---
title: "AMC E 810 Compressor and Turbine Blade Failure — CS-E · Avioverse"
description: "ED Decision 20187/014/R (1) General (a) Compliance with the specifications of CS-E 810(a) may be shown in accordance with either (i), (ii) or (iii) - (i)…"
url: "https://avioverse.aero/regulations/easa/cs-e/subpart-e/amc-e-810-compressor-and-turbine-blade-failure"
---

# AMC E 810 Compressor and Turbine Blade Failure

Engines (CS-E) · CS-E · EAR revision 7 Dec 2023

AMC Acceptable means of compliance

## AMC E 810 Compressor and Turbine Blade Failure

ED Decision 20187/014/R

(1) General

(a) Compliance with the specifications of CS-E 810(a) may be shown in accordance with either (i), (ii) or (iii) -

(i) by compliance with the tests detailed in (2) and (3),

(ii) by presentation of adequate evidence that substantiates the strength of the Engine either by blade Failure experience with Engines agreed by the Agency to be of comparable size, design and construction, or by blade Failures which have occurred during the development of the Engine, provided that the conditions of Engine speed, shutdown period, etc., are sufficiently representative;

(iii) by other evidence acceptable to the Agency.

(b) Tests for containment are detailed in (2) and those for running following blade Failure are detailed in (3), but where the most critical blade from the point of view of blade containment is the same as that for the subsequent out-of-balance running, it is acceptable to combine the tests of (2) and (3).

(c) In order to comply with CS-E 810(a), the threat represented by any blade Failure must be addressed. Therefore, the applicant should assess other possible blade Failure conditions (e.g. blade released at a different angular position, partial blade Failure, or release at speeds below the maximum to be approved) to ensure that the conditions of paragraphs (2)(c)(i) and (2)(c)(ii) below would be satisfied for any blade Failure. It is recognised that limitations on prediction capabilities exist, particularly in relation to the prediction of axially released debris. Engineering judgement based on available test and service experience may be used to evaluate these threats.

(d) When assessing the potential Hazardous Engine Effect resulting from other damage before Engine shutdown, the applicant should consider the long-term effects (e.g. unbalance loads) of blade Failures which would not be detected by the declared instrumentation, such as vibration detectors. If reliance is placed on Engine shutdown by the flight crew to prevent propagation to Hazardous Engine Effect, then appropriate means should be provided to ensure flight crew action and adequate instructions are required in the Engine manuals required by CS-E 20(d).

(2) Containment

(a) General. Containment tests should be made, either -

(i) on the complete Engine; or

(ii) on the individual stage concerned with the adjacent stators, where -

(A) the actual strength of casing under the anticipated operation conditions (e.g. temperature and pressure) is taken into account; and

(B) adequate evidence is available such as to indicate that the aircraft would not be endangered by the effect of the blade Failure on subsequent blade rows.

(b) Test Conditions. Separate tests on each compressor and turbine stage adjudged to be most critical from the point of view of blade containment (account being taken of blade size, material, radius of rotation, Rotational Speed and the relative strength of the adjacent Engine casing under operating temperature and pressure conditions) should be carried out in accordance with the conditions of (i) and (ii).

NOTE: Where the Engine design is such that potentially Engine Critical Parts overlie the compressor or turbine casing (e.g. bypass Engines, or reverse flow Engines where the combustion systems may be outside the rotors), consideration should also be given to possible hazardous internal damage caused by blades penetrating the rotor casings, even though they are contained within the external geometry of the Engine. Consideration should also be given to AMC E 520(c)(1) paragraph (2).

(i) Number of blades to be detached. One blade should be released at the top of the retention member.

(A) For composite fan blades only, the fan blade may be released at the inner annulus flow path line, provided that:

(1) tests and analyses, or other methods that are acceptable to the Agency, substantiate that the minimum material properties of the fan disk and fan blade retention system can withstand, without Failure, a centrifugal load that is equal to two times the maximum load that the retention system could experience within the approved Engine operating limitations;

(2) a procedure that is approved by the Agency is used to establish an operating limitation that specifies the maximum allowable number of start-stop stress cycles for the fan blade retention system. The life evaluation should include the combined effects of high-cycle and low-cycle fatigue. If the operating limitation is less than 100 000 cycles, this limitation should be specified in the Airworthiness Limitation Section of the Instructions for Continued Airworthiness;

(3) the effects of in-service deterioration, manufacturing variations, minimum material properties, and environmental effects should be accounted for during the tests and/or analyses.

