U.S. Department
of Transportation

Federal Aviation

Administration 

# ADVISORY
CIRCULAR

Subject: T urbojet, Turboprop, Turboshaft, and Turbofan Engine Induction System Icing and Ice Ingestion Date: 1 0 /22/14 Initiated by: A NE -110 AC No: 20

- 147A

This a dvisory circular ( AC) pr ovides gui dance f or de monstrating compliance w ith the e ngine induction system i cing and engine i nstallation ice r equirements of T itle 14 Code of F ederal Regulations ( 14 CFR) pa rts 23, 25, 27, 29, and 33. This A C di scusses t he i cing environments depicted in appendices C a nd O of pa rt 25, appendix C of pa rt 29, and appendix D of pa rt 33.

This AC d iscusses turboshaft engine in stallations, b ut not the r o tary w ing a ircraft they a re installed on.

If you have suggestions for improving this AC, you may use the Advisory Circular Feedback

form at the end of this AC.Contents Paragraph Page

1. P urpose.............................................................................................................................. 1

2. Applicability ..................................................................................................................... 1

3. Cancellation ...................................................................................................................... 2

4. Related Regulations.......................................................................................................... 2

5. References and Related Reading Material ........................................................................ 2 6. Background....................................................................................................................... 4

7. Definitions ......................................................................................................................... 5

8. Discussion ......................................................................................................................... 8 9. Induction System Icing and Operation Test..................................................................... 11

a. Critical Point Analysis ............................................................................................ 11

b. Test Versus Analysis .............................................................................................. 11

c. Test Facilities ......................................................................................................... 12

d. Elements of CPA .................................................................................................... 12

e. Test Point(s) Selection ........................................................................................... 14

f. Table 1 of § 33.68 Icing Conditions...................................................................... 15 g. Section 33.68(d) – Ground Operation................................................................... 17

h. Ground Idle Demonstrations................................................................................. 18 i. Ice Crystal Icing .................................................................................................... 19 j. Test Setup Considerations..................................................................................... 20

k. Direct Connect Inlet Engine Test Facilities .......................................................... 20

l. Testing in Mixed Phase or Glaciated Atmospheric Conditions............................ 20

m. Test Results and Compliance Issues ...................................................................... 21

n. Mechanical Damage .............................................................................................. 21 o. Engine Systems...................................................................................................... 23

p. Auto

- Recovery Systems ......................................................................................... 23

q. Operation Instructions ............................................................................................ 24

r. Special Considerations for Mixed Phase and Glaciated Atmospheric Conditions............................................................................... 24 s. Comparative Analysis Guidance for Ice Crystal Icing ........................................... 25

t. C ompliance Considerations.................................................................................... 28

u. Ice Crystal Accretion Explanation.......................................................................... 28

v. Ice Crystal Icing Assessment .................................................................................. 28

w. Appendix O of Part 25, Water Impingement.......................................................... 29

10. I ce Slab Ingestion .............................................................................................................. 33

a. Why the Ice Slab Ingestion Test is Conducted....................................................... 33 b. Compliance Considerations..................................................................................... 33

c. Validated Analysis with Equivalent Soft Body Tolerance Testing........................ 34

d. Elements of a Validated Analysis ........................................................................... 34

e. Ice Slab Break Up .................................................................................................... 36

f. Test Results ............................................................................................................. 37 g. Communication of Test Results.............................................................................. 3811. Induction System Icing Protection

...................................................... ............................. 38

a. Acceptable Mean s of Compliance

.............................................. ............................. 38

b. Inlet Design Point Selection

.................................................................................... 39

c. Addressing Specific Operational Co nditions

.......................................................... 40

d. Two

- minute Delayed Selection of Inlet Anti

- Ice Accretion Analysis

.................... 41

e. Extended Operations Performance Standards

............................ ................ ............. 41

f. Potential Airframe Ice Sources, Including Inlets

.................................................... 4 1

g. Wing Sourced Ice for Aft Mounted Engines

.............................. ............................. 42

h. Aft Mounted En gine Installations

........................................................................... 4 3

i. Compliance with Mixed Phase and Glaciated Atmospheric Conditions under § 25.1093(b)(1)

............................................................ ..... ............................. 43 j. Compliance with Supercooled Large Drop Conditions under § 25.1093(b)(1)

...... 4 4

k. Natural Icing Flight Tests

........................................................................................ 4 4

l. I dentification of Ice Source

........................................................ ............. ................ 45

m. Icing Test Point Monitoring

....................................................... ............................. 45

n. Compliance

... .............................................................................. .................. ........... 45

o. Falling and Blowing Snow

...................................................................................... 4 6

p. Test Results

....... ...................................................................................................... 4 7

q. Compliance for Auxiliary Power Unit Installations

................... ............................. 48

12. Conclusion

............................ ............................................................................................ 49

A ppendix A

............................................................................................................................. .. A1

- 11. P urpose

.

a. This advisory circular (AC) provides guidance for demonstrating compliance with the engine induction system icing and engine installation ice requirements of Title 14 Code of Federal Regulations (14 CFR ) parts 23, 25, 27, 29, and 33. This AC dis cusses turboshaft engine installations, but not the rotary wing aircraft they are installed on. See AC s 20

- 73A

, AC 27

- 1B,

and AC 29

- 2C for additional guidance on engine installation icing issues for rotary wing applications.

b. This AC, 20

- 147A, r eplaces AC 20

- 147, dated February 2, 2004

. A pplicants should us e this AC as their primary guidance concerning icing issues for engines and engine installations for parts 25 and 33

, other guidance on the subject notwithstanding

. For part 23 airplanes , AC 23

- 16A takes precedence on engine installation icing compliance guidance. For part s 27 and 29 rotorcraft, ACs 27

- 1B, and 29

- 2C, respectively, take precedence on engine installation guidance ( that is, § § 27 .1093 and 29.1093 )

. A dditionally

, AC 20

- 73A conta ins useful information on the understanding and characterization of the icing environment

, b ut it should not be viewed as guidance on engine icing methods of compliance discussed in this AC. It should be noted that

neither supercooled large droplet ( SLD ) ( refer to appendix O of p art 25) nor Ice Crystals ( refer to appendix D of p art 33) icing environments are currently required for turboshaft engines and their

aircraft installations. It should also be noted that the SLD (refer to appendix O of part 25) ici ng environment is currently not required for turbine engine in duction systems (refer to § 25.1093)

on part 25 airplanes with a maximum takeoff weight equal to or greater than 60,000 pounds , or for part 23 airplanes

.

2. A pplicability

.

a. The guida nce provided in this document is directed to engine and airframe manufacturers, modifiers, and foreign regulatory authorities.

b. This material is neither mandatory nor regulatory in nature

, and it does not constitute a regulation. It describes a cc eptable means, but not the only means, for demonstrating compliance with the applicable regulations. The FAA (we) will consider other methods of demonstrating compliance that an applicant (you) may elect to present. Terms such as “should,” “shall,” “may ,” and “must” are used only to ensure the applicability of this particular method of compliance when the method in this document is used. This guidance is derived from extensive FAA and industry experience in determining compliance with the relevant regul ations. If we determine that following this AC would not result in compliance with the applicable regulations, we will not be bound by the terms of this AC, and we may require additional subst antiation as the basis for finding compliance.

c . This m aterial does not change, create any additional, authorize change s in, or permit deviations from existing regulatory requirements.

3. Cancellation . AC 20

- 147, Turbojet, Turboprop, and Turbofan Engine Induction System Icing and Ice Ingestion, dated 2/2/04 , is canceled.4 . R elated Regulations

.

a. Part 23, Airworthiness Standards: Normal, Utility, Acrobatic, and Commuter Category Airplanes, §§ 23.901(d)(2), 23.1093, and 23.1419.

b. Part 25, Airworthiness Standards: Transport Category Airp lanes, §§ 25.1091, 25.1093

, 25.1419 and 25. 14 20

.

c. Part 27, Airworthiness Standards: Normal Category Rotorcraft, §§ 27.1093.

d. Part 29, Airworthiness Standards: Transport Category Rotorcraft, §§ 29.1093.

e . Part 33, Airworthiness S tandards: Aircraft Engines, §§ 33.68, 33.77(c), 33.77(e), 33.89(b), and 33.78

.

5 . R eferences and Related Reading Material

.

a. AC 20

- 73 A

, “ Aircraft Ice Protection ” , dated 8/16/200 6

.

b. AC 23

- 16A, “ Powerplant Guide for Certification of Part 23 Airplanes and Airships ” , dated February 23, 2004

.

c. AC 23.1419

- 2D, “ Certification of Part 23 Airplanes for Flight in Icing Conditions ”

, dated April 19, 2007.

d . AC 27

- 1B, “ Certification of Normal Category Rotorcraft ”

, dated September 30, 2008.

e . AC 29

- 2C, “ Certification of Transport Category Rotorcraft ”

, dated September 30, 2008.

f . AC 33

- 2B

, “ Aircraft Engine Type Certification Handbook ”, dated June 30, 1993

.

g . U.S. Department of Transportation. Federal Aviation Ad ministration. Report No. FAA

RD

- 77

- 76

, “ Engineering Summary of Powerplant Icing Technical Data ”

, by G. D. Pfeiffer and

G. P. Maier , dated July 1977

.

h

. U.S. Department of Transportation. Federal Aviation Administration. Report No. FAA

RD

- 77

- 78

, “ Engineering Summary of Powerplant Icing Technical Data ”

, dated July 1977.

i . U.S. Department of Transportation. Federal Aviation Administration. Report No.

DOT/FAA/AR

- 97/66, August, 1998. “ Snow and Ice Particle Sizes and Mass Concentrations at

Al titudes Up To 9 km (30,000 ft) ” , by R. K. Jeck.

j . U.S. Department of Transportation. Federal Aviation Administration. Report No.

DOT/FAA/AR

- 09 / 13

, “ Technical Compendium from Meetings of the Engine Harmonization Working Group ”

, by R. Mazzawy , date d April 2009

.k . U.S. Department of Transportation. Federal Aviation Administration. Report No.

DOT/FAA/AR

- 03/48, MAY 2003. “ Assessment of Effects of Mixed

- Phase Icing Conditions on Thermal Ice Protection Systems ”

, by Kamel Al

- KHalil

.

l

. R oyal Aircraft Establishment (Farnborough) Technical Note No: Mech. Eng. 283

. “ The Analysis of Measurements of Free ice and Ice/Water Concentrations in the Atmosphere of Equatorial Zone ”

, by Ian I. McNaughton, B.Sc., Dip. R.T.C

. , dated 1959

.

m

. “ The Icing of an Unheated Non

- rotating Cylinder in Liquid Water Droplet

- Ice Crystal

Clouds ”

, by E. P. Lozowski, J. R. Stallabrass & P. P. Hearty, National Research Council Canada Report LTR

- LT

- 96, dated February 1979.

n

. “ Further Icing Experiments on an Unheated Non

- rotating Cylinder ”

, LTR

- LT

- 105, by J.R. Stallabrass and P. F. Hearty, dated November 1979

.

o

. “ Snow Concentration Measurements and Correlation with Visibility ” , by J. R. Stallabrass, AGARD Conference Proceedings No. 236 Icing Testing f or Aircraft Engines , dated April 1978

.

p

. “ The Estimation of Snowfall Rate Using Visibility ” , by Rasmussen, R.M., J. Vivekanandan, J. Cole, B. Myers, and C. Masters, 1999: Journal of Applied Meteorology, 38,

1542

- 1563.

q

. “ Calibration and Acc eptance of Icing Wind Tunnels, Aerospace Recommended Practice (ARP) 5905 ” , issued 2003

- 09: Society of Automotive Engineers.

r. “Appendix D

– An Interim Icing Envelope”, by R. S. Mazzawy and J. W. Strapp, SAE

2007

- 01

- 3311, dated September, 2007.6

. B ackground

.

a. The i nduction s ystem i cing requirements of §§ 23.1093, 25.1093, 27.1093, 29.1093 and of § 33.68 are to provide protection for flight into icing with no adverse effect on engine operation

or power

. Propulsion systems that meet these requirements when operated in accordance with

the aircraft flight manual, have generally demonstrated safe operation upon expos ure to natural icing environments.

b. The successful demonstration of the test conditions outlined in the regulation is intended to address many potential engine power conditions . These conditions include aircraft flight conditions, and environmental conditions that could otherwise prove to be costly and difficult to realistically test.

c. A direct comparison betwe en the severity of compliance requirements for icing versus compliance requirements for rain and hail is in appropriate

. In § 33.68, the regulation references

a ppendices C and O of part 25

, and a ppendix D of part 33

. T he environmental threat s depicted

in all of these a ppendices have occur r ed

. However , operators have also encountered icing

conditions that are more severe

. Indeed, t he probable occurrence rate of encountering the icing

conditions depicted in a ppendix C of part 25 or 29 is far greater than t he remote threat posed by

the rain and hail environmental conditions depicted in a ppendix B of part 33

. As a result of

t hese differences in rates of occurrence, as well as fundamental differences of environment and corresponding assumed in

- flight profiles

, a comparison of compliance requirements across different threat types will be flawed , incongruous, and ultimately inappropriate

.

d

. The icing environments described in this AC are considered probable to encounter. Appendices C and O of part 25 an d appendix C of part 29

, address 99% of supercooled droplet icing conditions ; they are based on atmospheric icing data

. Additionally, the probability of exceeding both liquid water c ontent ( LWC ) and m edian v olume d iameter (MVD) while operating at or below the temperature specified in both appendi ces C

, of parts 25 and 29 is

estimated at approximately 10 - 3 encounters for each flight ; thus a probable encounter. Appendix

D of part 33 is derived from the available flight data ( see reference in paragraph 5.j

. of this AC ) and represents a 99% ice crystal condition based on the current data.

e

. Finally, t he icing environments depicted in each of the referenced appendices define single cloud lengths in terms of 17.4 nautical miles , while nature doesn’t prov ide icing clouds so discretely. E xperience has shown that actual icing environment can be a combination of

conditions . For example

, as depicted in appendix C of part s 25 and 29

, a continuous maximum cloud followed by an intermittent maximum cloud

, follow ed by a continuous maximum cloud,

and so on , is probably what a natural environment demonstrates

. To account for this difference between natural icing environments and the environments depicted in the appendices, this guidance effectively provides an equi valent icing condition that is representative of multiple clouds in sequence.

f

. The body of this AC is arranged in three areas corresponding to the applicable parts (that is, §§ 33.68, 33.89(b), 33.77, 23.1093, 25.1093, 27.1093, and 29.1093 ). The first area addresses part 33 induction system icing compliance guidance ( § 33.68 compliance) , in paragraph 9 of thisAC . The second area addresses engine ice slab ingestion compliance guidance ( § 33.77 compliance) , in paragraph 10 of this AC . The third area addresses installation icing requirements under parts 23, 25, 27, and 29 , in paragraph 11 of this AC

.

g. The icing conditions depicted in a ppendix O of part 25

, and a ppendix D of part 33 are not applicable to turboshaft engines or their install ations. T urboshaft engines need only comply with the icing conditions depicted in appendix C of part 29.

h. The icing conditions depicted in appendix O of part 25 are not applicable to turbine engine induction systems on part 25 airplanes with a max imum takeoff weight equal to or greater than 60,000 pounds , or on part 23 airplanes

. Turbine engine induction systems on part 25 airplanes with a maximum takeoff weight equal to or greater than 60,000 pounds need only comply with the icing conditions depi cted in appendix C of part 2 5 , appendix D of part 33

, and in falling and blowing snow within the limitations established for the airplane for such operation

.

7 . D efinitions . The following are defined for the purpose of this AC :

a. Auto

- R ecovery S ystems

. E ngine systems that ensure that engines operate just before or immediately after a n upset ( t hat is, power loss or stall) without operator intervention. Auto

recovery systems include auto

- relight systems, stall recovery systems, a nd other engine s ystem intended to recover the operability of an engine following a flameout, surge, stall, or a

combination of these.

b. Freezing F raction

. T he ratio or percentage of water that impacts a surface and freezes.

