All components critical to complying with the requirements of (d) of this section are included in the subassembly test;
    (amended 2008/10/30; no previous version)
  2. (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
    (amended 2008/10/30; no previous version)
  3. (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.
    (amended 2008/10/30; no previous version)

(7) Applicants shall demonstrate that an unsafe condition will not result if any engine operating limit is exceeded during the run-on period.
(amended 2008/10/30; no previous version)

**Table 4 to 533.76 -- Large Flocking Bird Mass and Weight**
Engine Inlet Throat Area (A)--square-meters (square-inches)	Bird quantity	Bird mass and weight kg (lbs)
A < 2.50 (3875)	none
2.50 (3875) £ A < 3.50 (5425)	1	1.85 (4.08)
3.50 (5425) £ A < 3.90 (6045)	1	2.10 (4.63)
3.90 (6045) £ A	1	2.50 (5.51)
(amended 2008/10/30; no previous version)

533.77 Foreign Object Ingestion - Ice
(amended 2001/03/05)

(a) Compliance with the requirements of this section must be demonstrated by engine ice ingestion test or by validated analysis showing equivalence of other means for demonstrating soft body damage tolerance.
(effective 2021/04/08)
(b) Reserved.
(amended 2001/03/05)
(c) Ingestion of ice under the conditions of this section may not
(effective 2021/04/08)
1. (1) cause an immediate or ultimate unacceptable sustained power or thrust loss; or
  (effective 2021/04/08)
2. (2) require the engine to be shut down.
  (amended 2001/03/05)
(d) For an engine that incorporates a protection device, compliance with this section need not be demonstrate with respect to ice formed forward of the protection device if it is shown that:
1. (1) Such ice is of a size that will not pass through the protective device;
2. (2) The protective device will withstand the impact of the ice; and
3. (3) the ice stopped by the protective device will 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.
  (effective 2021/04/08)
(e) Compliance with the requirements of this section must be demonstrated by engine ice ingestion test under the following ingestion conditions or by validated analysis showing equivalence of other means for demonstrating soft body damage tolerance.
1. (1) The minimum ice quantity and dimensions will be established by the engine size as defined in Table 1 of this section.
2. (2) The ingested ice dimensions are determined by linear interpolation between table values, and are based on the actual engine’s inlet hilite area.
3. (3) The ingestion velocity will simulate ice from the inlet being sucked into the engine.
4. (4) Engine operation will be at the maximum cruise power or thrust unless lower power is more critical.
  (effective 2021/04/08)

**Table 1—Minimum Ice Slab Dimensions Based on Engine Inlet Size**
(effective 2021/04/08)
Engine Inlet Hilite area (sq. inch)	Thickness (inch)	Width (inch)	Length (inch)
0	0.25	0	3.6
80	0.25	6	3.6
300	0.25	12	3.6
700	0.25	12	4.8
2800	0.35	12	8.5
5000	0.43	12	11.0
7000	0.50	12	12.7
7900	0.50	12	13.4
9500	0.50	12	14.6
11300	0.50	12	15.9
13300	0.50	12	17.1
16500	0.5	12	18.9
20000	0.5	12	20.0

533.78 Rain and Hail Ingestion

(a) All engines.
1. (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:
  1. (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. (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. (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. (2) Continuation of the rain ingestion during subsequent rapid deceleration to minimum idle, then
3. (3) Continuation of the rain ingestion during three minutes at minimum idle power to be certified for flight operation, then
4. (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. (1) The subject rain and hail constituents are of a size that will not pass through the protection device;
2. (2) The protection device will withstand the impact of the subject rain and hail constituents; and
3. (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.]
  (Change 533-5)

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;
(b) Permit on-off cycling;
(c) Be controllable within the intended range of operation;
(d) Upon a failure or malfunction of augmentor combustion, not cause the engine to lose thrust other than that provided by the augmentor; and
(e) Have controls that function compatibly with the other engine controls and automatically shut off augmentor fuel flow if the engine rotor speed drops below the minimum rotational speed at which the augmentor is intended to function.

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. (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. (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.
(Change 533-5)

533.84 Engine Overtorque Test

(amended 2010/05/27; no previous version)

(a) If approval of a maximum engine overtorque is sought for an engine incorporating a free power turbine, compliance with this section must be demonstrated by testing.
1. (1) The test may be run as part of the endurance test requirement of section 533.87. Alternatively, tests may be performed on a complete engine or equivalent testing on individual groups of components.
2. (2) Upon conclusion of tests conducted to show compliance with this section, each engine part or individual groups of components must meet the requirements of paragraphs 533.93(a)(1) and (a)(2).
(b) The test conditions must be as follows:
1. (1) A total of 15 minutes run at the maximum engine overtorque to be approved. This may be done in separate runs, each being of at least 2.5 minutes duration.
2. (2) A power turbine rotational speed equal to the highest speed at which the maximum overtorque can occur in service. The test speed may not be more than the limit speed of take-off or OEI ratings longer than 2 minutes.
3. (3) For engines incorporating a reduction gearbox, a gearbox oil temperature equal to the maximum temperature when the maximum engine overtorque could occur in service; and for all other engines, an oil temperature within the normal operating range.
4. (4) A turbine entry gas temperature equal to the maximum steady state temperature approved for use during periods longer than 20 seconds when operating at conditions not associated with 30-second or 2 minute OEI ratings. The requirement to run the test at the maximum approved steady state temperature may be waived by the Minister if the applicant can demonstrate that other testing provides substantiation of the temperature effects when considered in combination with the other parameters identified in paragraphs (b)(1), (b)(2) and (b)(3) of this section.