NOTE: The fan blade retention system includes the portion of the fan blade from the inner annulus flow path line inward to the blade dovetail, the blade retention components, and the fan disk and fan blade attachment features.

(ii) Engine Conditions at Release. The blade should be released at either

(A) the maximum rotational speed to be approved ( including the maximum Engine rotor shaft over-speed transient approved under CS-E 740(f)(5), but not including the maximum Engine over-speed inadvertent transient approved under CS-E 830) and the associated maximum casing temperature; or

(B) any likely combination of the non-transient rotational speed, intake temperature and casing temperature (including conditions associated with the maximum Engine exhaust gas over-temperature transient approved under CS-E 740(f)(5), but not including the maximum Engine exhaust gas over-temperature inadvertent transient approved under CS-E 870) that is considered to be more critical.

NOTE: Any deficiency in the required casing temperature may be compensated for by means of a suitable increase in the Engine speed.

(c) Condition after Tests. On completion of the tests, a complete Engine shutdown is acceptable, but there should be :

(i) radial containment by the Engine and no significant rupture or hazardous distortion of the Engine outer casing , and no unsafe condition from possible internal damage to the Engine as a result of blades penetrating the rotor casings, even though they are contained within the external geometry of the Engine;

(ii) no other Hazardous Engine Effect resulting from the blade Failure, including due to debris being released from the Engine, forward, rearward, or otherwise outside the containment structure, unless the Hazardous Engine Effect can be shown to occur at a rate not in excess of that defined as Extremely Remote (refer to AMC E 510). Note (1): The approximate size and weight of debris released during the test, along with an estimate of the trajectory and velocity, should be recorded to determine whether the debris could result in a Hazardous Engine Effect. This data should be documented in the Manuals required by CS-E 20(d). Note (2): The above assessment is required to demonstrate that the likelihood of a Hazardous Engine Effect due to blade Failure is low enough to be accepted for Engine certification (i.e. Extremely Remote). Additional considerations may be applied during aircraft certification to further mitigate the potential effects of blade Failures at the aircraft level. Note (3): A Hazardous Engine Effect following a blade Failure may be found acceptable under CS-E 810(a), providing it can be shown that it occurs at a rate not in excess of that defined as Extremely Remote. However, the applicant must ensure that this Engine Effect is acceptable for compliance with other specifications of CS- E. For instance, if a blade Failure results in a shaft failure, this would be unacceptable under CS-E 850(a)(1) and/or CS-E 840(c), even though it is shown to occur at a rate not in excess of that defined as Extremely Remote.

(3) Running Following a Blade Failure

(a) The tests should be conducted on a complete Engine, mounted in such a manner that the reactions induced by the out-of-balance on the Engine carcass and mounts will be representative of those which would occur in the installed condition. Alternatively, tests may be carried out on a rig but consideration should be given to the effects of shaft power input, further subsequent damage, heavy out-of-balance forces on other parts of the Engine, possible shaft Failure, etc., when interpreting the test results as being indicative that no hazardous damage would occur in a complete Engine.

(b) Test Conditions. Separate tests should be carried out on each compressor and turbine stage that is adjudged to be most critical from the point of view of Engine damage subsequent to blade Failure as a result of out-of-balance forces that exist during the period prior to the Engine shutdown.

(i) The Engine should be run, with an out-of-balance that is representative of the loss of a blade from the top of the retention member, at the maximum rotational speed to be approved (other than the Maximum Engine Over-speed) until either the Engine stops of its own accord, or a period of at least 15 seconds has elapsed.

(ii) During the run, the power setting should not be altered.

(c) Condition after Tests. On completion of the tests, the result should be such that there is no unsafe condition . A complete Engine shutdown is permitted.

\[Amdt No: E/1]

\[Amdt No: E/5]

AMC — CS-E · CS-E Easy Access Rules · EAR revision 7 Dec 2023

[All rules in SUBPART E – TURBINE ENGINES: TYPE SUBSTANTIATION](https://avioverse.aero/regulations/easa/cs-e/subpart-e)

Consolidated from the EASA Easy Access Rules (revision 7 Dec 2023, extracted 13 Oct 2026) for convenience. Not the official publication — verify against the Official Journal of the European Union and the EASA publications before operational use.