The freezing fraction is defined as a number between 0 and 1, and will determine the type of ice

formation.

c. Cloud Extent Factor

. T he distance that a cloud extends vertically (vertical extent) or horizontally (horizontal extent). Vertical extent is normally measured in feet while ho rizontal extent is measured in nautical miles. The cloud extent factor is a dimensionless number, which

uses the length of a cloud to determine an average liquid water content across the cloud.

d. H ighlight A rea

. T h e area bounded by the leading ed ge of the nacelle inlet. This may be different for turboshaft installations where complex inlet schemes are utilized.

e

. Ice F ormations

. These result from the impact of supercooled water droplets on propulsion system surfaces ; formations are class ified as follows:

(1) Glaze I ce

. A clear, hard ice, which forms at temperatures close to (but below) freezing, in air with high liquid water content and large droplet sizes. Droplets impacting the surface do not freeze immediately, but run bac k along the surface until freezing occurs. Glaze ice typically has a non

- aerodynamic shape and is more susceptible to aerodynamic forces that result in shedding. Glaze ice typically has both a lower freezing fraction and lower adhesive properties than ri me ice. Glaze ice is often a concern for static hardware while rime ice is often a concern for rotating hardware.(2) Rime I ce

. A milky white ice which forms at low temperatures, in air with low liquid

water content and small droplet sizes. R ime ice typically forms in an aerodynamic shape, on both rotating and static engine hardware. The freezing fraction is high for rime ice, typically approaching a value of 1.0. Rime ice typically has greater adhesion properties than glaze ice but

often a lower density. Adhesion properties increase with lower temperature up to a test point where no additional adhesion is gained with additional lower temperature.

(3) Mixed or I ntermediate I ce

. A combination of glaze and rime ice which forms with rime patches slightly aft of the glaze ice portions. This ice forms at temperatures, liquid water contents

, and droplet sizes between those that produce rime and glaze ice.

f

. Ice S hed C ycles

. T he time period required to build up and then shed ic e on a propulsion system surface for a given power and icing condition. A shed cycle can be identified visual ly (for example, with high

- speed camera s , and engine instrumentation such as vibration pickups, temperature probes, pressure probes, speed pickups , etc. ). F or rotating surfaces, such as fan blades, t he ice shed cycle is strongly influenced by rotor speed and the adhesive strength of ice to

the surface. In general, i ce adhesive strength increases as surface temperature decreases

.

g

. Icing C o nditions

. These m eteorological condition s are defined by the following parameters:

(1) Liquid Water Content (LWC)

. T he c oncentration of liquid water in air, typically expressed in grams of water for each cubic meter of air.

(2) Medi an Volume Diameter (MVD)

. T he drop diameter which divides the total water volume present in a droplet distribution in half ( that is, half the water volume is in larger drops

a nd half the volume in smaller drops )

. Note the MVD used in a ppendix O of part 2 5

, and

a ppendix D of part 33

, is equivalent to the MED in a ppendix C of part s 25 and 29 for an assumed

L angmuir type droplet distribution.

(3) Mean Effective Diameter (MED)

. Similar to MVD. See definition of MVD.

( 4 ) Median Mass Dim ension (MMD)

. T he particle size (sphere of equivalent mass) which divides the total ice mass present in an ice particle distribution in half (that is, half the ice mass is in larger particles and half the ice mass is in smaller particles )

.

( 5 ) Total Temperature

. T he ambient temperature plus the ram rise. For icing testing in test cells, the total engine inlet temperature includes static temperature of the cloud as depicted

in a ppendices C and O of part 25 , appendix C of part 29 , and a ppendix D of part 33, plus the assumed flight airspeed.

( 6 ) Static Temperature

. T he ambient temperature calculated from the local measured total

temperature on the aircraft

, minus the temperature rise from velocity effects. Appendices C and

O of part 2 5 , appendix C of part 29 , and a ppendix D of part 33 temperatures are assumed to be static ambient temperatures.h

. Power L oss I nstabilities

. These are caused by e ngine operating anomalies

. These types of

anomalies could include non

- recoverable or repeating surge, stall, rollback, or flameout, which can result in engine power or thrust cycling

. A rotating compressor stall is one kind of power loss instability that might be acceptable, upon review with the FAA : if it is not noticeable to the flight crew
, and if it does not result in unacceptable stresses or operability effects either at steady

- state conditio ns
, or during acceleration

.

i

. Scoop F actor (concentration factor )

. T he ratio of the nacelle inlet highlight area (A H ) to

the area of the captured air stream tube (A C ) [scoop factor = A H /A C ]. Scoop factor compares the liquid water available for ice formation in the engine inlet

, to that available in the low

- pressure

c ompressor or engine core

, as a function of aircraft forward airspeed and engine power condition. The scoop factor effect depends on the droplet diameter, the simulated airspeed

, and the engine power level

, as well as the geometry and size of the engine

. This may be different for turboshaft installations where complex inlet sc hemes are utilized.

j

. Serious L oss of P ower or T hrust

. A non

- recoverable or repeating engine surge, stall,

rollback

, or flameout occurrence , which can result in noticeable engine power or thrust loss.

k

. Stabilized Operation

. This is w hen the engine d emonstrate s steady, reliable, and smooth operation while operating at the test condition ( that is, during multiple ice build or shed cycles, if ice is accreting), and during throttle or power transients. Engine operation is considered

stabiliz ed when measured engine parameters are not changing, or when a regular repeatable shed cycle has been demonstrated through the recording of measured engine parameters with

s tabilized i ce a ccretions

.

l

. Stabilized Ice A ccretion

. A condition w hen ice accretion is not increasing on any engine

part

, or when the accreting ice has demonstrated a regular shed cycle when viewed by a video

camera or instrumentation indication

.

m

. Sustained P ower L oss

. A permanent loss in engine power or thrust. Typ ically

, sustained power loss is calculated at rated takeoff power. See definitions below on “ Temporary Power

Loss ” and “ Momentary Power Loss ” for p ower or thrust losses that are not sustained

.

n

. Descent Idle Engine Speed

. An altitude

- dependent

, m inimum flight

- idle engine speed that

is generally in effect from the top of descent ( TOD ) to where the approach phase of flight begins

. This is typically when flaps are deployed and idle speed increases for an approach

. Since this is

different for turbos haft installations, the applicant can propose appropriate criteria.

o

. Momentary P ower L oss

. A short

- duration reduction in engine power or thrust associated with a transient event ( for example

, ice shedding).

p

. Noticeable

. A tactile feel ing during an event

, or the use of typical engine test instrumentation or flight deck instrumentation (such as, N 1 , N 2 , vibrations, exhaust gas temp) to
indicate an event

.

q

. Temporary P ower L oss

. A reduction or loss in engine power or thrust occurrin g during an icing encounter ; this may be related to the effects of ingesting supercooled water or ice particles,or possibly the effects of ice accumulation within the flowpath. The amount of temporary power

loss should be communicated to the installer fr om the engine manufacturer.

r. Power rollback

. A n uncommanded reduction in engine rotor speeds

, and subsequent loss of power and power control.

8

. D iscussion

.

a. An icing encounter, including a prolonged encounter, should neither cause a s ignificant

increase in workload

, n or elicit concern from the flight crew

. Additionally, an icing encounter

should not result in damage to the engine

, its systems or subsystems, loss of performance or an operability deficiency

, or any other compromise of c ompliance with the engine certification

basis

. Relevant changes in bill of material hardware or software made during the compliance program should be re

- evaluated with respect to icing.

(1) The engine should have sufficient durability to operat e through prolonged or repeated

environmental icing encounters without special operational or maintenance interventions. For example, o perational procedures to assist ice shedding, such a s t hrottle manipulation or power change , should not be used to compl y with the in

- flight icing requirements of parts 23, 25, 27, 29

and 33.

(2 ) The applicant may provide procedures for engine throttle manipulation or power change (for example, power run

- ups) to shed ice accumulated during ground operations. The se procedures will be used as in

- service ground operation recommendation s , although they would

be mandatory if used during the compliance demonstration of §§ 33.68, 23.1093, 25.1093

, 27.1093, and 29.1093

.

(3) A pplicant s should ensure they collec t sufficient data to demonstrate regulatory compliance . V ideo or photographic coverage may be useful for supplement ing test results

. In

addition

, applicant s should determine the most critical measured parameters which indicate ice accretion and shedding within the engine. These parameters provide critical data for show ing regulatory compliance during the compliance plan . In addition they help validate the critical point analysis ( CPA ) assumptions and demonstration intent . The parameters may include both

visual and instrumented indications that need to be monitored during the icing test to show acceptable operation in icing conditions

.

( a ) This demonstration should include s tabilized i ce a ccretion (that is, s tabilized

o peration) with eithe r no ice buildup or no additional ice buildup on the engine or nacelle inlet. Normal engine control system responses during the ice accumulation process (for example, isochronous control response s to accreting ice) are acceptable if no sustained power los ses are encountered

. Any significant temporary power loss during operation in icing conditions should be reported to the installer by the engine manufacturer.

b. Mixed Phase or Glaciated Atmospheric Conditions

. The FAA and the aviation community

h ave recently become aware of the potential effects of m ixed phase icing conditions . They occur when mixtures of supercooled liquid water droplets and ice particles (mixed phase) or icecrystals alone (glaciated) exist in a cloud, often around the outskirt s of a deep convective cloud formation. Ice crystal icing condition s exist when all the liquid water particles in the cloud have frozen into ice particles. Mixed phase and ice crystal icing has caused more than 100 turbine engine power losses. Complianc e methodologies for these conditions are being introduced in

this AC for the first time for turbojet and turbofan engines

. These conditions do not apply to turboshaft engines.

(1) T raditionally

, i n dustry has c onsidered the susceptibility of tur bine engines to mixed phase or ice crystal conditions as minimal ly consequential

, with two exceptions: (1) pronounced inlet bends (such as particle

- separator inlets) or inlet flow reversals, and (2) high solidity, high turning front stage compressor stator s

. Industry has recognized t he inlet bends are a concern since that is where inlet flow can stagnate and accumulate ice . Similarly

, industry has recognized that h igh solidity, high turning front stage compressor stators are a concern because

stator airfo ils are susceptible to non

- aerodynamic ice buildup

. These flowpath blockages and significant turns can result in accretion zones and core airflow blockage from ice accretion

.

(2) Recently, operators have reported service experience on fixed win g applications with icing within the engine core stream due to mixed phase and glaciated atmospheric conditions

. These reported s ervice events include d ice crystal accretion within the engine ’ s flowpath and within engine inlet probes, including temperatur e and pressure probes. This ice accretion has produced significant adverse impacts, including uncommanded engine power rollback or

flameout

, with occasional core hardware damage.

(3) Operators have also reported t otal a ir t emperature (TAT) pro be malfunction s during many of these engine icing events

. A h eated TAT probe malfunction is a known indicator of ice crystals. Ice accumulation within the TAT probe produces a false TAT of air temperature near 32 0 Fahrenheit ( 0 0 degrees Celsius )

. Ice de tection systems should be evaluated for these

conditions.

(4) Events attributed to i cing within the engine core stream due to mixed phase or ice crystal conditions , and TAT probe malfunctions, suggests that ice crystal icing is of greater

concer n than originally thought

. The root cause of these events can be traced to ice buildup

within the core flowpath of the affected engine. In general, these events occur near convective clouds at ambient temperature s warmer than the international standard a tmosphere

, and outside the icing envelopes depicted in appendic es C and O of part 25

, and appendix C of part 29

.

(5) We have established app en d i x D of part 33 to define th e mixed phase or glaciated atmospheric envelope

.

c

. Auto

- R ecovery S ystems

. C ompliance with §§ 33 .68 and 33.77 is intended to demonstrate

no flameout , sustained power loss, surge or stall, or rundown will occur when operating in icing

conditions

. The a uto

- recovery provide s protection in

- service against multiple forms of power loss. However, a uto

- r ecovery protections should not be relied on during the compliance

demonstration because auto

- recovery systems are back

- up dev ices

. For additional information

about the use of auto

- recovery systems when demonstrating compliance to §§ 33.68 and 33.77, see paragraph s 9.p., and 10.f.(5) of this AC

. For additional information on auto

- recoverysystems when demonstrating compliance to §§ 23.1093, 25.1093, 27.1093, and 29.1093, s ee paragraph 11.p.(3) (a) of this AC

.

d

. Applicatio n of Cloud Extent Factors for § 33.68 Compliance Tests

. Cloud extent factors

are depicted in a ppendices C and O of part 25, appendix C of part 29

, and appendix D of part 33

. The horizontal cloud extent factor is a dimensionless number, which relates the length of a cloud to an average LWC across the cloud. It is used in the CPA to assess the probabl e icing content during various aircraft mission and performance analyses

.

(1) Typically

, the average LWC of a cloud is less for longer clouds. The horizon tal cloud

extent factor is not intended to be used to limit the severity of exposure to icing conditions,

assum ing the aircraft will be required to operate in that condition . F or example, when in a holding pattern which may require repeated passes through a severe icing environment, or continuously remain in that severe environment , the horizontal cloud extent factor should not be applied

.

(a) Engines and their inlet systems should demonstrate continuous operation in icing

condition s without regard to time. Typically

, an applicant assumes multiple clouds in their CPA,

with a cloud extent factor equal to 1.0. This is because actual cloud extent factors are not a
consideration for most engine operations, particularl y in an aircraft h old pattern. This approach

assure s proper engine and induction system operation within the atmospheric conditions depicted

in appendices C and O of part 25

, appendix C of part 29, appendix D of part 33 , and as experience indicates

, in actual icing environ ments. The following are s ome additional factors to

consider :

1

. These cloud extent factors are also applicable to airframe flight profiles . Airplane applicants us e the cloud extent factors depicted in the appendices during their eva luation of the straight

- line flight portion of the ir compliance demonstration. However , engines and induction systems have historically not been limited to or evaluated against a specific aircraft flight profile when considering icing environments. Inste ad

, they are evaluated for unlimited operation in icing.

2

. While appendix C of part s 25 and 29 define supercooled clouds

, a ppendix O of part 25 include s freezing precipitation (that is, rain and drizzle).

3

. A ppen dix D of part 33 depicts i ce crystals and mixed conditions (meaning, a mixture of supercooled water droplets and ice crystals). The cloud extent factor depicted in appendix D of part 33 relies primarily on limited data derived from the reference in paragr aph

5.k . of this AC. The data cover s several geographical regions including eastern Asia where ice crystal

- related service events are prevalent.

9

. I nduction System Icing a nd Operation Test (§§ 33.68 and 33.89(b))

.

a

. C ritical P oint A nalysis (CPA )

. A CPA is an analytical method utilizing engine test data to show an engine meets the certification requirements of part 33. This method derives critical test points fro m the collection and analysis of data on icing effects on engine performance.Comp liance with § 33.68 includes identifying, through analysis, the critical operating test points for icing within the declared operating envelope of the engine

. A CPA should include

consideration of the icing conditions depicted in appendi x C of part s 25 an d 29 , and as we gain experience

, in appendix D of part 33. A CPA should relate these a ppendices with the aircraft speed range and engine powers defined by the engine manufacturer . It should also include a

p rolonged flight operation in icing (for example, in

- flight hold pattern), or repeat icing

encounters. These combined elements within the CPA should identify the most critical operational icing conditions.

( 1 ) The applicant (you) should ensure that analysis is supported by test data. Your CP A

s hould also include environmental and engine operational effects on accumulation, accretion locations, and the most critical engine operating conditions for ice shed and ingestion. You may propose conditions outside the requirements of the a ppendices (f or example, recommending

conditions that are more severe based on actual service experience )

. You may also supplement a CPA with development test data (for example, wet and dry testing with thermocouple components).

( 2 ) Your CPA should include ice accretion calculation s that account for freezing fraction

and aerodynamic effects of the ice as it moves into the air inlet

. For example, water ingestion

into the fan module and core inlet s , water impingement rates for critical surfaces, forward

aircr aft airspeed effects, engine configuration effects such as inter

- compressor bleed, and altitude

effects ( such as bypass ratio effects )

. The CPA should also include an energy balance of critical engine surfaces ( for example, latent heat and heat of fusion effects, metal

- to

- ice heat transfer effects, and ice insulating effects )

.