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.
(b) A power check at sea level conditions must be accomplished on the endurance test engine after the endurance test and any change in power characteristics which occurs during the endurance test must be determined. Measurements taken during the final portion of the endurance test may be used in showing compliance with the requirements of this paragraph.
(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.

(Change 533-5)

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. (1) The runs must be made in the order found appropriate by the Minister for the particular engine being tested.
2. (2) Any automatic engine control that is part of the engine must control the engine during the endurance test except for operations where automatic control is normally overridden by manual control or where manual control is otherwise specified for a particular test run.
3. (3) Except as provided in paragraph (a)(5) of this section, power or thrust, gas temperature, rotor shaft rotational speed, and, if limited, temperature of external surfaces of the engine must be at least 100 percent of the value associated with the particular engine operation being tested. More than one test may be run if all parameters cannot be held at the 100 percent level simultaneously.
4. (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. (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.
  (amended 2010/01/29)
6. (6) Each accessory drive and mounting attachment must be loaded in accordance with paragraphs (a)(6)(i) and (ii) of this section, except as permitted by paragraph (a)(6)(iii) of this section for the test required under paragraph (f) of this section.
  (amended 2010/01/29)
  1. (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)
  2. (ii) The endurance test of any accessory drive and mounting attachment under load may be accomplished on a separate rig if the validity of the test is confirmed by an approved analysis.
    (amended 2010/01/29)
  3. (iii) The applicant is not required to load the accessory drives and mounting attachments when running the tests under paragraphs (f)(1) through (f)(8) of this section if the applicant can substantiate that there is no significant effect on the durability of any accessory drive or engine component. However, the applicant must add the equivalent engine output power extraction from the power turbine rotor assembly to the engine shaft output.
    (amended 2010/01/29)
7. (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. (8) If the number of occurrences of either transient rotor shaft overspeed, transient gas overtemperature or transient engine overtorque is limited, that number of the accelerations required by paragraphs (b) through (g) of this section must be made at the limiting overspeed, overtemperature or overtorque. If the number of occurrences is not limited, half the required accelerations must be made at the limiting overspeed, overtemperature or overtorque.
  (amended 2010/05/27)
9. (9) For each engine type certificated for use on supersonic aircraft the following additional test requirements apply:
10. (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.
11. (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.
12. (iii) During the simulated supersonic runs the fuel temperature and induction air temperature may not be less than the limiting temperature.
13. (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).
14. (v) During the runs at thrust settings for maximum continuous thrust and percentages thereof, the engine must be operated with the inlet air distortion at the limit for those thrust settings.
(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. (1) Take-off and idling. One hour of alternate 5-minute periods at rated take-off power or thrust and at idling power or thrust. The developed power or thrust 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 applicant. The applicant may, during any one period, manually control the rotor speed and power or 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 subparagraph (5) of this paragraph.
  (effective 2013/01/28)
2. (2) Rated maximum continuous and take-off power or thrust. Thirty minutes at:
  (effective 2013/01/28)
  1. (i) Rated maximum continuous power or thrust during fifteen of the twenty-five 6-hour endurance test cycles; and
    (effective 2013/01/28)
  2. (ii) Rated take-off power or thrust during ten of the twenty-five 6-hour endurance test cycles.
    (effective 2013/01/28)
3. (3) Rated maximum continuous power or thrust. One hour and 30 minutes at rated maximum continuous power or thrust.
  (effective 2013/01/28)
4. (4) Incremental cruise power or thrust. Two hours and 30 minutes at the successive power lever positions corresponding to at least 15 approximately equal speed and time increments between maximum continuous engine rotational speed and ground or minimum idle rotational speed. For engines operating at constant speed, the power or thrust may be varied in place of speed. If there is significant peak vibration anywhere between ground idle and maximum continuous conditions, the number of increments chosen may be changed to increase the amount of running made while subject to the peak vibrations up to not more than 50 percent of the total time spent in incremental running.
  (effective 2013/01/28)
5. (5) Acceleration and deceleration runs. Thirty minutes of accelerations and decelerations, consisting of 6 cycles from idling power or thrust to rated take-off power or 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 one 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 two seconds.
  (effective 2013/01/28)
6. (6) Starts. One hundred (100) starts must be made, of which twenty-five (25) starts must be preceded by at least a two (2)-hour engine shutdown. There must be at least ten (10) false engine starts, pausing for the applicant's specified minimum fuel drainage time, before attempting a normal start. There must be at least ten (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.
  (effective 2013/01/28)
(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. (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 applicant. 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) (6) of this section.
