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(a)
This part prescribes airworthiness standards for the
issue of type certificates and changes to those
certificates, for aircraft engines.
(b)
Each person who applies under part 21 for such a
certificate or change must show compliance with the
applicable requirements of this part and the
applicable requirements of part 34 of this chapter.
Each
applicant must show that the aircraft engine
concerned meets the applicable requirements of this
part.
The
applicant must prepare Instructions for Continued
Airworthiness in accordance with appendix A to this
part that are acceptable to the Administrator. The
instructions may be incomplete at type certification
if a program exists to ensure their completion prior
to delivery of the first aircraft with the engine
installed, or upon issuance of a standard
certificate of airworthiness for the aircraft with
the engine installed, whichever occurs later.
Each
applicant must prepare and make available to the
Administrator prior to the issuance of the type
certificate, and to the owner at the time of
delivery of the engine, approved instructions for
installing and operating the engine. The
instructions must include at least the following:
(a)
Installation instructions. (1) The location
of engine mounting attachments, the method of
attaching the engine to the aircraft, and the
maximum allowable load for the mounting attachments
and related structure.
(2)
The location and description of engine connections
to be attached to accessories, pipes, wires, cables,
ducts, and cowling.
(3)
An outline drawing of the engine including overall
dimensions.
(b)
Operation instructions. (1) The operating
limitations established by the Administrator.
(2)
The power or thrust ratings and procedures for
correcting for nonstandard atmosphere.
(3)
The recommended procedures, under normal and extreme
ambient conditions for—
(i)
Starting;
(ii)
Operating on the ground; and
(iii) Operating during flight.
(c)
Safety analysis assumptions.
The assumptions of the safety analysis as described in
§33.75(d) with respect to the reliability of safety
devices, instrumentation, early warning devices,
maintenance checks, and similar equipment or
procedures that are outside the control of the
engine manufacturer.
33.7 Engine
ratings and operating limitations.
(a)
Engine ratings and operating limitations are
established by the Administrator and included in the
engine certificate data sheet specified in §21.41 of
this chapter, including ratings and limitations
based on the operating conditions and information
specified in this section, as applicable, and any
other information found necessary for safe operation
of the engine.
(b)
For reciprocating engines, ratings and operating
limitations are established relating to the
following:
(1)
Horsepower or torque, r.p.m., manifold pressure, and
time at critical pressure altitude and sea level
pressure altitude for—
(i)
Rated maximum continuous power (relating to
unsupercharged operation or to operation in each
supercharger mode as applicable); and
(ii)
Rated take-off power (relating to unsupercharged
operation or to operation in each supercharger mode
as applicable).
(2)
Fuel grade or specification.
(3)
Oil grade or specification.
(4)
Temperature of the—
(i)
Cylinder;
(ii)
Oil at the oil inlet; and
(iii) Turbosupercharger turbine wheel inlet gas.
(5)
Pressure of—
(i)
Fuel at the fuel inlet; and
(ii)
Oil at the main oil gallery.
(6)
Accessory drive torque and overhang moment.
(7)
Component life.
(8)
Turbo-supercharger turbine wheel r.p.m.
(c)
For turbine engines, ratings and operating
limitations are established relating to the
following:
(1)
Horsepower, torque, or thrust, r.p.m., gas
temperature, and time for—
(i)
Rated maximum continuous power or thrust
(augmented);
(ii)
Rated maximum continuous power or thrust (unaugmented);
(iii) Rated take-off power or thrust (augmented);
(iv)
Rated take-off power or thrust (unaugmented);
(v)
Rated 30-minute OEI power;
(vi)
Rated 21/2-minute OEI power;
(vii) Rated continuous OEI power; and
(viii) Rated 2-minute OEI Power;
(ix)
Rated 30-second OEI power; and
(x)
Auxiliary power unit (APU) mode of operation.
(2)
Fuel designation or specification.
(3)
Oil grade or specification.
(4)
Hydraulic fluid specification.