( 3 ) For anti

- iced parts, the CPA should identify a critical test point determined from energy balance calculations of required heat loads , and encompassing the range of po ssible combinations of icing condition and engine power. In glaze ice conditions, assessing the effects

of non

- aerodynamic ice formations and their shedding is more complex

. FAA Report number

FAA

- RD

- 77

- 78

, titled “ Engineering Summary of Powerplant Icing Technical Data ”

, provides additional guidance on performing a critical test point icing analysis

.

b. Test V ersus Analysis

. The CPA is not meant to replace testing . Rather , it provides a means to predict critical test points . The CPA test points c an replace the standard table points

defined in Table s 1 and 2 of § 33.68

, when applicants can show they are equivalent. Otherwise they supplement the standard certification test points

. Today’s analysis tools are greatly improved over their predecessors ; h owever, experience has shown the CPA is best used to predict the critical icing conditions for a given design, then use the m in conjunction with the

standard certification test points for certificatio n c omplianc e test purposes

. Please note, even if

the FAA concur s with a type certificate holder ’ s generic CPA method , this does not constitute

our acceptance of the r esulting critical icing test points for future certification projects. The content of an icing certification program for any given certificat ion project will be evaluated on a

case

- by

- case basis.

c. Test Facilities

. Facility limitations or weather can delay § 33.68 compliance

demonstrations. These delays may often be avoided by using national and international icingtest facilities

. N ational and international icing test facilities for accomplishing engine icing compliance testing include the Unit ed States Air Force ( USAF ) McKinley Climatic Lab in Florida, the Arnold Engineering Development Center ( AEDC ) in Tennessee, and the Canada’s

N ational R esearch Council in Ottawa

, the Global Aerospace Center for Icing and Environment (GLACIER), and GE Aviation Engine Testing, Research and Development Centre (TRDC) in Winnipeg

. In Europe, there is the Delegation Generale Po ur L’Armement (DGA) Aero Engine test facility in France. Other test facilities are likely to become available in the future as test capabilities improve

.

(1) When considering any icing facility, the capability to provide simulated cloud

conditions that meet the requ irements of § 33.68 should be thoroughly considered

. Critical items

that that you should also consider are:

(a) A d escription of the icing facility

, including the spray system arrangement, spray nozzle(s), water flow system, airflow system

, and test facility operation.

(b) Instrumentation should includ e facility operation monitoring, cloud simulation

, and cloud property determination

.

(c) Air/water flow operating map

.

(d) Psych r ometrics

.

(e) Pre

- test icing rig functional checks

.

(f) Facility test procedures

.

(g) Test data measurement accuracy and capability description

.

d

. Elements of CPA

. Your CPA should address, at minimum, the following icin g issues:

(1) Ice S hed D amage

. S hed ice can cause engine damage if it impacts an engine surface with sufficient mass and velocity. The following types of damage ar e common , and you should

include them in your CPA with an assessment of each :

( a ) First Stage of Compressor (example, f an)

. Various rotating and non

- rotating parts
of the fan module

, or first stage compressor section for non

- fan engines, are susceptible to ice shed damage

. For example, a coustic panels, fan rub strips , and fan blade tips are susceptible to

ice shed from the inlet sensor(s), spinner, and the fan blade root.

( b ) Since engine icing compliance testing is performed in test cells and not actual flight conditions, t he effects of ice density, h ardness, and adhesion strength as it sheds should be assessed to realistic flight conditions. For example, in realistic flight conditions t he ice shed cycle for rotating surfaces such as fan blades

, or first stage compressor blades for non

- fan engines , is strongly influenced by the rotor speed and the adhesive strength of the ice to the surface. The adhesive strength of ice generally increases with decreasing surface temperature.The ice thickness and the rotor speed at the time of the shed

, defines the impact threat

. A vailable

data t ypically shows that shedding from unheated rotating and static parts can be quite variable and difficult to predict . This is due to uncertain factors such as adhesion strength properties, local thermodynamic properties, and ice structure (such as ice bridging). In some cases

, applicants have utilize d consistently demonstrated shedding trends in their CPA

, as opposed to

absolute values

, since there is often uncertainty in precise shed predictions.

( c ) When de termining the critical conditions for fan module damage, the surface

temperature, exposure time

, and rotor speed are important considerations

, as well as atmospheric icing condition s and scoop factor

. In particular, extended operation in a holding conditi on power level during very cold and continuous maximum icing conditions will maximize the adhesion of ice on rotating first stage compressor or fan components . This can result in large ice accretions and resulting sheds which can damage the engine or caus e power loss

.

( 2 ) Compressor D amage

. When i ce formations on static components shed

, they often result in damage. This type of damage generally occurs on the first blade set in the high

- pressure compressor ( that is, intermediate pressure compre ssor for three spool engines

, or first stage of compression for conventional turboshaft engines ). E stablishing the critical conditions for these glaze ice accretions requires careful consideration

, since they occur at specific limited conditions

of low fr eezing fractions

, over a range of local mach numbers and air densities. The critical conditions may not occur during any of the power settings discussed in this AC ( for example, flight

- idle, 50 % and 75 % of maximum continuous or 100 % maximum continuous) ;

c onsequently

, you should evaluate the power setting at the critical condition

. Finally, since icing

is a common environmental condition , you should evaluate a ny engine compressor damage that

result s from ice testing against the possibility of multiple occu rrences.

( 3 ) Engine O perability and C ompressor R ematch . Ice shed from upstream components may enter the core compressor. The presence of ice

, or water from melted ice

, in the gas path

may cause engine component cycle changes

. T he engine shoul d be capable of accelerating from minimum flight

- idle to takeoff power , and should demonstrate takeoff power set procedures on the ground at any icing condition, without unacceptable power loss or p ower l oss i nstabilities

. Ice sheds should not result in f lameout, rollback, or surge. Any anomalous engine behavior

should be raised to the FAA Engine or Aircraft Certification Office ( ECO or ACO) for

evaluation

, and if found acceptable, it should be documented in the engine’s installation manual.

Y ou should c onsider as part of the CPA

, the engine accelerations and decelerations relative to operability challenges (for example, surge and stall)

. Assume that t he minimum engine bleed schedule allowed for the condition being tested minimize s the operability margin

. CPA testing

should demonstrate those conditions where the minimum operability margin is expected.

( 4 ) Core and Booster I ce B lockage . Ice accretion on internal engine vanes from glaze ice accretions may affect air flow capacity and rematch of the engine cycle

. This should be

considered in your CPA. For engine powers that can sustain flight, you should reconcile ice accretion through a demonstration of several ice build shed cycles

. S how that there are no adverse operating effects from ice bu ilds or sheds.( 5 ) Sensor F ouling . Ice accretion and blockage of control sensors can result in erroneous engine pressure and temperature measurements. Critical sensors include inlet total pressure and temperature probes

, and inter

- compressor t emperature probes. A power loss or power loss instability could result if erroneous measurements are used by the engine control system ’ s control law to establish power or thrust ratings, or if used to schedule other systems required to operate the engine (for example, variable stator vanes). Critical sensors should be designed to
operat e when exposed to the conditions defined in 33.68 and the associated appendices, with

minimal ice accreti on and no erroneous measurement s that would result in an unacceptab le operating characteristic

, such as a power loss

. I n accordance with 33.68 icing requirements,

t urboshaft engines need only be evaluate d against the environmental icing conditions depicted in appendix C of part 29. The effects of installation on the loc al icing conditions at the probe

should be accounted for in all types of applicable installations

. Additionally, ice accretion on upstream sensors can shed and cause engine damage to downstream rotating hardware. You

should evaluate e ngine inlet probes f or supercooled droplet icing susceptibility, as well as ice crystal icing susceptibility

.

e

. T est P oint ( s ) S election . Test points selected for a supercooled droplet environment should

address the applicable icing envelopes depicted in a ppendices C and O of part 25

, and appendix

C of part 29. Test points for an ice crystal icing environment should be representative of the meteorological conditions depicted in a ppendix D of part 33 . Typically, the supercooled droplet test points include those define d in the regulation (either § 33.68 or § 25.1093) and any additional test points identified as part of your CPA. You should consider pertinent service experience as well as the anticipated use of the aircraft when selecting critical icing test points.

Do consider the following when constructing an icing test matrix:

(1) Section 33.68

, Acceptable Means of Compliance . The engine should be capable of operating acceptably under the meteorological conditions depicted in a ppendices C and O of part
2 5

, and appendix C of part 29

, covering the engine operating envelope defined in § 33.68(b)(2)(ii)(B)

, and under the conditions of ground icing defined in § 33.68( d ). Section 33.68( c ) permits eliminating specific standard certification test points so long as you choose to

run a similar CPA test condition that is more severe in terms of ice accretion mass at critical engine locations or to produce an equivalent level of safety to the standard test points

.

( 2 ) Rotorcraft Turboshaft Engines . Engine s installed on rotorcraft present their own unique conditions. Holding phase test conditions are not applicable to turboshaft applications. Therefore, turboshaft engine testing should include test conditions that address prolonged hover or continued oper ation in a relatively small area of operation. For example, prolonged operations in the local vicinity of the airport or possibly local operations around oil rig s at sea. This may require additional test, analysis, or both, to show installation configura tion dilution

benefits in prolonged operations, based on the aircraft induction system design.

f

. Table 1 of § 33.68 I cing C onditions

.

(1) The supercooled LWC shown in Table 1 of § 33.68 represent s ambient icing conditions for an open inl et ground test facility or equivalent

, within the inlet duct of a direct connect testfacility. An a nalysis of local enrichment for flight effects

, as would normally be done for a CPA

LWC

, is not necessary for the table points

.

(2) For Table 1 test conditions 1 and 2, applicants should run the engine under both rime and glaze icing conditions

. You should test under both conditions for at least 10 minutes each at
100 %, 75%, and 50% maximum continuous power
, and for 10 minutes at a flight

- idle

. F ollow

e ach power setting with a snap acceleration to takeoff power. If ice is building at the end of 10 minutes at the three higher power settings, continue running the test point until the engine demonstrates stabilized building and shedding

, or until t he engine no longer operate s satisfactorily.

(3) Table 1, c onditions 1 and 2 are intended to partially represent a broad test matrix of environmental and engine operating conditions to be used when showing compliance with § 33.68. This test mat rix includes power settings from idle to 100% maximum continuous

. They are representative of exposure to conditions typical of both high altitude where rime ice formations occur, as well as conditions typical of low altitude where glaze ice formations oft en

occur.

(a) Icing conditions are normally run for a minimum of 10 minutes for all engine powers that can sustain level flight, or longer if the natural engine ice shed cycle is not established. Special consideration and tests should be c onducted to adequately demonstrate engine inlet screens and inlet air passages that might accumulate snow or ice due to restrictions

or contours. At low power

, such as idle descent power

, the required time period is limited to 10 minutes. The rime icing condition should be run at an engine speed associated with top of descent operation for fixed wing aircraft

, as depicted in a ppendix C of part 25

. The glaze icing condition should be run at the minimum engine speed associated with lower altitudes at the e nd of the descent phase of flight for fixed wing applications

. Applicants for rotorcraft applications should propose an appropriate flight profile to address this low power test period duration.

(4) Conditions 3 and 4 ( h olding phase)

. The engi ne should continue to meet part 33 requirements while demonstrating it can operat e in definitely in a flight hold pattern under icing

conditions

. Guidance on test durations and procedures to achieve this are outlined throughout

this AC. The test program f or turbofan and turboprop applications should include test point

condition s (for example, icing condition s and power setting s ) to address the effects of prolonged exposure in icing conditions that are typical of in

- flight holding patterns.

( a ) Condition 4 in T able 1 represents a rime icing condition that is typically encountered on transport category airplanes. Condition 3 in Table 1 represents a mixed rime and glaze icing condition that was originally derived from the E uropean Union’s JAR

- E ice requirements of LWC, temperature

, and droplet size. The engine and inlet should be capable of prolonged exposure to the conditions specified in T able 1. For fixed wing aircraft, a 45

- minute test exposure will typically demonstrate several ice shed cycles. You should follow each accretion point for condition 3 and 4 with a snap acceleration to tak e off power. To demonstrate unlimited icing operation, t he engine should exhibit stabilized operation at the conclusion of the

45

- minutes test condition , p rior to the snap acceleration

.(5) For A ll C onditions . Engine operation in all icing conditions should be uneventful

, uninterrupted, and without any significant adverse effects . It should also include the ability to continue in operation and a ccelerate and decelerate normally without adverse operational effects. Some power reduction is acceptable at idle power settings in icing conditions, due to the cycle effects of pumping ice and water ; however, all other operation should be unaffected.

(a) You should determine what parameters need to be monitored to determine the

stabilized operation of the engine during the icing test. Engine operation is considered stabilized

either when measured engine parameters are not changing, or whe n a regular, repeatable shed cycle has been demonstrated through the recording of measured engine parameters

. Variations in measured parameters are acceptable during the performance of the ice test, as long as the long

- term trend (typically the duration o f several shed cycles) is stable and not trending upwards

or downwards.

( b ) E ngines with manually activated icing protection systems (including systems for probes), should be tested. You may test the se systems by stabiliz ing the engine for at least 2

minutes in the icing atmosphere (1

- minute for turbo shaft engines) with the protection systems

off (that is, prior to turning the icing protection system on ) . Conduct this test at flight

- idle and

above. Doing so simulate s the delay expected f or the pilot to recognize the icing condition.

(c) Fully automatic systems may use an appropriate delay. Systems that are automatic and controlled by the full authority digital electronic engine control (FADEC) do not require the

2

- minute delay in the ice protection system demonstration. Where the engine’s anti

- icing system relies on an ice detector to indicate the presence of icing conditions, a delay in the anti

- ice selection is likely

. Therefore, delayed selection testing should still be demonstrated.

(d) Stable engine operation should occur under the tested icing conditions

. A stable engine operation is intended to address both stabilized ice accretion and shed cycles during steady

- state engine operation. Ice accretio ns are considered stabilized when either ice is not forming on any engine part or the accreting ice has demonstrated a regular shed cycle when viewed by a video camera

, or as indicated by instrumentation. E ngine operation is considered

stabilized when the measured engine parameters are not ch angin g, o r a regular and repeatable shed cycle has been demonstrated through the recording of measured engine parameters. You

may provide a justification to the FAA for a performance change while operating a steady

- st ate test point

. M omentary performance changes s uch as a thermodynamic engine response to shed

ice ingestion may be acceptable

. These momentary performance changes are due to an ice shed which can momentarily affect the thermodynamic cycle.

( 6 ) Turboprop a nd Turboshaft E ngines E quipped w ith a n I nlet S creen

. You may perform engine icing test s utilizing the aircraft inlet system or alternately with a bellmouth. Some

airframe installations utilize an inlet screen , which in some cases is included as part of the testing

for the part 33 compliance demonstration

. The inlet screen compliance demonstration for icing may be addressed separately from the § 33.68 engine ice compliance testing . If addressed separately, then you should note the requirement for an engine inlet screen in the engine

installation manual.(a) If the inlet screen is not addressed under § 33.68, then compliance could be

demonstrated under the installation requirements of §§ 23.1093, 25.1093

, for turboprop, and §§ 2 7.1093, or 29.1093 , for turboshaft installations

. Icing testing that utiliz es inlet screens should demonstrate acceptable operation

, including multiple shed cycles representative of long

term or unlimited operation in icing conditions. Protruding or unus ual inlet features may require specialized tests or analysis to assure they don’t adversely affect operations in icing conditions

.

( b ) At the conclusion of each steady

- state icing test point, the engine should be accelerated to takeoff powe r (throttle or power lever movement of one second or less) or shedding should be induced by other methods, depending on the critical shed methodology. The

throttle or power lever motion should be the most critical when considering the ice shed effects

on engine operability. In some cases, a quick deceleration before accelerating to takeoff power may be more critical to the ice shed effects on engine operation. A ssess this effect and account

for your assessment in your test proposal.

g

. Section 33. 68( d )

– Ground Operation

. Section 33.68( d ) provides two icing test points that represent freezing fog conditions ; a falling or blowing snow icing encounter during ground operation ; and a supercooled large droplet ( SLD ) test point that represents freezing rain or

drizzle . The SLD condition is similar to freezing fog in terms of ambient temperature range and

the LWC levels . However, the larger droplets can penetrate f a rther back on the surface of the engine spinner with the potential for ice shedding into the fan blades.