  (effective 2013/01/28)
2. (2) Rated maximum continuous and takeoff power. Thirty minutes at —
  (amended 2010/01/29)
  1. (i) Rated maximum continuous power during fifteen of the twenty-five 6-hour endurance test cycles; and
    (amended 2010/01/29)
  2. (ii) Rated take-off power during ten of the twenty-five 6-hour endurance test cycles.
    (amended 2010/01/29)
3. (3) Rated maximum continuous power. One hour at rated maximum continuous power.
  (amended 2010/01/29)
4. (4) Rated 30-minute OEI power. Thirty minutes at rated 30-minute OEI power.
  (amended 2010/01/29)
5. (5) Incremental cruise power. Two hours and 30 minutes at the successive power lever positions corresponding with not less than 15 approximately equal speed and time increments between maximum continuous engine rotational speed and ground or minimum idle rotational speed. For engines operating at constant speed, power may be varied in place of speed. If there are significant peak vibrations anywhere between ground idle and maximum continuous conditions, the number of increments chosen must be changed to increase the amount of running conducted while being subjected to the peak vibrations up to not more than 50 percent of the total time spent in incremental running.
  (effective 2010/01/29)
6. (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.
  (amended 2010/01/29)
7. (7) Starts. On hundred starts, of which twenty-five (25) starts must be preceded by at least a (2)-hour engine shut-down. There must be at least ten (10) false engine starts, pausing for the applicant's specified minimum fuel drainage time, before attempting a normal start. There must be at least ten (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.
  (amended 2010/01/29)
(d) Rotorcraft engines for which a continuous OEI rating is desired. For each rotorcraft engine for which a continuous OEI power rating is desired, the applicant must conduct the following series of tests:
1. (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 applicant. 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 (d)(6) of this section.
  (effective 2013/01/28)
2. (2) Rated maximum continuous and take-off power. Thirty minutes at:
  1. (i) Rated maximum continuous power during fifteen of the twenty-five 6-hour endurance test cycles; and
  2. (ii) Rated take-off power during ten of the twenty-five 6-hour endurance test cycles.
3. (3) Rated continuous OEI power. One hour at rated continuous OEI power.
4. (4) Rated maximum continuous power. One hour at rated maximum continuous power.
5. (5) Incremental cruise power. Two hours at the successive power lever positions corresponding with not less than 12 approximately equal speed and time increments between maximum continuous engine rotational speed and ground or minimum idle rotational speed. For engines operating at constant speed, power may be varied in place of speed. If there are significant peak vibrations anywhere between ground idle and maximum continuous conditions, the number of increments chosen must be changed to increase the amount of running conducted while being subjected to the peak vibrations up to not more than 50 percent of the total time spent in incremental running.
6. (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.
7. (7) Starts. One hundred starts, 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.
(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. (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 applicant. 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 paragraphs (b)(5), (c)(6) or (d)(6) of this section, as applicable.
  (effective 2013/01/28)
2. (2) The tests required in paragraphs (b)(2) through (b)(6), or (c)(2) through (c)(7), 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)(4) 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.
  (effective 2013/01/28)
(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)
1. (1) Take-off power. Three minutes at rated take-off power.
2. (2) 30-second OEI power. Thirty seconds at rated 30-second OEI power.
3. (3) 2-minute OEI power. Two minutes at rated 2-minute OEI power.
4. (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)
5. (5) 50 percent take-off power. One minute at 50 percent take-off power.
6. (6) 30-second OEI power. Thirty seconds at rated 30-second OEI power.
7. (7) 2-minute OEI power. Two minutes at rated 2-minute OEI power.
8. (8) Idle. One minute at flight idle.
  (amended 2010/01/29)
(g) Supersonic aircraft engines. For each engine type certificated for use on supersonic aircraft the applicant must conduct the following:
1. (1) Subsonic test under sea level ambient atmospheric conditions. Thirty runs of one hour each must be made, consisting of:
  1. (i) Two periods of 5 minutes at rated take-off augmented thrust each followed by 5 minutes at idle thrust;
  2. (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;
  3. (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
  4. (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. (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:
3. (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;
4. (ii) 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 80 percent of rated maximum continuous augmented thrust;
5. (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;
6. (iv) One period repeating the runs specified in paragraph (g)(2)(i) and (ii) of this section; and
7. (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.
8. (3) Starts. One hundred starts must be made, of which 25 starts must be preceded by an engine shutdown of at least 2 hours. There must be at least 10 false engine starts, pausing for the applicant's specified minimum fuel drainage time before attempting a normal start. At least 10 starts must be normal restarts, each made no later than 15 minutes after engine shutdown. The starts may be made at any time, including the period of endurance testing.
  (Change 533-1 (87-01-01))
  
  (Change 533-2 (89-01-01))
  
  (Change 533-5)

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 533.28(k), 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)
(c) A separate test vehicle may be used for each test condition.
(amended 2010/01/29)
(Change 533-5)