(5)
Temperature of—
(i)
Oil at a location specified by the applicant;
(ii)
Induction air at the inlet face of a supersonic
engine, including steady state operation and
transient over-temperature and time allowed;
(iii) Hydraulic fluid of a supersonic engine;
(iv)
Fuel at a location specified by the applicant; and
(v)
External surfaces of the engine, if specified by the
applicant.
(6)
Pressure of—
(i)
Fuel at the fuel inlet;
(ii)
Oil at a location specified by the applicant;
(iii) Induction air at the inlet face of a
supersonic engine, including steady state operation
and transient overpressure and time allowed; and
(iv)
Hydraulic fluid.
(7)
Accessory drive torque and overhang moment.
(8)
Component life.
(9)
Fuel filtration.
(10)
Oil filtration.
(11)
Bleed air.
(12)
The number of start-stop stress cycles approved for
each rotor disc and spacer.
(13)
Inlet air distortion at the engine inlet.
(14)
Transient rotor shaft over speed r.p.m., and number
of over-speed occurrences.
(15)
Transient gas over temperature, and number of over
temperature occurrences.
(16)
For engines to be used in supersonic aircraft,
engine rotor windmilling rotational r.p.m.
(a)
Requested engine power and thrust ratings must be
selected by the applicant.
(b)
Each selected rating must be for the lowest power or
thrust that all engines of the same type may be
expected to produce under the conditions used to
determine that rating.
This
subpart prescribes the general design and
construction requirements for reciprocating and
turbine aircraft engines.
The
suitability and durability of materials used in the
engine must—
(a)
Be established on the basis of experience or tests;
and
(b)
Conform to approved specifications (such as industry
or military specifications) that ensure their having
the strength and other properties assumed in the
design data.
(a)
The design and construction of the engine and the
materials used must minimize the probability of the
occurrence and spread of fire. In addition, the
design and construction of turbine engines must
minimize the probability of the occurrence of an
internal fire that could result in structural
failure, overheating, or other hazardous conditions.
(b)
Except as provided in paragraphs (c), (d), and (e)
of this section, each external line, fitting, and
other component, which contains or conveys flammable
fluid must be fire resistant. Components must be
shielded or located to safeguard against the
ignition of leaking flammable fluid.
(c)
Flammable fluid tanks and supports which are part of
and attached to the engine must be fireproof or be
enclosed by a fireproof shield unless damage by fire
to any non-fireproof part will not cause leakage or
spillage of flammable fluid. For a reciprocating
engine having an integral oil sump of less than
25-quart capacity, the oil sump need not be
fireproof nor be enclosed by fireproof shield.
(d)
For turbine engines type certificated for use in
supersonic aircraft, each external component which
conveys or contains flammable fluid must be
fireproof.
(e)
Unwanted accumulation of flammable fluid and vapor
must be prevented by draining and venting.
(a)
Engine design and construction must minimize the
development of an unsafe condition of the engine
between overhaul periods. The design of the
compressor and turbine rotor cases must provide for
the containment of damage from rotor blade failure.
Energy levels and trajectories of fragments
resulting from rotor blade failure that lie outside
the compressor and turbine rotor cases must be
defined.
(b)
Each component of the propeller blade pitch control
system which is a part of the engine type design
must meet the requirements of 35.42 of this chapter.
Engine design and construction must provide the
necessary cooling under conditions in which the
airplane is expected to operate.
(a)
The maximum allowable limit and ultimate loads for
engine mounting attachments and related engine
structure must be specified.
(b)
The engine mounting attachments and related engine
structure must be able to withstand—
(1)
The specified limit loads without permanent
deformation; and
(2)
The specified ultimate loads without failure, but
may exhibit permanent deformation.
The
engine must operate properly with the accessory
drive and mounting attachments loaded. Each engine
accessory drive and mounting attachment must include
provisions for sealing to prevent contamination of,
or unacceptable leakage from, the engine interior. A
drive and mounting attachment requiring lubrication
for external drive splines, or coupling by engine
oil, must include provisions for sealing to prevent
unacceptable loss of oil and to prevent
contamination from sources outside the chamber
enclosing the drive connection. The design of the
engine must allow for the examination, adjustment,
or removal of each accessory required for engine
operation.