( 1 ) The ground

- fog icing, SLD

, and falling or blowing snow demonstrations of § 33.68(d)

should continue for at least 30 minutes

, or until acceptable operation is demonstrated. If you cannot achieve acceptable operation then you should demonstrate the periodic engine speed run

ups. These run

- ups will become mandatory in the engine operating manual and aircraft flight

manual for operating in these icing conditions

, since they were required to comply with the icing testing requirem ents. Turboshaft engines do not need to address SLD conditions.

(2) Falling and Blowing Snow

. Service experience has shown that compressor damage

occurs as a result of exposure to prolonged periods of falling snow ingestion during ground

opera tion. Based on our review of service events we have found that airports have co ntinued to operate with falling snow concentrations that result in 0.25 mile or less visibility (about 0.9 gm/m 3 of snow)

. While visibility in snow can be a poor indicator of p recipitation rates ( see reference in paragraph 5 .o. of this AC ) , the following calculation is based on equivalent rainfall

rate and gives a similar result as the 0.25 mile visibility criteria :

(a) Visibility and Snow Fall Rates

. The maximu m precipitation rate for moderate snow

is equivalent to 2.5 millimeters per hour ( mm/hr ) of rainfall over approximately 30 minutes

. Using a typical fall speed for snow of 0.8 meters per second ( m/s ) , this translates into a snow /ice

concentration of 0.9 g m /m 3 . From a Transport Canada data set of 338,000 minutes of snowfall data, the 95% and 99% values were 2 and 4 mm/hr respectively

, showing that a 2.5 mm/hr threshold provides an extreme value of snowfall rates. H oldover t ime tables for de

- icing fluids

ar e only good for the extreme values of moderate snow

, defined as 2.5 mm/hr ( or a visibility of

1⁄4 of a statute mile )

. This is equivalent to the engine certification levels described here.( b ) Engine core icing service events in snow conditio ns confirm t he critical snow

accretion temperature range is 2 5 0 to 32 0 Fahrenheit (

- 4 0 to 0 0 Celsius )

. The engine service

events have demonstrated that a snow environment is conducive to ice accretion behind the fan

in front compressor stages of the engin e , at low engine power

. The engine test for snow

conditions w as developed to represent the engine conditions where snow can form accretions aft of the fan on the core inlet and first stages. You should choose a temperature within the range provided above to achieve icing behind the fan.

h

. Ground Idle Demonstrations

. Section 33.68 requires operation at ground idle setting for at

least 30

- minutes under the icing conditions defined in § 33.68( d )

, followed by acceleration to

takeoff power or thrust

. Since a broad temperature range is provided within the regulation

, you should identify the most critical temperature, as determined by the CPA, and target that range. The CPA analysis can be used to demonstrate that colder conditions are less critical to the engine’s operation than the CPA condition that is demonstrated during ground icing tests. Turboprop and turboshaft engine inlet screen icing compliance demonstrations may be addressed separately from the § 33.68 engine ice compliance testing

. I f so

, you should note the requirement for an engine inlet screen in the engine installation manual. We recommend you give special attention to these inlet features :

(1) Unless data is provided indicating otherwise, the minimum ground idle engine spe ed

will be considered the most critical engine operating condition. U nless a more critical ground operating speed is identified, you should show that the engine continues to perform acceptably at its minimum ground idle speed attainable within all icing c onditions. T his test demonstration generally establishes the maximum allowable ground icing operation time for the engine. In

addition , this test usually establishes the maximum allowable time in the engine operation instructions between engine run

- ups t o shed ice. This time may be prescribed as a single duration or, if s upported by test conduct and results, the maximum time interval allowed between performing a repetitive, intentional, ice shedding engine run

- up procedure. You should

demonstrate s tabil ized i ce a ccretion, either naturally or with run

- ups, to achieve a ground icing taxi time which is appropriate for long taxi times in icing conditions.

(2) Engine ice shed characteristics ( for example, period, extent, location(s) ) during the

du ration of the § 33.68(d) test demonstration

, should be examined ( that is, by means of visual and/or suitable engine flowpath instrumentation) to determine the engine’s natural (unassisted by operator input) tendency to shed compressor system rotating and s tatic surface ice accumulations as well as the effectiveness of any intentional ( that is, prescribed, operator initiated) ice shedding engine run

- up p rocedure(s). Understanding these natural and, if applicable, assisted ice shedding

characteristics will a id in determining the necessary ground operational limitations in icing requirements ( for example, total and periodic exposure durations, prescribed procedures) for the engine.

(3) The resulting substantiated ground operating procedure in icing must be included in the engine’s operating instructions. If required, the applicable aircraft flight manual should specify the maximum total time allowed for ground icing operation and, if appropriate , the maximum time period between run

- ups consistent w ith the compliance test demonstration.Alternate run

- up procedures are also acceptable if you can show these will result in acceptable engine acceleration to takeoff power.

i

. I ce C rystal I cing ( part 33 t urbofan and t urboprop, only)

.

( 1 ) I ce crystal icing conditions that are hazardous to turbine engines are found in and

around cloud formation s contain in g deep convection

, in scales ranging from isolated and organized convective cells and complexes, to tropical storms ( for example, monsoons
, hurricanes, typhoons ) . Information on the mixed phase icing condition is located in the

reference shown in paragraph 5.i

. of this AC

. Total water content ( TWC ) up to 5 g m /m 3 can exist at these high altitude s , composed mostly of ice crystals

. The maxim um size of ice particles

within convective weather cloud s can be up to 10 mm or larger

, but mass is typically

concentrated at a much smaller median mass diameter of approximately 200 microns.

( 2 ) Appendix D of part 33 define s compliance requirem ents in terms of ice crystal and mixed phase icing environment s

. This mixed phase environment extends beyond the icing envelopes depicted in appendices C and O of part 25 or appendix C of part 29

. TWC

concentration levels in appendix D of part 33 are bas ed on predicted “adiabatic” condensation rates in convective storm updrafts. This has been shown to be a reasonable upper estimate based

on available flight test data , mostly from the 1950s

. However

, the accuracy of these flight test data is no longer ad equately traceable

, and the flight patterns adopted were not optimum for deriving TWC extent (cloud length) statistics. Therefore, the TWC extent probability of 99% put forth by appendix D is likely increasingly conservative with increasing cloud length

. Until a more comprehensive mixed phase and glaciated atmospheric envelope is defined from flight test data using modern cloud instrumentation to determine TWC and ice crystal size measurement

, appendix D of part 33 should be treated as the critical ice cr ystal icing condition when demonstrating compliance for § 33.68

. Likewise, a pplicants should evaluate e ngine inlet probes

for the conditions depicted in a ppendix D of part 33

.

j

. T est S etup C onsiderations . The LWC levels depicted in a ppendices C a nd O of part 25 and appendix C of part 29 are intended as supercooled droplet ambient icing conditions. Tests may

be conducted with a simulated cloud which is produced outside of the inlet and ingested into the engine. Under such a test environment, the LWC within the inlet ducting should replicate the engine operating conditions of an airplane in icing conditions at actual airspeeds. The inlet icing

concentration or dilution effect is dependent on droplet size, engine fan speed (for turbofan engines) , a nd simulated forward airspeed. For example, engine operation at idle descent power with simulated forward airspeed that is less than flight speed

, due to facility limitations

, might require a compensating increase in the test level of LWC concentration ab ove what is depicted in appendices C and O of part 25 . This increase would be greater for larger supercooled droplet

diameters. The engine size is also a variable that affects the difference in the LWC inlet concentration between flight conditions and th e engine test environment, with small engines potentially needing the greatest compensation. Rotorcraft engine applications may have different criteria and applicant s should propose a test that addresses those specific turboshaft

installation issues.k

. Direct Connect Inlet Engine Test Facilities

. Icing tests that provide a simulated icing cloud by direct connection of facility piping to the front flange of the engine, where no inlet air spillage is allowed, may cause test facility effects to alter the test parameters (for example, LWC and MVD). You must provide the F AA with data that demonstrat es the simulated test conditions are representative of an installed engine operating in the icing environment s as depicted in

a ppendices C and O of part 25 and appendix C of part 29, as applicable

. Suitable data may be

direct measurement data concerning LWC within the inlet

, or an acceptable validated analysis of water droplet trajectories for the test setup. In some cases, you may need to adjust the LWC to address any effects of the test setup (for example, non

- uniformity across the engine face).

l

. T esting i n M ixed P hase o r Glaciated Atmospheric Conditions ( part 33 t urboprop and

t urbofan e ngines only)

.

(1) Flight testing is a method of de monstrating engine operation in icing condition s, particularly mixed phase or glaciated atmospheric conditions

. Under these conditions, t wo important flight test considerations are : the measurement of ambient meteorological data

, and the ability to correl ate the measured engine performance to a more severe icing point

. To address

the correlation

, it is necessary that you have a fully instrumented engine with temperature sensors strategically located in the core flow passage . This instrumentation will col lect data during the icing environment encounter. Since the environment encountered may not be as

severe as the level s needed , scaling of the measured data to show satisfactory engine operation at

the more severe point should be attempted to the extent po ssible

. Any proposals for scaling should be supported by data from experiments

, and the scaling approach should be conservative.

(a) An accurate measurement of the ambient meteorological condition is essential. Past engine flight tests in this environment demonstrated that a combination of LWC probe, TWC probe , particle sizing and imaging probe, ice accretion detector

, and TAT probe are necessary to fully characterize the ambient environment.

(b) An o nboard real

- time meteor ological data display with GPS positioning capability

i s also helpful to find high ice crystal c oncentration area s . This allows the pilot to position the aircraft for the test. Since testing around a thunderstorm elevates the risk level of a flight test, an

on

- site rapid data reduction capability would give a timely indication if the test objective is met

, and minimizes the number of flights into this hazardous environment.

( 2 ) Simulation o f t he C ritical M ixed P hase a nd Glaciated Atmospheric C o ndition i n a G round Icing F acility

. As indicated in the reference shown in paragraph 5.i

. of this AC , facility simulation of ice crystal icing conditions is difficult and not currently done routinely. It is not known how well any facility simulation meth od actually replicates the natural environment. Consequently, no testing standard for this icing condition currently exists . Therefore, ground

simulation to demonstrat e compliance to the ice crystal icing certification requirement s will be

evaluated on a case

- by

- case basis until uniform industry standards are developed

.

m

. T est R esults a nd C ompliance I ssues . During all icing tests ( that is, supercooled droplet, or ice crystal icing), the engine should operate without accumulati ng ice that would ad versely affect engine operation . Accumulating ice can adversely affect engine operation (for example,flameout, surge, stall, run

- down, high vibrations, slow acceleration

, or lack of throttle or power lever response)

, or can cause a sustained loss of powe r or thrust. Additionally, the applicant should accurately monitor icing test point conditions . V ideo surveillance or other

instrumentation may provide the means to identify any source of ice damage, especially in those instances where test apparatus may also shed ice (for example, icing nozzles, special test

instrumentation , or icing tunnel walls ).

(1) Sustained L oss o f P ower o r T hrust a nd P ower L oss I nstabilities . The engine should

no t experience a sustained power loss while operating at app roved ratings in icing conditions.

(2) Temporary Power Loss

. Temporary power losses below engine power and thrust ratings selected under § 33.8 are acceptable if you show that the engine has a sufficient margin to

avert any power loss instabili ty, such as rollback, surge, stall, high vib ration s or flameout.

(3) Momentary Power Loss

. Momentary power loss caused by pumping or processing of ice debris through the fan module and compressor during the ice shed ingestion process is usually acceptable. Any accepted t emporary or m omentary p ower l oss or temporary high

vibrations must be documented in the engine installation manual.

n

. Mechanical Damage . The engine should not exhibit more than limited engine damage from any cause becau se of § 33.68 icing testing

. S ome limited damage can be acceptable if the

resultant power loss is minor

. Additionally you must fully account for cumulativ e damage from

repeat encounters

. Limited engine damage will be considered acceptable provided you s atisf y the following criteria:

( 1 ) Continued I n

- S ervice U se

. You should e valuate a ny resultant damage and demonstrate

that it does not affect the engine performance for continued in

- service use . Th is includes both continued safe operation, wit h no imminent failures expected, and no significant power loss

.

( 2 ) Sustained P ower L osses

. The engine should not experience any sustained power loss beyond 1.5% ( that is

, the nominal accepted level considered to be within measurement capabilit y )

.

( 3 ) Temporary or Momentary P ower L oss

. T emporary or m omentary p ower l oss should be reviewed by both the airframe and engine manufacturer to assess installed operational impact

. A ny resultant t emporary or m omentary p ower l oss
, or high vibra tions found acceptable by the FAA during the compliance demonstration , should be recorded in the installation manual by the engine manufacturer

.

( 4 ) Validation B asis . Analytical tools used to substantiate the criteria for determining acceptable damage should demonstrate a n acceptable validation basis . F or example, validation

could utilize engine tests or rig tests to substantiate the accuracy of r esults

. A n acceptable analytical tool validation basis includes test data which yield conservative results.

( 5 ) Engine Damage

. D amage to the engine or engine component s as a result of icing compliance testing should not exceed the maintenance manual limits

. C umulative damage forrepeat ed encounters should be considered part of this assessm ent

. A ny damage findings should be brought to the attention of the FAA for approval.

(6) High Vibrations

. You must e nsure that h igh vibrations are assessed for the specific aircraft installation. The acceptability of high vibration s is evalua ted by the airframer. Bring high vibrations to the attention of t he ACO or ECO approving the engine ; it will contact the ACO approving the engine installation before the final approval

, and inform them of th is

condition

. Any high vibrations should be inc luded in the installation manual by the engine

manufacturer

.

( 7 ) Communication o f R esults . The i nstallation and o perating m anuals required by § 33.5 should provide information describing all engine condition s observed during engine certificatio n icing tests. Prior to the se test s
, t he engine manufacturer should provide the FAA with a process

to evaluate the acceptability of any potential damage that could occur during icing tests.

(a) N atural icing flight tests are conducted to d emonstrate compliance with §§ 23.1093

, 25.1093 , 27.1093, or 29.1093 ; they require a n FAA preapproved process for evaluating damage resulting from icing tests. Also, if periodic engine power run

- ups are used to minimize icing damage during the ground icing operation demonstration defined in § 33.68( d ), then the applicant must document the run

- up procedure

. Complete d ocumentation must contain a description of the run

- up requirements and the required run

- up intervals

, and it must be

contained in the operatio n manual as mandatory with in icing conditions

. Prior to demonstration,

t he engine manufacturer should coordinate the proposed run

- up shedding procedures with the airplane manufacturer , so as to assure that procedures are practical and appropriate for the intended operations.

(b) Any acceptable power loss anomalies from for example, ice accretion, ice shed,

water runback and refreeze and shed , and the effects on performance and operation

, should be

documented in the installation m anual. Bot h the engine certifying ACO and the installing ACO

should carefully evaluate any high vibrations induced from ice accretions during ice testing

. T hese vibrations should be documented by the engine manufacturer in the engine’s i nstallation

m anual.

o

. Engine S ystems

. Applicants may use an automatic engine control system to initiate an ice protection system to meet the requirements defined in § 33.68

, provided that system operation

does not result in potential crew action. Examples of engine characte ristics that may be

noticeable to the flight crew are exhaust gas temp fluctuations , large speed fluctuations , or audible surging. Further, a ny engine control system required for engine ice protection

certification under part 33 should not react adverse ly with other aircraft systems, aircraft handling qualities and performance, or human factor considerations.

(1) Any unacceptable adverse interactions may result in the engine being deemed un

installable under part s 23 or 25 airplanes or part s 27 or 29 rotorcraft

. C ritical issues which you

must address include : c rew interface, uncommanded thrust or power changes, thrust or power setting, asymmetric engine behavior, pilot workload and appropriate flight deck indication and procedures, the effect on air craft handling, pilot ability, and human factors. Additionally, anyengine system required to show compliance under § 33.68 should meet the following requirements:

( a ) System R eliability . Demonstrate the capability of the system for r eliably sensing the conditions, which enables the engine system function, throughout the approved operating envelope

.