533.89 Operation Test

(a) The operation test must include testing found necessary by the Minister to demonstrate:
1. (1) Starting, idling, acceleration, overspeeding, ignition, functioning of the propeller (if the engine is designated to operate with a propeller);
2. (2) Compliance with the engine response requirements of 533.73; and
3. (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:
  1. (i) No bleed air and power extraction for aircraft use.
  2. (ii) Maximum allowable bleed air and power extraction for aircraft use.
  3. (iii) An intermediate value for bleed air and power extraction representative of that which might be used as a maximum for aircraft during approach to a landing.
4. (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 applicant, except an applicant for an engine being type certificated through amendment of an existing type certificate or through supplemental type certification procedures, must complete one of the following tests on an engine that substantially conforms to the type design to establish when the initial maintenance inspection is required:
(effective 2013/01/21)

(a) an approved engine test that simulates the conditions in which the engine is expected to operate in service, including typical start-stop cycles.
(effective 2013/01/21)
(b) an approved engine test conducted in accordance with 533.201(c) through (f).
(effective 2013/01/21)

533.91 Engine System and Component Tests

(amended 2010/01/29)

(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.
(amended 2010/01/29)
(b) Temperature limits must be established for those components that require temperature controlling provisions in the aircraft installation to assure satisfactory functioning, reliability, and durability.
(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.
(amended 2010/01/29)
(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.

(Change 533-5)

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. (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. (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. (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. (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
  (amended 2010/01/29)
(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).
(Change 533-5)

533.94 Blade containment and rotor unbalance tests

(a) Except as provided in paragraph (b) of this section, it must be demonstrated by engine tests that the engine is capable of containing damage without catching fire and without failure of its mounting attachments when operated for at least 15 seconds, unless the resulting engine damage induces a self shutdown, after each of the following events:
1. (1) Failure of the most critical compressor or fan 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.
2. (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:
1. (1) That test, of the two prescribed, produces the least rotor unbalance; and
2. (2) The analysis is shown to be equivalent to the test.

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:

(a) Feathering operation: 25 cycles.
(b) Negative torque and thrust system operation: 25 cycles from rated maximum continuous power.
(c) Automatic decoupler operation: 25 cycles from rated maximum continuous power (if repeated decoupling and recoupling in service is the intended function of the device).
(d) Reverse thrust operation: 175 cycles from the flight-idle position to full reverse and 25 cycles at rated maximum continuous power from full forward to full reverse thrust. At the end of each cycle the propeller must be operated in reverse pitch for a period of 30 seconds at the maximum rotational speed and power specified by the applicant for reverse pitch operation.

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.
(c) One hundred engine starts and stops with the propeller brake engaged.
(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).
(Change 533-1 (87-01-01))

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

(a) Each applicant may, in making a block test, use separate engines of identical design and construction in the vibration, calibration, endurance, and operation tests, except that, if a separate engine is used for the endurance test it must be subjected to a calibration check before starting the endurance test.
(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.
(c) Each applicant must furnish all testing facilities, including equipment and competent personnel, to conduct the block tests.

Subchapter G - Special Requirements: Turbine Aircraft Engines

(effective 2013/01/21)

533.201 Design and Test Requirements for Early ETOPS Eligibility

An applicant seeking type design approval for an engine to be installed on a two-engine aeroplane approved for extended operations (ETOPS), without the service experience specified in Chapter 525, Appendix K, K525.2.1 of this manual, must comply with the following:
(effective 2019/08/25)

(a) The engine must be designed using a design quality process acceptable to the Minister that ensures the design features of the engine minimize the occurrence of failures, malfunctions, defects, and maintenance errors that could result in an in-flight shutdown (IFSD), loss of thrust control, or other power loss.
(b) The design features of the engine must address problems shown to result in an IFSD, loss of thrust control, or other power loss in the applicant's other relevant type designs approved within the past 10 years, to the extent that adequate service data is available within that 10-year period. An applicant without adequate service data must show experience with and knowledge of problem mitigating design practices equivalent to that gained from actual service experience in a manner acceptable to the Minister.
(c) Except as specified in paragraph (f) of this section, the applicant must conduct a simulated ETOPS mission cyclic endurance test in accordance with an approved test plan on an engine that substantially conforms to the type design. The test must:
1. (1) include a minimum of 3,000 representative service start-stop mission cycles and three simulated diversion cycles at maximum continuous thrust or power for the maximum diversion time for which ETOPS eligibility is sought. Each start-stop mission cycle must include the use of take-off, climb, cruise, descent, approach, and landing thrust or power and the use of thrust reverse (if applicable). The diversions must be evenly distributed over the duration of the test. The last diversion must be conducted within 100 cycles of the completion of the test;
2. (2) be performed with the high speed and low speed main engine rotors independently unbalanced to obtain a minimum of 90 percent of the recommended field service maintenance vibration levels. For engines with three main engine rotors, the intermediate speed rotor must be independently unbalanced to obtain a minimum of 90 percent of the recommended production acceptance vibration level. The required peak vibration levels must be verified during a slow acceleration and deceleration run of the test engine covering the main engine rotor operating speed ranges;
3. (3) include a minimum of 3,000,000 vibration cycles for each 60 rpm incremental step of the typical high-speed rotor start-stop mission cycle. The test may be conducted using any rotor speed step increment from 60 to 200 rpm provided the test encompasses the typical service start-stop cycle speed range. For incremental steps greater than 60 rpm, the minimum number of vibration cycles must be linearly increased up to 10,000,000 cycles for a 200 rpm incremental step;
4. (4) include a minimum of 300,000 vibration cycles for each 60 rpm incremental step of the high-speed rotor approved operational speed range between minimum flight idle and cruise power not covered by paragraph (c)(3) of this section. The test may be conducted using any rotor speed step increment from 60 to 200 rpm provided the test encompasses the applicable speed range. For incremental steps greater than 60 rpm the minimum number of vibration cycles must be linearly increased up to 1,000,000 for a 200 rpm incremental step;
5. (5) include vibration surveys at periodic intervals throughout the test. The equivalent value of the peak vibration level observed during the surveys must meet the minimum vibration requirement of 533.201(c)(2).
(d) Prior to the test required by paragraph (c) of this section, the engine must be subjected to a calibration test to document power and thrust characteristics.
(e) At the conclusion of the testing required by paragraph (c) of this section, the engine must:
1. (1) be subjected to a calibration test at sea-level conditions. Any change in power or thrust characteristics must be within approved limits.
2. (2) be inspected in accordance with the recommendations and limits contained in the Instructions for Continued Airworthiness submitted in compliance with 533.4.
3. (3) be completely disassembled and inspected
  1. (i) in accordance with the applicable inspection recommendations and limits contained in the Instructions for Continued Airworthiness submitted in compliance with 533.4;
  2. (ii) with consideration of the causes of IFSD, loss of thrust control, or other power loss identified by paragraph (b) of this section; and
  3. (iii) in a manner to identify wear or distress conditions that could result in an IFSD, loss of thrust control, or other power loss not specifically identified by paragraph (b) of this section or addressed within the Instructions for Continued Airworthiness; and
4. (4) not show wear or distress to the extent that could result in an IFSD, loss of thrust control, or other power loss within a period of operation before the component, assembly, or system would likely have been inspected or functionally tested for integrity while in service. Such wear or distress must have corrective action implemented through a design change, a change to maintenance instructions, or operational procedures before ETOPS eligibility is granted. The type and frequency of wear and distress that occurs during the engine test must be consistent with the type and frequency of wear and distress that would be expected to occur on ETOPS eligible engines.
(f) An alternative mission cycle endurance test that provides an equivalent demonstration of the unbalance and vibration specified in paragraph (c) of this section may be used when approved by the Minister.
(g) For an applicant using the simulated ETOPS mission cyclic endurance test to comply with 533.90(a), the test may be interrupted so that the engine may be inspected in accordance with the recommendations and limits contained in the Instructions for Continued Airworthiness submitted in compliance with 533.4 or other method, using criteria acceptable to the Minister, after completion of the test cycles required to comply with 533.90(a). Following the inspection, the ETOPS test must be resumed to complete the requirements of this section.

Appendix A - Instructions for Continued Airworthiness
Appendix B - Certification Standard Atmospheric Concentrations of Rain and Hail
Appendix C - (Reserved)
Appendix D - Mixed Phase and Ice Crystal Icing Envelope (Deep Convective Clouds)

Appendix A - Instructions for Continued Airworthiness

Content last revised: 1999/12/01

A533.1 General

(a) This appendix specifies requirements for the preparation of Instructions for Continued Airworthiness as required by 533.4.

(b) The Instruction for Continued Airworthiness for each engine must include the Instructions for Continued Airworthiness for all engine parts. If Instructions for Continued Airworthiness are not supplied by the engine part manufacturer for an engine part, the Instructions for Continued Airworthiness for the engine must include the information essential to the continued airworthiness of the engine.

(c) The applicant must submit to the Minister a program to show how changes to the Instructions for Continued Airworthiness made by the applicant or by the manufacturers of engine parts will be distributed.

A533.2 Format

(a) The Instructions for Continued Airworthiness must be in the form of a manual or manuals as appropriate for the quantity of data to be provided.

(b) The format of the manual or manuals must provide for a practical arrangement.

A533.3 Content

The Instructions for Continued Airworthiness must contain the following manuals or sections, as appropriate, and information:

(a) Engine Maintenance Section.

(1) Introduction information that includes an explanation of the engine's features and data to the extent necessary for maintenance or preventive maintenance.

(2) A detailed description of the engine and its components, systems and installations.

(3) Installation instructions, including proper procedures for uncrating, deinhibiting, acceptance checking, lifting and attaching accessories, with any necessary checks.

(4) Basic control and operating information describing how the engine components, systems, and installations operate, and information describing the methods of starting, running, testing and stopping the engine and its parts including any special procedures and limitations that apply.

(5) Servicing information that covers details regarding servicing points, capacities of tanks, reservoirs, types of fluids to be used, pressures applicable to the various systems, locations of lubrication points, lubricants to be used and equipment required for servicing.

(6) Scheduling information for each part of the engine that provides the recommended periods at which it should be cleaned, inspected, adjusted, tested, and lubricated, and the degree of inspection, the applicable wear tolerances, and work recommended at these periods. However, the applicant may refer to an accessory, instrument, or equipment manufacturer as the source of this information if the applicant shows that the item has an exceptionally high degree of complexity requiring specialised maintenance techniques, test equipment, or expertise. The recommended overhaul periods and necessary cross-references to the Airworthiness Limitations section of the manual must also be included. In addition, the applicant must include an inspection program that includes the frequency and extent of the inspections necessary to provide for the continued airworthiness of the engine.