( b ) Dispatch . The engine system function should be available for all dispatch configurations. The system should be confi gured in its most critical dispatch state for certification icing tests

.

( c ) Electronic F aults

. Applicants should demonstrate that engine system function s will

not be lost due to any probable electronic fault ( s )

.

( d ) Other E nvironmental T esting

. Engine system function s should not be affected when the system and any associated electronic systems are exposed to the required operating environments, including high intensity radiated fields ( HIRF) and lightning

.

( e ) Power R equirements . For those systems powered solely with a dedicated engine alternator (either directly or using another engine system such as a FADEC), applicants should

demonstrate the system sensing and performance function at minimum certified r otor speeds. Minimum certified engine speed is the minimum idle rotor speed achievable anywhere in the icing envelope

.

p

. Auto

- R ecovery S ystems . Auto

- recovery systems should not be needed under § 33.68 testing since these icing conditions are cons idered to be within the engine’s certified operational envelope and a probable encounter . The intent of § 33.68 is to certify engines to perform and operate reliably in the icing conditions depicted in a ppendices C and O of part 25, appendix C of part 29

, and a ppendix D of part 33 , as applicable

.

(1) Auto

- recovery systems are considered back

- up devices that are only needed following rare ice ingestion events result ing from icing conditions outside appendices C and O of part 25

, appendix C of par t 29, and appendix D of part 33

. If auto

- recovery sy s tems activate during the icing test

, then the applicant should alert the installing ACO of the activation

. Auto

- recovery systems are not the primary protection for continued safe engine operation durin g normal ice sheds or accretion while operating in icing conditions depicted in appendices C and O of part 25

, appendix C of part 29, and appendix D of part 33

.

(2) Therefore, applicants may perform the § 33.68 compliance testing with auto

- recov ery systems enabled, but the systems should not activate during the § 33.68 test sequence. Additionally, continuous ignition should not be selected during the § 33.68 compliance testing.

To assure non

- activation of an enabled auto

- recovery system, the ap plicant should have displayed an instrumented signal that monitors the auto

- recovery system activation

. If activation monitoring cannot be accomplished, then disabling the auto

- recovery system will be necessary.q

. Operation I nstructions . Any oper ating procedure (for example, ground run

- up procedures) required to ensure continued operational compliance with ground icing conditions or falling and blowing snow evaluated under § 33.68(d), and the installation requirements of §§ 23.1093(b)(1)(ii), 23.1 093(b)(2)

, 25.1093 , 27.1093(b)(1)(ii) or 29.1093(b)(1)(ii) , should be

communicated to the installer in the o perating i nstructions as a requirement

. The requirement for each applicable operating procedure should further be included in the limitation sectio n of the airplane flight m anual. It may be necessary to coordinate with the installer on these procedures

to ensure that they can be effectively implemented in

- service. The installer may translate the operating instructions into a procedure tailored at t he airplane level if it results in an operation

that is equivalent to the demonstrated operating procedure. An equivalent operation should be conservatively representative of the demonstrated operating procedure.

r

. Special Considerations f or Mixed Phase a nd Glaciated Atmospheric Conditions ( part 33

t urbofan and t urboprop only)

. P art 33 requirements for the type certification approval of engines operating in an ice crystal environment are located in § 33.68

, and appendix D of part 33 . These requir ements have been developed in response to service events. The root cause of these events can be traced to ice buildup within the core flow path of the affected engines. Engine inlet probe blockage and the resulting signal corruptio n should also be analyz ed

. Adverse effects created by this type of icing on the engine include un

- commanded rollback of power or flameout

, as well as compressor stall and core hardware damage.

(1) Ice crystals have only been recognized as a threat to turbine engines in recent years. In

response, the FAA has worked with industry to develop standards t hat address this threat. During this process, we found that although capabilities are developing, the present ice crystal tools and test techniques have not been fully d eveloped and validated sufficiently

. Therefore

, we developed a phased

- in approach to address th e ice crystal threat during an engine type certification program. The following paragraphs describe this phased

- in approach.

(2) N ew engines must ad dress the known in

- service experience in ice crystal environments ( for example, core damage and engine flameout events ). Until ice crystal tools and test techniques have been developed and validated, the engine manufacturer should use a comparative analys is to specific field events. This analysis approach should show that new engine cycle or design features, or both, will result in acceptable engine operation.

(3) Acceptable operation include s the absence of rollback, rundown, stall, flameout, and unacceptable compressor blade damage as described earlier in paragraph 8.a of this AC

. Additionally, we recommend the engine manufacturer incorporat e developed and validated technology as it becomes available, and work toward a full engine test substa ntiation of operation in ice crystal environments

.

(4) Long

- term, we anticipate an acceptable demonstration will eventually include a CPA of an ice crystal environment as depicted in appendix D of p art 33. All engine power levels

, including in

- flight idle operation

, will be evaluated in these conditions. The critical conditions

will be demonstrated to the FAA through a combination of testing and validated analysis using the latest tools and technology when proposing the compliance methodology. Computationaltools to be used in this analysis would be calibrated by either rig calibration test data or engine

test measurements.

s

. Comparative Analysis Guidanc e for Ice Crystal Icing ( part 33 t urbofan and t urboprop only)

. As stated earlier, un til ice crystal tools and test techniques have been developed and validated, the engine manufacturer should use a comparative analysis to specific field events. This analysis

should show the new engine cycle or design feature , or both, would result in an acceptable engine operation when subjected to the ice crystal environment depicted in appendix D of p art 33. This comparative analysis should consider both suspected susceptible design features

, as well as mitigating design features.

(1) Suscep tible D esign F eatures

. The se features could include:

(a) Stagnation points which could provide an increased accretion potential.

(b) Exposed core entrance (as opposed to hidden core).

(c) High turning rates in the inlet, in the booster and core flowpath (particularly compound turning elements)

.

(d) Protrusions into the core flowpath ( for example

, bleed door edges and measurement probes)

.

(e) Unheated surfaces on booster and fro nt core stages.

(f) Narrow vane

- to

- vane circumferential stator spacing leading to a small stator passage hydraulic diameter

.

(2) Mitigation Features

. These susceptible design features can be significantly mitigated by one or mor e of the following design features. Mitigating design features could include:

(a) Heated surfaces in the fan, booster

, and forward core compressor stages.

(b) Elevated rotor speeds.

(c) Hidden core entrance.

(d) Low frontal cross

- sectional area on flowpath probes.

(e) Inter

- compressor bleed scheduling to remove both the ice crystal media and any up

stream shedding, from the flowpath

.

(f) Circumferential spacing of stators set to enhance tolerance to ice blockage (generally denoted by the hydraulic diameter of the stator passage)

.(g) Increasing c ore compressor airfoil tolerance to soft

- body damage

. Soft

- body damage is typically airfoil bending, w hereas hard

- body damage results in tearing and fracture.

(3) S imilarity t o E ngines P roven S afe t o O perate i n M ixed P hase o r Glaciated Atmospheric C onditions (part 33)

. Although ice crystal icing conditions are hazardous to turbine engine operat ion, severe incident s i nvolving this type of meteorological condition are not common. Many currently certified engine designs have been proven by their field service experience to be safe to operate in these conditions. New engine design s that are simila r to those

of proven engine design s are allowed to show compliance by comparative analysis.

(a) Several steps are required to demonstrate compliance by similarity analysis. First,

you should identify the baseline (certified) engine

, and su pply evidence that this engine is safe to operate in mixed phase or glaciated icing condition s . This evidence can be field service experience and/ or a certification report.

(b) Second

, identify the icing

- pertinent engine features that may influence mixed phase or glaciated ice accretion within the target (certifying) engine ; then compare them to the baseline engine . This comparison should establish that the new engine model is less than or equally susceptible to icing as the baseline model

.

(4) Comparative Analysis Versu s C PA

. A comparative severity analysis should be performed to show the operation al envelope of the target engine does not make it more susceptible to ice crystal icing than the baseline engine. A s the icing comm unity’s knowledge improves through research , in the long term applicants should perform a CPA to fully assess ice crystal icing. This severity analysis should consider both environmental and engine operational effects on accumulation, accretion locations, as well as ice shedding.

(5) Innovative o r Novel Designs

. If the new engine cycle or design feature , or both

, contain s innovative design ideas such that a comparative analysis with current engines and specific engine events is not possible, th en you should demonstrate that the innovative feature (s) will not be susceptible to any adverse effects of operating in an ice crystal environment. Such a demonstration entails a two part process. In p art one you should document the physical basis regard ing why this design would result in acceptable operation in ice crystals. In p art two you should generate physical evidence to substantiate the claims of p art one

. For purposes of substantiation, w ater particles may be used instead of ice crystals

. Howe ver, data collected with water particles should be corrected to account for ice crystal characteristics , including liquid to solid phase thermodynamic effects

. The use of water to simulate crystals will be re

- evaluated by the FAA as ice crystal test techn iques improve through research and development.Table A

. Example S teps T hat M ay B e U sed

Part 1 Part 2

Discuss consequences of ice accretion. Identify test article (engine or component). Identify design features that promote tolerance

to critical ice accretion. Establish test conditions promoting ice accretion (with allowance for use of liquid water if testing with ice crystals is not feasible). Discuss consequences of ice crystals on promotion of ice accretion. Evaluate tolerance of new design featur es over a range of test points.

t

. Compliance Considerations ( p art 33)

. Engine icing events in an ice crystal environment appear to be associated with accretion of ice deeper in the compressor

, as compared to supercooled droplet environments where accretion occurs farther forward

.

u

. Ice Crystal Accretion Explanation

. The accretion process is believed to be a result of ice crystals passing through the engine inlet and into the front compressor stages . This process

continues until the local conditions within the engine are conducive to forming a liquid layer on

vane and blade surfaces

. Once the local conditions are conducive to the form ation of a liquid layer, the vane and blade surface condition allows impinging crystals to melt and stick to the liquid layer and accrete on the surface. This is a combination of conductive and convective heat

transfer from ice melting, evaporation

, and surface contact.

( 1 ) The combination of both the liquid water and solid ice crystal draw signifi cant heat away from the impinged surface. An impingement surface that was initially as high as 120 0

Fahrenheit ( 49 0 Celsius) can be cooled down to 32 0 Fahrenheit (0 0 Celsius) by this dual phase ice

and water media on the surface.

v. Ice Crystal Ici ng Assessment ( p art 33 , t urbofan and t urboprop only)

. You should assess engine operation across the whole icing envelope . C ruise, hold

, and descent power settings

should be evaluated in mixed phase and glaciated atmospheric conditions as depicted in

a ppe ndix D of part 33

. Note that n o established standardized CPA for i ce crystal icing currently

exists

. However, t o aid you
, we are recommend ing some possible CPA point selections

described in Table B below

. We also recommend applicants use the associated general criteria

when selecting proposed conditions to be evaluated.Table B

. General Criteria f or Selecting Conditions to be Evaluated

Service H istory This suggests that a selection of conditions at high ambient

temperatures depicted in a ppendix D of part 33 icing envelope

,

are consistent with higher TWC levels. However, l ower temperatures might push the accretion aft, where there may be

a more critical ice accretion site with lower TWC. Therefore, both high and low temperatures within the envelope s hould be

evaluated.

Power Levels To promote ice crystal melting, p ower levels with internal total air temperature within the core flowpath should be

between 32 0 Fahrenheit (0 0 Celsius) and approximately 120 0

Fahrenheit (49 0 Celsius)

. The p ower level a djusts the accretion site forward or aft within the engine.

High Altitude High altitude (low air density) allow s greater ice accretion

mass before shedding from static surfaces.

(1) The operating conditions for this evaluation should be cho sen from a ppendix D of part
33

, covering the flight phases to includ e climb, cruise, idle

- descent, and holding

. The water

contents depicted in a ppendix D represent the level for a standard exposure distance of 17.4

nautical miles

. To adjust the water con tent level as required, so that it reflect s the expected icing exposure period

, us e the distance scale factor depicted in a ppendix D. Service experience suggests that straight line exposure distances of 20 to 80 nautical miles may be encountered , as well as holding in cloud

. As our understanding of water content in relation to distance flown becomes more precise, this encounter distance may be adjusted. Currently we hold that applicants who test to this greater distance will assure their engines are robu st and meet the minimum standards of the regulations.

(2) As noted earlier, ice crystals promote icing at engine sites rearward and at higher local air temperatures than would exist with only supercooled liquid. These rearward internal flowpath accretion sites will also have higher air loads, which can limit ice accretion due to shedding. Therefore , the CPA should consider not only ice accretion, but also the likelihood of ice shedding.

w

. A ppendix O of Part 25

, Water Impingement

.

(1) The water impingement region may become greater when encountering the large water droplets depicted in a ppendix O of part 25. This occurs because larger droplets have greater inertia and follow a more ballistic trajectory less prone to influence by local airflow streamlines that bend to deflect air around an object . A way to illustrate this droplet inertia effect is through the use of the modified inertia parameter variable , K
o , as shown in Figure 1 of this AC (see

reference g, in section 5 of this AC

. This parameter is a relationship based on the momentum of the droplet, air viscosity

, and size scale of the impingement region (with a correction forReynolds n umber). It is used to correlate the streamline effect o f the airflow in deflecting a droplet and prevent ing it from impinging on a n object. The equation for K
o is provided below for a droplet with diameter (d) with an upstream velocity (V

o ) traveling in an airflow with a viscosity (μ).

(2) The impingement efficiency E

m is simpl y a decimal fraction of the water droplets that are aligned with an object far upstream that actually do impinge on the surface. For very large droplets, impingement efficiency approaches unity and is progressively less for smaller droplets.

Published co rrelations exist for various object shapes of interest for ice impingement. Such a

correlation between K

o and impingement efficiency E

m , is shown below in Figure 1 for several object shapes.Figure 1

. Droplet I mpingement E fficiency f or D ifferent S hape s

Figure 1 a. Definition of Inertia Parameter, Equation 3.10 from Reference 5.g of this AC

Inert i a Parameter K = 2(d/2) 2 

# ρ H2 0 V 0 (Di mensionless) EQU 3.10 9 L C 

### μ Where Characteristic length L C of ob ject is:

Airfoils L C = Chord Cylinders L C = Radius Spheres L C = Radius Cones L C = Base Radius( 3 ) Lower i mpingement efficienc ies (less than a value of one) are indicative of droplet

deflection from the object

. When dro plets are deflected, they don’t accrete on the object. This deflection is caused by streamlines moving around the object. For most shapes

, a modified inertial parameter of 10 or greater implies a ballistic path for the droplets. A ballistic path means

t he droplet does not follow stream lines and can impinge the vane f a rther back than a non

ballistic droplet. This results in ice accretion f a rther back on the vane.

( 4 ) Impingement of a supercooled water droplet on a surface is dependent on the droplet size, and the size of the geometric feature. The smaller the drop size and the bigger the feature, the more likely the drop will follow the aerodynamic streamlines around the body and fail to impinge. Larger drops, with more inertia are less like ly to follow streamlines, and therefore they

will impinge on larger features. In the engine, such features include the inlet and spinner. The applicant must account for the potential impingement of larger drops on engine geometric

features of this scale. Figure 2 of this AC illustrates the consequences of encountering the larger water droplets depicted in a ppendix O of part 25. Figure 2 also shows what happens with target objects of typical size ranges for larger transport engines. The shaded area in F igure 2 cover s

a ppendix C of part 25 and appendix C of part 29, as applicable

. For a given modified inertia parameter value (K

o ) of 10, t he droplet size range is seen to behave ballistic al ly for small “target” object s
, such as engine airfoils, airfoil cas cades

, and probes. Consequently, large droplets as depicted in a ppendix O of part 25 will not significantly affect surface water impingement and resulting ice accretions on small airfoils within engines

. For larger targets, however, applicants should ant icipate greater impingement on surfaces , f or example : the nacelle and engine spinner . Other surfaces unprotected by surface heating systems are also at risk for water impingement by the sized droplets depicted in a ppendix C of part s 25 and 29, as applicab le

.

Figure 2 . Ballistic Boundary as a Function of Target and Droplet Size

10 100 1000 10000 Droplet Diameter

- microns Airfoils Probes Nacelle Lip Spinner Inlet

Diameter

Cascades

.01

.1 1 10“Target”Dimension-ft

Ballistic (Minimal Deflection by Streamlines) Locus of Modified Inertia Parameter = 10

Droplet
Size Range

of Appdx C Ko > 10 (example) Ko = 10

Ko < 10 (example)

10 100 1000 10000 Droplet Diameter

- microns Airfoils Probes Nacelle Lip Spinner Inlet

Diameter

Cascades

.01

.1 1 10“Target”Dimension-ft

Ballistic (Minimal Deflection by Streamlines) Locus of Modified Inertia Parameter = 10

Droplet
Size Range

of Appdx C Ko > 10 (example) Ko = 10

Ko < 10 (example)10

. S ection 33.77, Ice Slab Ingestion

.

a. Why the Ice Slab Ingestion Test is Conducted

. The intent of the ice slab ingestion test is

to demonstrate tolerance to ice ingestion due to shedding from inlet ( nacelle ) surfaces in the front of the engine

. This test also establish es limits for ice released from other aircraft or rotorcraft surfaces during part s 23

, 25 , 27, and 29 certification s

. The dimensions of the test s lab are related to engine size (defined by inlet highlight area), based on service experience. Manufacturers of turboshaft engines with complex inlet designs should propose an equivalent inlet highlight area once they have defin ed the ice slab size. E ngi ne manufacturer s should also include in their compliance plan, an analysis of the potential installation effects of the engine induction system.

(1) T he engine manufacturer and the installer should close ly coordinat e the ice slab sizing

and dens ity . This coordination ensure s that ice accumulation sites on the airframe which can potentially be ingested by the engine are addressed under § 33.77. Airframe locations that

should be considered include for example, the inboard section of the wing for an aft fuselage mounted engine , the radome

, and antenna. Ice slab ingestion is demonstrated under § 33.77 , but

must be addressed and shown as acceptable for the installation under §§ 23.901(d)(2), 23.1093, 25.1091(e), 25.1093 , 27.1093, or 29.1093 (see par agraph 11 of this AC).

( 2 ) The airframer responsible for the induction system should assess the inlet cowl ice accumulations in accordance with §§ 23.901(d)(2), 23.1093, 25.1093 , 27.1093, or 29.1093

. They should also provide pertinent test vari ables to the engine manufacturer for incorporation

into a § 33.77 test demonstration. If an application or product inlet has not been selected at the time of engine certification , the engine manufacturer should include in the engine installation

manual al l pertinent inlet icing assumptions

, test data

, and results. See paragraph 11 of this AC for installer assessment guidance.

b. Compliance Considerations

. Compliance may be demonstrated by the standard engine ice

slab test or by using a validated a nalysis procedure that uses equivalent soft body testing . The

test demonstration should use ice slab trajectories aimed at critical engine locations . Applicants should pick locations based on the ice accretion and shed characteristics of the induction sy stem likely to be installed on the engine. The most critical impact location should be tested.

(1) Slab Dimension Considerations

. If the engine manufacturer type certificate applicant

l ack s specific knowledge of the icing characteristics of an aircraft inlet system

, it may select test conditions typical of an installation already in

- service , in combination with conservative assumptions on the future installation . The ice slab size, thickness, and density defined in the § 33.77 compliance demon stration should be evaluated against parts 23 , 25, 27

, and 2 9 engine installation and inlet system icing requirements

. Applicants should use a m inimum ice slab density equivalent to a 0.9 specific gravity

, unless a different value is considered more

appro priate ( see AC 20

- 73 A , AC 23.1419

- 2D , AC 25.1419

- 1

, AC 27

- 1B , or AC 29

- 2C (or latest revisions) for more guidance on ice shedding )

.

(2) Engine Test Considerations

. Applicants should determine if the ice slab size, thickness, and density are app ropriate for the specific engine installation . If appropriate, the testresults of § 3 3. 77 might be used by the airframe manufacturer to comply with the natural icing flight test requirements of §§ 23.1093

, 25.1093 , 27.1093, and 29.1093 . Execution of the ice slab ingestion test typically involves targeting the slab to enter the air stream ahead of the fan

, intact

, at the outer diameter of the inlet duct. This is intended to mimic the ice release from the inlet and results in impact on the outer diameter of the fan.

c. Validated Analysis w ith Equivalent Soft Body Tolerance Testing

. Compliance may also be shown by a validated analysis procedure that uses appropriate soft body damage testing. If the applicant elects this alternate compliance approac h , the engine will be certified for the minimum standard ice slab consistent with the engine inlet area as defined in § 33.77. We recognize that alternate soft body damage testing may
, in some circumstances

, include objects that are larger than the standa rd ice slab based on inlet area. However, c ertification of an ice

slab larger than the standard size by this validated analysis method is not currently allowed.

d. Elements o f a Validated Analysis

. This analytical model may be used alone or in

con junction with the results of a certification medium bird or other soft body ingestion test. A validated analysis must contain sufficient elements to show compliance

. T hese elements may

include:

• Full fan (fan engines) or first stage compressor ( non

- fan e ngines) blade model ing using

the latest techniques such as finite element analysis ;

• Blade material properties for yield o r failure , or both, as appropriate ;

• Dynamic and time variant capability ;

• Thrust or power variance prediction if required to account for blade or other damage ; and

• Appropriate engine or component testing , or both, with impact at the outer 1/3 of the first stage blade span location. The fan is the first stage blade row for turbofan engines.

(1) The analysis of the ice slab impa ct on the fan must properly account for critical controlling parameters:

• Relative kinetic energy normal to the leading edge chord ;

• Incidence angle

– relative slab speed and blade speed ;

• Slab dimensions ; and

• Slab orientation

.

(2) Any predicted power loss or blade damage (distortion, cracking, tearing) must be assessed against the criteria of this AC

. F igures 3 and 4 of this AC help describe the contribution of these parameters when establishing the threat to the blade.Figure 3

. Normal C ompo nent o f Kinetic Energy o f I ce S lab

Figure 4

. Ice S lab O rientation E ffects

Dependent on orientation at fan face the “bite” can be increased substantially

Load shared by 6

- 7 blades Load shared by 3 blades rotation

slab 

### Relative KE component

normal to leading edge is

critical(3) The relative kinetic energy of the ice slab should be determined from an assessment of flight conditions that control engine rotor speed versus ice slab velocity

. Engine test results from previous ice slab testing may be used to support the predicted ice slab velocity. The applicant’s analysis should assume the most critical orientation

, unless they can show that an alt ernate ice

slab orientation is more conservative for ice slab testing purposes

.

e. Ice Slab Break Up

. Typi cally, the ice slab breaks into smaller pieces during ingestion

. The applicant’s analysis should use the largest slab size consistent with a conservative assessment of slab “break up” that can occur within the air stream ahead of the fan. Data derived

from a number of tests are included in F igure 5 below

. T he largest ice piece is typically 1/3 to 1/2 the original size. For compliance by anal ysis , the applicant should assume 1/2 of the original slab length unless evidence suggests that this is not conservative relative to ice slab testing.

Figure 5

. Ice S lab B reak U p E xperience

1⁄2 slab dimension

– minimum allowedf

. Test Results . Section 33.77(c) requires that unde r the conditions in § 33.77(e), ice ingestion must not cause a sustained power or thrust loss, or require the engine to be shutdown. The following paragraphs cover these points in greater detail :

(1) Sustained P ower L oss

. A pplicants should eva luate the impact of a ny first stage blade bending or damage on potential sustained engine power loss. Power loss associated with blade damage from the slab should be less than 1.5%. As soft body blade damage is common from medium bird ingestion , a pplican ts may use the medium bird ingestion test results to show compliance with this requirement. If the medium bird ingestion test results is less than 1.5% permanent power loss

, and no cracks, tears

, or blade piece breakout occurs due to a bird

introduced at the outer 1/3 of the fan blade span

, then the § 33.77 requirement is met.

(a) If power loss exceeds 1.5% when utilizing the bird test , the manufacturer must provide a validated analysis that shows consistency with the bird test results . Th e manufacturer

must also demonstrate the standard ice slab would produce less than the 1.5% power loss.

(b) Applicants must also demonstrate by test that any cracks, tears

, or blade piece

breakout will not result in “ unacceptable sustained power or thrust loss ” within 100 flight cycles

. This is considered sufficient to allow continued engine use until the next scheduled “A” check for fixed wing applications, or similar inspection periods for rotorcraft applications

. Note that

any damage re sult ing from this test must be documented in the engine installation manual.

(2) Engine O perability

. Engine damage should not cause surge, flameout, or prevent transient operation.

(3) In

- S ervice C apability

. Engine damage should not result in a failure or a performance loss that would prevent continued safe operation during a conservative flight operations scenario

. F or example, an engine should be able to operate within the time period for an “A” check or greater

, if appropriate te sting validates a continued period of in

- service capability . The period of

in

- service capability to be demonstrated may vary with installation if the damage is not readily evident to the crew or visible on preflight inspection (for example, tail mounted e ngine location ).

(4) Other A nomalies

. Engine d amage should not result in any other anomaly (for example, vibration) that may cause the engine to exceed operating or structural limitations.

(5) Auto

- R ecovery S ystems . If during the ic e slab ingestion testing under § 33.77, an engine incur s a momentary flameout and auto

- relight, then the acceptance of that test is predicated on inclu ding the auto

- relight system as a required part of the engine ’ s type design

. Furthermore

, additional cri teria would be required where the ignition system is fully operable

before each dispatch. The reason for the additional dispatch criteria is to ensure the ignition system’s critical relight function is reliably available during the next flight. We allow use of the

auto

- recovery system during § 33.77 certification testing to account for ice accretion and shedding that result s from an inadvertent delay in actuating the anti

- icing system . We consider

this delay to be an abnormal operational result where ope rability effects

, like momentary flameout and relight

, may be accepted.g

. Communication o f Test Results

.

(1) The installation and operation instructions required by § 33.5 should provide information on the size, thickness, and density of the ice slab ingested ; any anomalous behavior such as high vibrations

, and any affect on the engine ’ s ability to operate at the commanded power setting or rating. By including t he information in the installation manual , you provide

the se test results to the installer.

(2) The icing certification report addressing § 33.77 compliance should include information regarding ice slab orientation and trajectories, slab breakup, impact locations, description s of any resultant damage, as well as all othe r pertinent data defining the engine's capability or response to the ice ingestion event. Additionally, if the auto

- recovery system is required to comply with § 33.77, then the functional state of the recovery system (for example, one igniter inoperative) becomes a limitation that needs to be communicated to the installer via the engine installation manual to ensure compliance with the delayed activation requirements of §§ 23.1093

, 25.1093 , 27.1093, or 29.1093

.

1 1

. I nduction System Icing Protection (§§ 23.1093, 25.1093, 27.1093, and 29.1093)

. A pplicants usually use the results of § 33.77 for compliance with §§ 23.1093, 25.1093, 27.1093,

and 29.1093

, instead of a test d emonstration of the 2

- minute delayed activation of inlet anti

- ice (1

- minute for turbo shaft engines) . The § 33.77 results have also been used to show compliance

with §§ 23.1093, 25.1093, 27.1093, and 29.1093 for other airframe ice sources. T he engine

manufacturer and the airframer should coordinate closely to ensure the § 33.77 test cover s all potential ice sources. Similarly the icing compliance demonstration of § 33.68 should be

coordinat ed with the installer to verify the engine icing compliance demonstration is applicable

to the installation icing compliance requirements

. The icing c onditions depicted in appendix O of part 25 are only applicable to turbine engine induction systems on part 25 airplanes with a maximum takeoff weight less than 60,000 pounds. Turbine engine induction systems on part 25 airplanes with a maximum takeoff we ight equal to or greater than 60,000 pounds need only comply with the icing conditions depicted in appendix C of part 2 5 , appendix D of part 33 and in falling and blowing snow within the limitations established for the airplane for such operation.

a. Acceptable Means of Compliance (§§ 23 .1093(b)(1)

, 25.1093(b)(1) , 27.1093 (b)

, and

29.1093 (b) )

. As a general rule, engine induction systems should operate continuously in icing

conditions without regard to time , as in a hold condition

. For fixed wing air craft, a 45

- minute

test duration in a standard continuous maximum cloud is an acceptable demonstration of

continuous in

- flight operation, so long as the test demonstration result isn’t indicating an

imminent test criteria failure. An exception would be fo r low engine power conditions where a sustained level flight is not possible. Even then, a conservative approach must be used when a

series of multiple horizontal and vertical cloud extent factors are assumed. T he horizontal cloud

extent factor i s not in tended to be used to limit the severity of exposure to those icing conditions where it is reasonable to assume the aircraft will be required to operate. (1) A pplicant s must adequately analyze the engine inlet anti

- ice system performance and

addre ss any potential ingestion hazard s to the engine from predicted ice buildup on the engine

inlet

. This includ es runback and lip ice. If an applicant can show that both the inlet anti

- ice system performance and the first stage compressor blade ( for example , fan blade for turbofanengines) capability are equivalent to previous certification experience, then certification may be

shown through similarity to previous designs.

(2) You must show by analysis, and verify by test, that the engine inlet an ti

- ice system provides adequate protection under all flight operations. Your analysis should include the

conditions shown in Table C of this AC

. N ote that additional critical points may be needed depending on the specifics of the airplane and engine desi gn. Bleed crossover points may also need to be analyzed, if applicable.

b. Inlet Design Point Selection

. If the inlet is evaporative under the critical points in continuous maximum icing conditions, and if it is running wet under intermittent maxi mum icing

conditions ( refer to a ppendix C of part s 25 and 29 ) , then the design is satisfactory. Turbofan and turboprop e ngine inlet anti

- ice systems have historically demonstrated good service experience

with inlet anti

- ice systems that run wet

, when runb ack ice ha s been evaluated at the conditions

described in Table C of this AC

. For t urbofan and turboprop i nlets that are certified for unlimited operation in icing

, runback ice formation is allowed during hold, descent

, and

diversions under the conditions shown in T able C of this AC

. Turbofan and turboprop installers

should determine the critical accretion conditions for their inlets in each of the scenarios in Table

C below

. T hey should then compare each result to the amount of ice the engine was satisf actorily demonstrated to ingest during engine certification ( refer to § 33.77).

Table C

. Inlet Lip a nd Runback Ice ( part 33 t urbofan and t urboprop engines only )

Flight

Condition Descent Hold Straight

- L ine F light

Icing Design

Condition 6500 foot

descen t through

a continuous

maximum

cloud , as depicted in

a ppendix C of part 25

, with a cloud extent
factor of 1, followed by an

intermittent

maximum cloud

exposure, as depicted in

a ppendix C of part 25

. 45

- minute hold in a continuous

maximum cloud

, as depicted in appendix C of part 25, with a cloud

extent factor of 1

. 45

- minute exposure in a continuous maximum cloud, as depicted in appendix C of part 25,

with a cloud extent factor as depicted in appendix C of part 25

, followed by an intermittent

maximum cloud exposure, as depicted in appendix C of part 25.c. Addressing Specific Operational C onditions

.

( 1 ) Descent

. If the engine inlet anti

- ice system is not fully evaporative during descent through a continuous maximum cloud as depicted in a ppendix C of part 25

, then the total amount
of ice accretion (including inlet lip and runback ice ) must be calculated for both the continuous maximum and intermittent maximum icing conditions depicted in a ppendix C , and for the operating conditions shown i n Table C above. Airspeed and scoop factor (if applicable for the inlet design) should be part of this assessment. Ingestion of these calculated quantities of ice buildup for a given condition must not result in more ice ingestion damage ( that is, based on slab size or kinetic energy ) than the amount of ice the engine was satisfactorily demonstrated to ingest during the engine certification ( refer to § 33.77).

(2) Holding

. T he engine inlet anti

- icing system must be capable of safe operation in the

cloud conditions depicted in a ppendix C of part s 25 or 29 for extended operations. The extended holding condition is for 45 minutes ( except, 30 minutes is used for p art s 27 and 29) at the critical

accretion ice conditions for the continuous maximum i cing cloud

, using a horizontal cloud extent
factor of one.