(7) Troubleshooting information describing probable malfunctions, how to recognise those malfunctions, and the remedial action for those malfunctions.

(8) Information describing the order and method of removing the engine and its parts and replacing parts, with any necessary precautions to be taken. Instructions for proper ground handling, crating and shipping must also be included.

(9) A list of the tools and equipment necessary for maintenance and directions as to their method of use.

(b) Engine Overhaul Section.

(1) Disassembly information including the order and method of disassembly for overhaul.

(2) Cleaning and inspection instructions that cover the materials and apparatus to be used and methods and precautions to be taken during overhaul. Methods of overhaul inspection must also be included.

(3) Details of all fits and clearances relevant to overhaul.

(4) Details of repair methods for worn or otherwise substandard parts and components along with the information necessary to determine when replacement is necessary.

(5) The order and method of assembly at overhaul.

(6) Instructions for testing after overhaul.

(7) Instructions for storage preparation including any storage limits.

(8) A list of tools needed for over-haul.

For an applicant seeking eligibility for an engine to be installed on an aeroplane approved for Extended Range Twin-Engine Operations (ETOPS), the Instructions for Continued Airworthiness must include procedures for engine condition monitoring. The engine condition monitoring procedures must be able to determine, prior to flight, whether an engine is capable of providing, within approved engine operating limits, maximum continuous power or thrust, bleed air, and power extraction required for a relevant engine inoperative diversion.

Information Note: For an engine to be installed on a twin-engine aeroplane approved for ETOPS, the engine condition monitoring procedures must be validated before ETOPS eligibility is issued.

A533.4 Airworthiness Limitations Section

The Instructions for Continued Airworthiness must contain a section titled Airworthiness Limitations that is segregated and clearly distinguishable from the rest of the document.
(amended 2010/01/29)

(a) For all engines:
(amended 2010/01/29)

(1) The Airworthiness Limitations section must set forth each mandatory replacement time, inspection interval, and related procedure required for type certification. If the Instructions for Continued Airworthiness consist of multiple documents, the section required by this paragraph must be included in the principal manual.
(amended 2010/01/29)

(2) This section must contain a legible statement in a prominent location that reads: “The Airworthiness Limitations Section is approved by the Minister and specifies maintenance required by any applicable airworthiness or operational rule, unless an alternative program has been approved by the Minister.”
(amended 2010/01/29)

(b) For rotorcraft engines having 30-second OEI and 2-minute OEI ratings:
(amended 2010/01/29)

(1) The Airworthiness Limitations section must also prescribe the mandatory post-flight inspections and maintenance actions associated with any use of either 30-second OEI or 2-minute OEI ratings.
(amended 2010/01/29)

(2) The applicant must validate the adequacy of the inspections and maintenance actions required under paragraph (b)(1) of A533.4.
(amended 2010/01/29)

(3) The applicant must establish an in-service engine evaluation program to ensure the continued adequacy of the instructions for mandatory post-flight inspections and maintenance actions prescribed under paragraph (b)(1) of A533.4 and of the data for 533.5(b)(4) pertaining to power availability. The program must include service engine tests or equivalent service engine test experience on engines of similar design and evaluations of service usage of the 30-second OEI or 2-minute OEI ratings.
(amended 2010/01/29)

FAR: The Instructions for Continued Airworthiness must contain a section titled Airworthiness Limitations that is segregated and clearly distinguishable from the rest of the document.
(amended 2010/01/29)

(a) For all engines:
(amended 2010/01/29)

(1) The Airworthiness Limitations section must set forth each mandatory replacement time, inspection interval, and related procedure required of type certification. If the Instructions for Continued Airworthiness consist of multiple documents, the section required by this paragraph must be included in the principal manual.
(amended 2010/01/29)

(2) This section must contain a legible statement in a prominent location that reads: “The Airworthiness Limitations Section is FAA approved and specifies maintenance required under 43.16 and 91.403 of the Federal Aviation Regulations unless an alternative program has been FAA approved.”
(amended 2010/01/29)

(Change 533-1 (87-01-01))

(Change 533-3 (91-11-01))

Appendix B - Certification Standard Atmospheric Concentrations of Rain and Hail

Content last revised: 1999/12/01

Figure B1, Table B1, Table B2, Table B3, and Table B4 specify the atmospheric concentrations and size distributions of rain and hail for establishing certification, in accordance with the requirements of Sec.533.78(a)(2). In conducting tests, normally by spraying liquid water to simulate rain conditions and by delivering hail fabricated from ice to simulate hail conditions, the use of water droplets and hail having shapes, sizes and distributions of sizes other than those defined in this appendix B, or the use of a single size or shape for each water droplet or hail, can be accepted, provided that applicant shows that the substitution does not reduce the severity of the test.