(a) If the inlet is evaporative under continuous maximum icing conditions, and it is running wet under intermittent maximum icing conditions, then the design is satisfactory

. T his is because the descent condition provides less power to the anti

- ice system and will always

become the critical accretion condition in intermittent maximum icing conditions

.

(b) If the inlet is running wet in a maximum continuous atmospheri c condition, then the applicant should calculate the amount of ice that would accumulate during the holding

conditions as described in Table C above. The total amount of calculated ice should include runback and lip ice for the condition being analyzed. Ingestion of this ice must not result in more damage than the amount of ice the engine was satisfactorily demonstrated to ingest during engine certification test s ( refer to § 33.77). This damage assessment should be based on the kinetic energy at the lead ing edge of the first stage compressor blade ( for example, the fan for turbofan engines) , and should include consider ations of ice sl ab size, velocity, density

, and compressor (or fan) rotor speed.

(3) Straight

- Line Flight

. For straight

- line fl ight ( for example, cruise, and diversion to alternate airport) the engine inlet anti

- icing system should be capable of operating sufficiently to

assure safe operation for extended time in icing conditions . The straight

- line flight evaluation

must be evalu ated and may includ e the use of the cloud extent factor depicted in appendix C of

part s 25 or 29 . If the inlet is evaporative under continuous maximum icing conditions, and if it is running wet under intermittent maximum icing conditions, then the design is satisfactory. If the inlet is running wet in a maximum continuous atmospheric condition, then the applicant should

calculate the amount of ice that would accumulate during the straight

- line flight conditions

identified in Table C of this AC ( that is, c ontinuous maximum exposure followed by an intermittent maximum exposure). Ingestion of these calculated quantities of ice ( that is, the total

amount of ice includ ing runback and lip ice for the condition) should not result in more iceingestion damage tha n the amount of ice the engine was satisfactorily demonstrated to ingest during engine certification test s ( refer to § 33.77).

( 4) Running Wet Inlet Anti

- ice Systems

. For running wet systems, an alternative method of compliance would be to demo nstrat e the normal (no crew action required) ice build and shed cycles during exposures representative of descent, holding, or straight

- line flights ( t urboshaft engines can propose alternate representative exposures). This may be accomplished through

tank er or ground testing, with conservative corrections for water contents consistent with exposures discussed in paragraphs (1) through (3) above. Ingestion of this ice must not result in more damage than the amount of ice the engine has satisfactorily demon strated to ingest during engine certification tests (§ 33.77)

. The period of in

- service capability to be demonstrated may

vary with installation if the damage is not readily evident to the crew or visible on preflight inspection (for example

, tail mounted positions). Damage should be reviewed with the FAA and documented in the certification report.

d

. Two

- m inute Delayed Selection o f Inlet Anti

- I ce Accretion Analysis ( part 33 t urbofan and

t urboprop engines only)

. The § 33.77 testing criteria were d eveloped to account for historical means of compliance regarding the 2 minute delayed selection of inlet anti

- ice. For traditional pitot

- style engine inlet design s
, a part 23 or 25 applicant does not need to consider this scenario if the applicant shows c ompliance with § 33.77 testing criteria that utilizes a standard slab size

.

(1) For inlet designs o ther than the traditional pitot

- style inlets, a 2

- minute delay calculation may be required

, given that i nlet lip ice can form during the delayed a ctivation of the engine inlet anti

- ice system. Applicants should calculate the amount of inlet lip ice that forms using a continuous maximum condition depicted in appendix C of part 25, with a cloud extent factor of one. Of the total lip ice, only the ic e on the inner barrel side of the stagnation point would be ingested into the engine.

( 2 ) For the 2

- minute delay approach, a pplicants may assume that 1/3 of the ice on the inlet perimeter is ingested as one piece . This assumption is consistent with the historical approach taken by the engine manufacturers. T he maximum damage to the fan blade occur s at a high fan speed ; the critical condition occurs with the maximum inlet nacelle heat . T he refore, 2 minutes is

a reasonable time to include the pi lot reaction to select the anti

- ice system and

, for the engine thereafter, to shed the ice from the inlet perimeter

.

e

. Extended Operation s Performance Standards ( ETOPS ) ( p art 25 installations only)

. If

ETOPS certification is desired, the applicant should consider the maximum ETOPS diversion scenarios. For example, if a lightweight two

- engine diversion occurs at a 10,000 f oo t altitude followed by a cabin depressurization , then this put s the airplane into icing conditions that will

promote ice build up

. T herefore

, the inlet icing should be evaluated. I ngestion of the ice must not result in more ice ingestion damage than has been satisfactorily demonstrated during engine

certification ice ingestion ( refer to § 33.77).

f

. Potential A irframe Ice Sources

, Including Inlet s

. The engine manufacturer should consider the potential installation effects of the engine induction system and work with the installer to ensure that potential ice accretions formed at airframe ice accumulation sites are adequat elytested during the engine ice ingestion testing of § 33.77

. F uselage sources of ice accretion include the inboard section of the wing for an aft fuselage mounted engine, and a radome for any installation configuration

, to name two

. The induction syste m manufacturer or installer should

assess ice accumulations in accordance with parts 23, 25, 27

, or 29 (that is, §§ 23.901(d)(2), 23.1093, 25.1093 , 27.1093, or 29.1093 )

. They should also provide pertinent test variables to the engine manufacturer for inco rporation into the § 33.77 test demonstration

. If an application or product inlet has not been selected by the time of engine certification, the engine manufacturer

should list all pertinent inlet assumptions, test data

, and results in the engine installa tion manual for use by the future installer.

( 1 ) It is normally sufficient to show that ice accretions from these sites are smaller in size

, and therefore have less or equal kinetic energy potential than the ice slab used in § 33.77 test ing

. As de scribed in paragraph 10.e of this AC , the slab typically breaks into pieces when entering

the inlet. For the comparison of airframe i ce to the results of the § 33.77 test, i t may be assumed

that airframe ice sources break in half when entering the inle t.

(2) The applicant may elect to compare the ice based on the kinetic energy of the ice slab. Kinetic energy may be used as an acceptable method for comparing the airframe ice source to

the results of the § 33.77 test. Any kinetic energy meth od must be agreed to by the applicant’s

A CO

.

(3) Demonstration of ice shedding with natural or icing tanker testing can be an acceptable

means of compliance for airframe ice that accumulat es in front of the engine , while in the

conditions depict ed in a ppendix C of part 25 or 29

. However, the applicant must demonstrate conservatism relative to long durations in the se icing conditions (such as for several successive cloud extent exposures)

. Ingestion of shed ice during natural or icing tanker tes ting must not result in more damage than the amount of ice the engine was satisfactorily demonstrated to ingest during engine certification tests ( refer to § 33.77)

. The period of in

- service capability to be demonstrated may vary with the installation if the damage is not readily evident to the crew or visible on preflight inspection (for example

, tail mounted positions).

g

. Wing Sourced Ice f or Aft Mounted Engines ( part 33 t urbofan and t urboprop e ngines, only)

. Clear ice may occur on the upper sur faces of the wing when cold

- soaked fuel (due to aircraft prolonged operation at high altitude) remains in contact with the fuel tanks’ upper

surfaces after landing , and during time on the ground when the airplane is exposed to conditions of atmospheric moi sture ( for example, fog, precipitation, and condensation of humid air) at ambient temperatures above freezing. A tmospheric moisture, when in contact with cold wing surfaces, may freeze. If undetected and still present during takeoff, the ice is most like ly to shed when the wings flex at takeoff rotation. Simultaneous ice shedding from both wings of an airplane with aft mounted engines has resulted in ice ingestion damage and power loss in both engines during takeoff.

(1) When fuel management p rocedures allow wing fuel tanks to remain mostly full after landing , cold fuel is in contact with the wing upper skin . As a result, clear ice may form on cold wing surface s during the time on the ground ; this holds true even if conditions remain abovefre ezing and are not expected to be, or recognized as, icing conditions

. Fuel temperature may need to be analyzed

, including the effect of fuel schedules and refueling procedures.

h. Aft Mounted Engine Installations

. For aircraft with aft mounted eng ines, applicant s should demonstrate that any undetected ice in front of engine inlets is not greater than what was shown during engine and aircraft certification . A pplicants should a lso demonstrate that shedding

the ice accumulated due to cold

- soaked fuel will not result in hazardous engine operation.

(1) For fixed wing applications, a wing mounted ice detection system may allow minimal

accretion prior to alerting the flight crew as long as the accumulated ice thickness is limited to a
defined v alue

. This defined value should demonstrate a sufficient margin relative to both engine ingestion (refer to § 33.77 ) and performance and handling qualities as demonstrated under § 25.1419 . T he largest clear ice slab area that could be ingested into an en gine is the same area as the engine’s inlet highlight area. In cases where this large slab is very thin

, i t may not be possible to reach the engine in one piece . T herefore

, consistent with a conservative assessment of slab “break up” that can occur withi n the air stream ahead of the fan

, the applicant may

consider the ice slab as multiple pieces. Y ou may evaluate any potential resulting damage to individual engine components

, such as fan blades

, by comparison to other ice ingestion tests

such as the ice slab discussed in paragraph 10 of this AC . Additionally, you may compare the ingestion results to other soft body ingestion test results, such as the 2

- inch hail ingestion tests

, as long as the mass and kinetic energy effects are accounted for

. Finally i t may be appropriate to address the potential effect on the engine fan shaft

, or other structure

, if the combined impact of

a large sheet of ice (or multiple pieces) striking many fan blades simultaneously could generate unacceptable loads on the engine co mponents. This might in some instances be addressed by analysis.

i

. Compliance w ith Mixed Phase a nd Glaciated Atmospheric Conditions Under § 25.1093(b)(1) ( part 33 t urbofan and t urboprop e ngines only)

. The results of previous airfoil testing in a mixed phase icing environment in dicat e s these icing conditions do not appreciably accrete on unheated aircraft wings. Furthermore the testing showed mixed phase environment

results in the same or less ice accretion than an equivalent amount of supercooled liquid water.

T he overall power required by the running

- wet ice protection system was essentially unchanged

between all

- liquid and mixed phase conditions.

(1) However, in the running

- wet mode, the local power density was much higher around

the stagnation area in the mixed phase conditions

, compared to the purely liquid conditions.

This is due to the power required to offset the thermodynamic heat

- of

- fusion necessary to melt the impacting ice particles that either fully or partially stick to th e surface.

( 2 ) This may also explain why pitot

- style inlets have not proved to be susceptible to mixed phase ice accretion within the inlet

, and why the compliance methods to § 25.1093 for the icing

conditions depicted in a ppendix C of part 25

, have historically adequately address ed those inlets

. Engines designed with reverse flow intakes, or with intakes involving considerable changes in

airflow direction should demonstrate acceptable operation in the icing conditions depicted in

a ppendix D of p art 33. Compliance for pitot

- style inlets, without considerable changes inairflow direction, may be shown through qualitative analysis of the design

, and supported by similarity to previous design s that have shown successful service histories

.

j

. Compliance w ith t he Supercooled Large Drop (SLD) Conditions Under § 25.1093(b)(1) ( part 33 t urbofan and t urboprop e ngines only)

. The icing conditions depicted in appendix O of part 25 are not applicable to turbine engine induction systems on airplanes wi th a maximum takeoff weight equal to or greater than 60,000 pounds.

(1) In

- F light Exposure to Appendix O of Part 25 Conditions

.

(a) For airframe ice sources , the installer may use either quantitative or qualitative analysis of th e design . The qualitative analysis must be supported by similarity to previous design s that have shown a successful service history in order to have confidence that the historical methodology fo r certification as represented by § 33.77

, Table 1

– “ Minimum Ice Slab Dimensions Based on Engine Inlet Size ”, is appropriate.

1

. For engine inlets, compliance may be shown through quantitative or qualitative analysis of the design . Qualitative analysis must be supported by similarity to previo us design s that ha ve shown a successful service history . C ompliance to the “Large Droplet Condition” of Table 1 of § 25.1093 (b)(2) , titled “Table 1

- Icing Conditions for Ground Tests ”, would still be required for ground idle taxi conditions. If similari ty is not shown, then an assessment of inlet impingement limits, differing catch efficiency, distribution effects, and water contents depicted

in appendix O of part 25 should be accomplished.

2

. To demonstrate compliance with § 25.1093 , the exposures should be consistent

with

, but conservative to the aircraft operation al approvals, as well as the applic able definitions

in Part II of a ppendix O of part 25

, for both the ice protection system and the potential airframe

accumulations with r espect to shedding.

(2) Ground Taxi Exposure t o Appendix O of Part 25 Conditions

. The s ervice experience considered for this advisory material indicates that engine fan damage events exist which

resulted from exposure to SLD during ground taxi operations. For this reason, an additional compliance condition was added to § 25.1093 (b)( 2 ) for a 30

- minute, idle power exposure to SLD on the ground. Applicants should include the terminal falling velocity of SLD ( for example, freezing rain, freezing d rizzle) in their trajectory assessment

, relative to the protected sections of

the inlet. The 100 micron minimum mean effective diameter ( MED ) was selected as a reasonable achievable test condition

, given current technology. We recommend however, that

app licant s who choos e to certify by analysis

, should evaluate engine operation in the icing

conditions depicted in a ppendix O of part 25 . This evaluation should include various drop sizes

, with the appropriate drop size distribution

, to find a critical condi tion.

k

. N atural Icing Flight Tests

. Natural icing flight tests are intended to demonstrate that each turbine engine is capable of operating throughout the flight power range of the engine (including idling), without an adverse e ffect . This includ es the accumulation of ice on the engine, inlet system components, or airframe c omponents that would have an adverse e ffect on the engine operation or cause a serious loss of power or thrust. The icing conditions depicted in appendix Oof part 25 are only applicable to turbine engine induction systems on part 25 airplanes with a maximum takeoff weight less than 60,000 pounds.

(1) Based on multiple natural ice engine damage and operability events experienced on natural icing flight test s and in

- s ervice airplanes, natural ice encounters must be used to show compliance with §§ 23.1093(b)(1) or 25.1093(b)(1) , and in some cases for §§ 27.1093(b)(1) and 29.1093(b)(1)

. F or airplanes that are not intended to be certificated for flight in icing condition s, other methods may be used to replace the natural icing flight tests . When replacing the natural icing flight tests on these airplanes, you can show compliance to § 23.1093(b)(1) by analysis such as ground testing, dry air flight testing, and similarity

.

(2) In addition to showing your engine inlet icing analysis model is accurate , several other key issues exist that the natural ice encounter addresses. These include :

• T he adequacy of flight crew procedures when operation in icing conditions ;

• T he acceptability of control indications to the flight crew as the airplane responds to engine fan blade ice shedding during various conditions ;

• T he performance of the engine vibration indication system

, as well as other engine indication systems ; and

• T he confirmation that the powerplant installation performs satisfactorily while in icing conditions. This powerplant installation is comprised of the engine, inlet,

and the anti

- ice system.

l

. Identification o f I ce S ource

. A pplicants should provid e a means to identify the source of any ice that may be ingested by the engine during the natural icing certification testing. Special attention should be given to non

- representative ice accretions on flight test instrumentation probes or other surfaces f orward of the engine during prolonged operation in icing conditions.

m

. Icing T est P oint M onitoring

. The applicant should provide acceptable monitoring of the icing test point condition (that is, LWC, droplet diameter, temperature) for sufficient t ime

exposure to ensure the icing encounter is representative of the environmental icing conditions depicted in a ppendices C and O of part 25 certification icing conditions , appendix C of part 29

,

or appendix D of part 33 conditions , as applicable

.

n

. Compliance

. Compliance with §§ 23.1093 or 25.1093 is required even if flight into icing

conditions is not intended or approved (§§ 23.1419 or 25.1419 compliance)

. Applicants must therefore, propose an acceptable icing engine

- installation test. The pr oposed test should include

the natural ice encounter criteria. The icing conditions depicted in appendix O of part 25 are only applicable to turbine engine induction systems on part 25 airplanes with a maximum takeoff weight less than 60,000 pounds.