Figure B1 - Illustration of Rain and Hail Threats, Certification concentrations are obtained using Tables B1 and B2

**Table B1 -** Certification Standard Atmospheric Rain Concentrations
Altitude (feet)	Rain Water Content (RWC) (grams water/ meter³ air
0...............................................	20.0
20,000.......................................	20.0
26,300.......................................	15.2
32,700.......................................	10.8
39,300.......................................	7.7
46,000.......................................	5.2

RWC values at other altitudes may be determined by linear interpolation.

Note: Source of data--Results of the Aerospace Industries Association (AIA) Propulsion Committee Study, Project PC 338-1, June 1990.

**Table B2 -** Certification Standard Atmospheric Hail Concentrations¨
Altitude (feet)	Hail Water Content (HWC) (grams water/ meter³ air
0........................................................	6.0
7,300..................................................	8.9
8,500..................................................	9.4
10,000.................................................	9.9
12,000.................................................	10.0
15,000.................................................	10.0
16,000.................................................	8.9
17,700.................................................	7.8
19,300.................................................	6.6
21,500.................................................	5.6
24,300.................................................	4.4
29,000.................................................	3.3
46,000..................................................	0.2

HWC values at other altitudes may be determined by linear interpolation. The hail threat below 7,300 feet and above 29,000 feet is based on linearly extrapolated data.

Note: Source of data--Results of the Aerospace Industries Association (AIA Propulsion Committee (PC) Study, Project PC 338-1, June 1990.)

**Table B3 -** Certification Standard Atmospheric Rain Droplet Size Distribution
Rain Droplet Diameter (mm)	Contribution Total RWC (%)
0-0.49 ...............................................	0
0.50-0.99 ..........................................	2.25
1.00-1.49 ..........................................	8.75
1.50-1.99 ..........................................	16.25
2.00-2.49 ..........................................	19.00
2.50-2.99 ..........................................	17.75
3.00-3.49 ..........................................	13.50
3.50-3.99 ..........................................	9.50
4.00-4.49 ..........................................	6.00
4.50-4.99 ..........................................	3.00
5.00-5.49 ..........................................	2.00
5.50-5.99 ..........................................	1.25
6.00-6.49 ..........................................	0.50
6.50-7.00 ..........................................	0.25
Total ..................................................	100.00

Median diameter of rain droplets in 2.66 mm

Note: Source of data--Results of the Aerospace Industries Association (AIA Propulsion Committee (PC) Study, Project PC 338-1, June 1990.)

**Table B4 -** Certification Standard Atmospheric Hail Size Distribution
Hail Diameter (mm)	Contribution Total HWC (%)
0-4.9 ..................................................	0
5.0-9.9 ...............................................	17.00
10.0-14.9 ...........................................	25.00
15.0-19.9 ...........................................	22.50
20.0-24.9 ...........................................	16.00
25.0-29.9 ...........................................	9.75
30.0-34.9 ...........................................	4.75
35.0-39.9 ...........................................	2.50
40.0-44.9 ...........................................	1.50
45.0-49.9 ..........................................	0.75
50.0-55.0 ...........................................	0.25
Total ..................................................	100.00

Median diameter of hail is 16 mm

Note: Source of data--Results of the Aerospace Industries Association (AIA Propulsion Committee (PC) Study, Project PC 338-1, June 1990.)

(Change 533-5)

Appendix C - (Reserved)

(effective 2021/04/08)

Reserved.
(effective 2021/04/08)

Appendix D - Mixed Phase and Ice Crystal Icing Envelope (Deep Convective Clouds)

(effective 04/08/2021)

The ice crystal icing envelope is depicted in Figure D1 of this Appendix.
(effective 2021/04/08)

Figure D1 – Convective Cloud Ice Crystal Envelope

Figure D1 – Convective Cloud Ice Crystal Envelope

The diagram determines the altitude and ambient temperature range within the ice crystal icing (deep convective clouds) envelope. The horizontal axis represents altitude from 0 to 50,000 feet. The vertical axis represents ambient temperature from 0 degrees Celsius to -70 degrees Celsius. The ice crystal icing envelope starts at around -3.5 degrees Celsius on a horizontal line from approximately 4,048 feet to 20, 000 feet altitude. At 20, 000 feet, a line is drawn on a negative slope till it reaches -60 degrees Celsius and around 47,381 feet. At this point, the line becomes horizontal and extends to the left till 39, 524 feet altitude. At 39,524 feet, a line is drawn on a slope till the -40 degrees Celsius and 30,000 feet point. The line is then extended to the left on the horizontal -40 degrees line to approximately 18,810 feet altitude. From this point, a line is drawn till the -10 degrees Celsius and 4,048 feet altitude point. Finally, a vertical line is drawn to -3.5 degrees Celsius and 4,048 feet altitude. The white area enclosed by the joined lines is the envelope.