(1) T ypically

, an adequate test sequence includes three natural fan ice shed cycles at each of the following conditions (inlet anti

- ice turned "on") : descent (flight

- idle , or anti

- ice idle ), holding (power set to maintain level flight through a range of anticipated airplane gross weight), and maximum climb, unless a more critical engine power setting exists. These fan shed cyclesshould result from natural ice accumulation and shedding

, not induced or forced shedding by

throttle or power lever excurs ions or manipulations during each condition.

(2) Th is test should be conducted at a steady

- state engine thrust level . The test may also involve flying through the same icing cloud multiple times (lapping) for the fan to accumulate enough ice fo r a shed cycle to occur

. The airplane may exit the icing conditions between each fan shed cycle as needed to clear any other unprotected airplane surfaces from ice. Applicants should ensure the test engine ignition system is off during the icing conditio ns. This will avoid masking any adverse engine operating conditions dur ing the natural icing encounter. This may require pulling several airplane circuit breakers to disable the test engine’s auto

- ignition or recovery system, or both . We recommend that applicants establish engine damage criteria and obtain FAA approval of th ese criteria before conducting the natural ice encounter test. Our experience s ha ve shown that pass or fail criteria

, such as allowable limits of engine damage

, are best agreed on be fore certification testing.

o

. F alling a nd Blowing Snow

. Sections 23.1093(b)(1)(ii)

, 25.1093(b)(1) , 27.1093(b)(ii), and 29.1093(b)(ii) require engines to operate satisfactorily in falling and blowing snow throughout the flight power range, and grou nd idle. Ground idle would include aircraft operations on or around the airport environment

, both before and af ter flight. Accordingly, the effect of ingesting snow during ground operations should be evaluated

.

(1) C onsistent with service expe rience

, a irports continue operat ions with falling snow concentrations that result in a 0.25 mile or less visibility (about 0.9 gm/m 3 , see paragraph 9.g.(2 ) of this AC)

. For turbine engines operating in these conditions, i n

- flight service experience

shows that snow shed s from engine or aircraft accumulation sites can cause severe operability

affects when ingest ed by these engines

. T ypical turbofan or turbo jet installations, using simple pitot (stra ight duct) inlets, have minimal areas for snow accumulation

. For these inlets, in

- flight icing tests have generally been found to be more critical than snow tests. The applicant should

evaluate e ngine inlets that incorporate plenum chambers, screens, particle

- s eparators, variable

geometry, inlet barrier filters (rotorcraft), or any other featur e , such as an oil

- cooler, struts or fairings, which may provide a potential accumulation site for snow. Further , any airframe accumulation sites upstream of the engine inlet should also be considered by the applicant

.

(2) A temperature range of 25 0 to 32 0 Fahrenheit (

- 4 0 Celsius to + 0 0 Celsius) is common in

a heavy snow environment

, and can result in “wet sticky snow” which may accumulate on unheated surfaces (airframe and engine) that are subject to impingement. E ngine service events

have demonstrated that a snow environment is conducive to ice accretion behind the fan in front compressor stages of the engine, at low engine power. Service experience has demonstrated compressor damage as a result of exposure to prolonged periods of falling snow during ground operation. The new engine test for snow conditions is developed to represent the engine

conditions where snow can form accretions aft of the first stage compressor ( for example, the fan on turbofan engines) on the core inlet and first stages of the core compressor . The applicant should choose a temperature within the range provided above to achieve icing behind the first stage compressor ( for example, the fan for turbofan engines)

.p

. T est Results

. T he applicant should carefully consider all evidence of ingestion and damage to the engines

, and their potential sources. If damage is incurred, the possible test

outcomes include:

(1) Acceptable D amage . The extent of damage is equivalent to or less than what was incurred and accepted during the engine certification testing.

(2) All S ystems O perating N ormally . The extent of damage is equivalent or less than what

was incurred and accepted during the § 33.68 tests.

(3) Delay ed A ctivation o f I nduction S ystem A nti

- I cing . If ice ingestion tests under § 33.77 do not adequately represent the particular airframe installation, then the delayed anti

icing system activation test should be considered ( c aution should be used and a ll sa fety precautions exercised )

. For this condition, the acceptance criteria d escribed in paragraph 10 of this AC should be used. The airframe manufacturer still must consider all potential ice shedding sites (for example, inboard wing and radome).

(a) Similar to the § 33.77 ice slab ingestion test compliance , the use of engine auto

ignition and recovery systems are allowed when show ing compliance with the delayed activation tests of parts 23

, 25 , 27, or 29 , as long as these automatic systems cannot be easily turned off by the flight crew

.

(b) The difference in anti

- iced inlets versus de

- iced inlets is significant

. Anti

- iced inlets are designed so that no ice accretes on the inlet while operating in icing conditions. De

iced i nlets allow some ice accretion

, but are designed for a cyclic shedding of ice from the engine inlet into the engine . De

- iced inlets typically incorporate, as part of their design, an inlet particle

- separator that stops the ingestion of ice into the core o f the engine. E ngine auto

recovery system s should not be a compensating design feature utilized to minimize the negative effects of an inadequate particle

- separating inlet that is not in full compliance with §§ 23.1093

, 25.1093 , 27.1093, or 29.1093

.

( 4 ) Damage F rom T esting i n N on

- R epresentative I cing C onditions

. When d amage result s from icing test conditions that fall significantly outside the icing envelopes depicted in appendices C and O of part 25, appendix C of part 29, or appen d i x D of par t 33

, or when the flight test is conducted in an abnormal manner and results in excessive ice shed damage, this

may result in a test failure relative to the pretest pass or fail criteria. Any abnormal conditions

should be discussed with the FAA to determi ne if the test can be deemed “passed ”

. An example

of an abnormal operation is flying with one engine at idle while the aircraft is operated in level flight.

( 5 ) Unacceptable D amage . The icing test conditions are representative of in

- service

en counters , but t he resultant airframe or engine ice sheds caused damage that exceed ed the

acceptance criteria described in paragraph 8. a.(1) of this AC.

q. Compliance for Auxiliary Power Unit (APU) Installations

. Compliance for an auxiliary

power un it (APU) installation is demonstrated under §§ 23

. 901(f), 25.901(d) or 29.901(d) , butAPUs must also be addressed and shown as acceptable for installation under §§ 23. 1093(b)

, 2 5.1093(b) or 29.1093 (b) (see section 11 of this AC). An essential APU is used to provide air and/or power necessary to maintain safe aircraft operation. A non

- essential APU is used to provide air and/or power as a matter of convenience and may be shut down without jeopardizing safe aircraft operation. The icing conditions depicted in appendix O of part 25 are not applicable

to APU induction systems on airplanes with a maximum takeoff weight equal to or greater than 60,000 pounds.

(1) Non

- Essential APU

. A n applicant for a transport category airplane with a non

essential APU installation that is restricted from use in icing conditions is not required to show compliance with § 2 5.1093(b). This includes the icing conditions depicted in appendices C and O of part 25

, appendix D of part 33

, and in falling and blowing snow con ditions. An applicant for a transport category airplane with a n on

- essential APU installation that is not restricted from use in icing conditions or falling and blowing snow must show compliance with § 2 5.1093(b)

. To do this, you must demonstrate that th e APU’s operation will not affect the safe operation of the airplane when subject to those conditions, both on the ground and in

- flight.

(2) Essential APU

. An applicant for a transport category airplane with an essential APU installation is req uired to demonstrate that it can safe ly operat e in natural icing conditions, both on the ground and in

- flight. Thus

, if you are applying for a transport category airplane icing approval with an essential APU installation , you must show compliance with § 2 5.1093(b). This

includes the icing conditions depicted in appendices C and O of part 25, and appendix D of part 33, and in falling and blowing snow within the limitations established for the airplane for such operation.

(a) Acceptable oper ation of an APU in the icing conditions depicted in appendix C to part 25 should be demonstrated using the same compliance methods for turbine engine

installations described in this AC.

(b) Acceptable operation in the icing conditions depic ted in appendix O to part 25 and appendix D to part 33

, should be demonstrated using the same compliance methods for turbine engine installations described in this AC. However, many transport category airplane APU

installation s have an inlet located in th e empennage section of the airplane fuselage that can generate a shadowing effect , and thereby prevent large icing droplets (SLD) and ice crystals from entering the APU inlet

. For this type of installation, the § 25.1093 compliance for the icing

environme nt depicted in appendix O to part 25 and appendix D to part 33

, may be demonstrated by an acceptable validated analysis

. To be acceptable, a validated analysis should show that icing conditions depicted in appendix O to part 25 and appendix D to part 33, are adequately addressed by the compliance demonstration for the conditions depicted in appendix C to part 25.

(c) APU c ompliance with the falling and blowing snow conditions, within the limitations established for the airplane for such ope ration

, should be demonstrated using the same compliance methods for turbine engine installations described in this AC

.1 2 . C onclusion

. Although applicants may conduct representative tests under §§ 33.68, 33.77, 23.1093, 25.1093 , 27.1093, and 29.1093

, flight test events during aircraft certification may still occur which appear inconsistent. In all likelihood, these results will not be inconsistent when

evaluated based on the scope, intent, and limitations of the certification testing. Only through

ap propriate test criteria, appropriate critical point selection, reliable instrumentation

, and adequate documentation, can the impact of icing on turbine engines be fully addressed . Because of the relative frequen cy with which operators experience icing enc ounters

, and because of the potential adverse impact on safety, we recommend all applicants adopt a conservative approach

when establishing icing testing plans and demonstrations of compliance.10/22/14 AC 20

- 147A

This document does not represent final agency action on this matter and should not be viewed as a guarantee that any final action will follow in this or any oth er form. A PPENDIX A

Reference: L ow engine speed part 33 compliance t esting ( part 33 t urbofan and t urboprop e ngines only)

1. The Federal Aviation Administration (FAA ; us ), Engine and Propeller Directorate

, Standards Staff, ANE

- 110, has been requested to provide a position relative to test time duration during § 33.68 ici ng tests at non

- flight sustainable engine speeds. The following paragraphs provide a technical rationale for the standardized compliance test times for Appendix C conditions. The

FAA ’ s position has traditionally been, and continues to be , that engines ar e held to a higher standard than aircraft for icing certification and the engines must demonstrate unrestricted and unlimited operation in icing conditions. Therefore, during the § 33.68 compliance test demonstration, engines must remain on

- point until st able operation is demonstrated.

Some applicants have described this requirement as being burdensome when operating the engine at non

- flight sustainable rotor speeds (meaning, below lightweight hold engine operation). In response to this observation, the FAA has developed criteria that allow for a reasonable test period and termination time for icing testing at non

- flight sustainable engine speeds. It is emphasized that this criteria is only applicable to icing testing when the engine is operating at non

- flight sustainable engine speeds.

2. The maximum intermittent maximum liquid water cont ent (LWC) severity depicted in

a ppendix C of part 25 is over three times more severe than the continuous maximum LWC

. Due to the significant differences in severity b etween intermittent maximum and continuo us maximum icing conditions depicted in a ppendix C of part 25 , two separate criteria have been

considered

. T hey are : (1) short test sequences for the more severe intermittent maximum conditions, and (2) longer test sequences for continuous maximum conditions

.

3

. Appendix C of part 25 , paragraph (a)

- Continuous Maximum Icing Conditions

. In an effort to preserve flight safety in icing conditions, the FAA strives to retain a margin of safety through

suitable demonst rations of compliance. W e have established the criteria for part 33 icing testing through c onservative assessments

, the better to ensure that compliance demonstrations are reflective of real world conditions

. Our rationale is detailed below.

a. Co nservative t est perspective

. E ngine designs should be capable of operations exceed ing the single cloud definitions depicted in appendix C of part 25

. A robust design is necessary

, since an encounter of LWC at a stated inlet temperature

, at any mean effec tive diameter (MED) is a probable event for the cloud length specified

. This expected occurrence rate is too frequent to be addressed by tests durations based on a single cloud extent. Thus for demonstration of compliance with § 33.68

, we use a test dura tion representative of multiple cloud extents

.

A 1

- 110/22/14 AC 20

- 147A

b. C onservative test period: icing test demonstrations at non

- flight sustainable speeds . T he FAA has developed a conservative methodology for assessing the worst case condition to be expected in service (that is, long descent loiter). The following analysis is intended to show the

rationale behind the test duration we have selected for low engine speed testing under § 33.68.

Th is analysis is not based on expected or representative flight prof iles or typical icing conditions.

(1) A typical descent gradient for pilots is 3 miles horizontal transit for every 1000

- foot vertical descent (15.8:1 gradient). Approximately, t his work s out to a 60

- mile descent initiation test point for a 20, 000

- foot descent , with a nominal descent slope of about 3.6 degree s

. If a conservative 2

- degree descent slope is considered, then the descent initiation would start at 120 miles out for a 22,000

- foot vertical descent. A total descent altitude of 22,000 f eet above ground level (agl) was used because Appendix C goes from sea level to 22,000 feet, although it is recognized that icing occurs at altitudes greater than 22,000

.

(2) In order to convert a descent gradient into a meaningful descent rate in feet each minute, a horizontal airspeed must be assumed. For this conservative analysis, an average horizontal air speed of 220 knots (250 miles per hour) is assumed. This average airspeed is for the complete descent through all 22,000 feet

, and assum es periodic head winds and profile variations for all types of fixed wing aircraft. This makes the assumption more conservative for transport category aircraft and less conservative for small lightweight prop eller aircraft. T his is a conservative approxi mation above 10,000 feet agl

, where speed is not limited.

(3) A t a 2

- degree descent gradient and 220 knots (250 mph) horizontal air s peed

, t he result of these inputs is about a 750 f ee t descent each minute from 22,000 feet

, for a total time of 3 0

minutes

. Consequently, when stabilized operation has not been demonstrated

, it is reasonable that engine manufacturers should test engines at flight

- idle and up to engine speeds that result in

non

- sustainable flight (that is, less than light hold power) for a total of not more than 30

- minutes of continuous maximum icing operation in the conditions depicted in appendix C of part 25

.

4

. Appendix C of part 25

, part 1, paragraph (b)

- Intermittent Maximum Icing Conditions

. T he

intermitte nt maximum icing c onditions depicted in a ppendix C of part 25

, are more severe than

the continuous maximum conditions. Just as the previous continuous maximum discussion provides for multiple cloud extents in a series, the same philosophy will be applied here for

intermitt ent maximum conditions, to assure robust design demonstrations. Th e increased severity of an intermittent cloud depicted in a ppendix C

, in conjunction with the reduced

horizontal and vertical extents of these clouds in a ppendix C

, necessitates a proportio nately reduced maximum compliance test demonstration period.

The FAA has a very long and successful historical experience base for compliance test demonstrations in intermittent maximum icing conditions. This historical basis has shown that a

10

- minute t est period has been sufficient. Additionally

, the 10

- minute period would represent about five to ten sequential inte rmittent maximum clouds on a horizontal extent basis

, depending

on aircraft speed, as depicted in a ppendix C of part 25

.

A 1

- 210/22/14 AC 20

- 147A

6

. Conclus ion: A ppendix A of this AC provide s the rationale for the standard test time

durations described in this AC and the icing rule ( 14 CFR 33.68 )

. In summary, shorter compliance test demonstration periods have been rationalized here for low power descent

con ditions and for intermittent maximum conditions where there are shorter cloud extents.

A1

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- 147A

APPENDIX B

Advisory Circular Feedback Information

If you have comments or recommendations for improving this advisory cir cular (AC), or suggestions for new items or subjects to be added, or if you find an error, you may let us know about by using this page as a template and 1) emailing it to 9

- AWA

- AVS

- AIR500

- Coord@faa.gov or 2) faxing it to the attention of the AIR Directive s Management Officer at 202

- 267

- 3983.

Subject (insert AC number and title) Date: (insert date)

Comment/Recommendation/Error: (Please fill out all that apply)

An error has been noted:

Paragraph ____________________

Page ______

Type of error (check all that apply): Editorial:

----- Procedural

-----

Conceptual____

Description/Comments:______________________________________________

_________________________________________________________________

_________________________________________________________________

Recommend paragraph ______ on page ______ be changed as follows: (attach separate sheets if necessary)

_________________________________________________________________

In a future change to this advisory circular, please include coverage on the following subject: (briefly describe what you want added attaching separate sheets if necessary)

________________________________________________________________

Name: __________________________

A 2

- 1