Within the envelope, total water content (TWC) in g/m3 has been determined based upon the adiabatic lapse defined by the convective rise of 90% relative humidity air from sea level to higher altitudes and scaled by a factor of 0.65 to a standard cloud length of 17.4 nautical miles. Figure D2 of this Appendix displays TWC for this distance over a range of ambient temperature within the boundaries of the ice crystal envelope specified in Figure D1 of this Appendix.
(effective 2021/04/08)

Figure D2 – Total Water Content
(effective 2021/04/08)

Figure D2 – Total Water Content

Within the ice crystal icing envelope in Figure D1, the total water content (TWC) in g/m³ has been determined based upon the adiabatic lapse defined by the convective rise of 90% relative humidity air from sea level to higher altitudes and scaled by a factor of 0.65 to a standard cloud length of 17.4 nautical miles. Figure D2 displays TWC for this distance over a range of ambient temperature within the boundaries of the ice crystal envelope specified in Figure D1. The horizontal axis represents altitude from 0 to 50,000 feet. The vertical axis represents TWC from 0 to 6 g/m³. Curves are drawn from approximately 4,048 feet to 47,381 feet altitudes. Several curves are plotted as a function of ambient temperatures ranging from 0 degrees Celsius to -60 degrees Celsius.

Ice crystal size median mass dimension (MMD) range is 50–200 microns (equivalent spherical size) based upon measurements near convective storm cores.

The TWC can be treated as completely glaciated (ice crystal) except as noted in the Table 1 of this Appendix.
(effective 2021/04/08)

**Table 1—Supercooled Liquid Portion of TWC**
(effective 2021/04/08)
Temperature range—deg C	Horizontal cloud length—nautical miles	LWC— g/m³
0 to -20	≤50	≤1.0
0 to -20	Indefinite	≤0.5
< -20		0

The TWC levels displayed in Figure D2 of this Appendix represent TWC values for a standard exposure distance (horizontal cloud length) of 17.4 nautical miles that must be adjusted with length of icing exposure.
(effective 2021/04/08)

Figure D3 – Exposure Length Influence on TWC
(effective 2021/04/08)

Figure D3 – Exposure Length Influence on TWC

The TWC levels displayed in Figure D2 represent TWC values for a standard exposure distance (horizontal cloud length) of 17.4 nautical miles that must be adjusted with length of icing exposure. Figure D3 displays the TWC factor for the horizontal extent. The diagram is displayed on a logarithmic scale starting on the horizontal axis representing cloud horizontal extent from 1 to 1000 nautical miles. The vertical axis represents total water content (TWC) scale factor from 0.6 to 1.2. At approximately 4.5 nautical miles and slightly above 1.12, a line is drawn along a negative slope till the point defined by the 200 nautical mile and 0.74 TWC scale factor.

(effective)

Airworthiness Manual Chapter 533 - Aircraft Engines - Canadian Aviation Regulations (CARs)

Table of contents

Interpretation Provision for Part V Standards

Preamble

General

1. General

2. FAR Amendments

3. CARAC Proposed Amendment Recommendations

Subchapter A - General

533.1 Applicability

533.3 General

533.4 Instructions for Continued Airworthiness

533.5 Instruction Manual for Installing and Operating the Engine

533.7 Engine Ratings & Operating Limitations

533.8 Selection of Engine Power and Thrust Ratings

Subchapter B - Design & Construction: General

533.11 Applicability

533.13 (Reserved)

533.14 (Removed)

533.15 Materials

533.17 Fire Protection

533.19 Durability

533.21 Engine Cooling

533.23 Engine Mounting Attachments and Structure

533.25 Accessory Attachments

533.27 Turbine, Compressor, Fan and Turbosupercharger Rotors Overspeed

533.28 Engine Control Systems

533.29 Instrument Connection

Subchapter C - Design & Construction: Reciprocating Aircraft Engines

533.31 Applicability

533.33 Vibration

533.34 Turbocharger Rotors

533.35 Fuel &Induction System

533.37 Ignition System

533.39 Lubrication System

Subchapter D - Block Tests Reciprocating Aircaft Engines

533.41 Applicability

533.42 General

533.43 Vibration Test

533.45 Calibration Tests

533.47 Detonation Test

533.49 Endurance Test

533.51 Operation Test

533.53 Engine System and Component Tests

533.55 Teardown Inspection

533.57 General Conduct of Block Tests

Subchapter E - Design and Construction Turbine Aircraft Engines

533.61 Applicability

533.62 Stress Analysis

533.63 Vibration

533.64 Pressurized Engine Static Parts

533.65 Surge &Stall Characteristics

533.66 Bleed Air System

533.67 Fuel System

533.68 Induction System Icing

533.69 Ignitions System

533.70 Engine Life-limited Parts

533.71 Lubrication System

533.72 Hydraulic Actuating Systems

533.73 Power or Thrust Response

533.74 Continued Rotation

533.75 Safety Analysis

533.76 Bird Ingestion

533.77 Foreign Object Ingestion - Ice (amended 2001/03/05)

(b) Reserved.

533.78 Rain and Hail Ingestion

533.79 Fuel Burning Thrust Augmentor

Subchapter F - Block tests Turbine Aircraft Engines

533.81 Applicability

533.82 General

533.83 Vibration Test

533.84 Engine Overtorque Test

533.85 Calibration Tests

533.87 Endurance Test

533.88 Engine overtemperature test

533.89 Operation Test

533.90 Initial Maintenance Inspection

533.91 Engine System and Component Tests

533.92 Rotor Locking Tests

533.93 Teardown Inspection

533.77 Foreign Object Ingestion - Ice
(amended 2001/03/05)