[Amdt.
33–10, 49 FR 6851, Feb. 23, 1984]
(a)
Turbine, compressor, fan, and turbo-supercharger
rotors must have sufficient strength to withstand
the test conditions specified in paragraph (c) of
this section.
(b)
The design and functioning of engine control
devices, systems, and instruments must give
reasonable assurance that those engine operating
limitations that affect turbine, compressor, fan,
and turbo-supercharger rotor structural integrity
will not be exceeded in service.
(c)
The most critically stressed rotor component (except
blades) of each turbine, compressor, and fan,
including integral drum rotors and centrifugal
compressors in an engine or turbo-supercharger, as
determined by analysis or other acceptable means,
must be tested for a period of 5 minutes—
(1)
At its maximum operating temperature, except as
provided in paragraph (c)(2)(iv) of this section;
and
(2)
At the highest speed of the following, as
applicable:
(i)
120 percent of its maximum permissible r.p.m. if
tested on a rig and equipped with blades or blade
weights.
(ii)
115 percent of its maximum permissible r.p.m. if
tested on an engine.
(iii) 115 percent of its maximum permissible r.p.m.
if tested on turbo-supercharger driven by a hot gas
supply from a special burner rig.
(iv)
120 percent of the r.p.m. at which, while cold
spinning, it is subject to operating stresses that
are equivalent to those induced at the maximum
operating temperature and maximum permissible r.p.m.
(v)
105 percent of the highest speed that would result
from failure of the most critical component or
system in a representative installation of the
engine.
(vi)
The highest speed that would result from the failure
of any component or system in a representative
installation of the engine, in combination with any
failure of a component or system that would not
normally be detected during a routine preflight
check or during normal flight operation.
Following the test, each rotor must be within
approved dimensional limits for an overspeed
condition and may not be cracked.
Each
control system which relies on electrical and
electronic means for normal operation must:
(a)
Have the control system description, the percent of
available power or trust controlled in both normal
operation and failure conditions, and the range of
control of other controlled functions, specified in
the instruction manual required by 33.5 for the
engine;
(b)
Be designed and constructed so that any failure of
aircraft-supplied power or data will not result in
an unacceptable change in power or thrust, or
prevent continued safe operation of the engine;
(c)
Be designed and constructed so that no single
failure or malfunction, or probable combination of
failures of electrical or electronic components of
the control system, results in an unsafe condition;
(d)
Have environmental limits, including transients
caused by lightning strikes, specified in the
instruction manual; and
(e)
Have all associated software designed and
implemented to prevent errors that would result in
an unacceptable loss of power or thrust, or other
unsafe condition, and have the method used to design
and implement the software approved by the
Administrator.
(a)
Unless it is constructed to prevent its connection
to an incorrect instrument, each connection provided
for powerplant instruments required by aircraft
airworthiness regulations or necessary to insure
operation of the engine in compliance with any
engine limitation must be marked to identify it with
its corresponding instrument.
(b)
A connection must be provided on each turbojet
engine for an indicator system to indicate rotor
system unbalance.
(c)
Each rotorcraft turbine engine having a 30-second
OEI rating and a 2-minute OEI rating must have a
provision for a means to:
(1)
Alert the pilot when the engine is at the 30-second
OEI and the 2-minute OEI power levels, when the
event begins, and when the time interval expires;
(2)
Determine, in a positive manner, that the engine has
been operated at each rating; and
(3)
Automatically record each usage and duration of
power at each rating.
This
subpart prescribes additional design and
construction requirements for reciprocating aircraft
engines.
The
engine must be designed and constructed to function
throughout its normal operating range of crankshaft
rotational speeds and engine powers without inducing
excessive stress in any of the engine parts because
of vibration and without imparting excessive
vibration forces to the aircraft structure.
Each
turbocharger case must be designed and constructed
to be able to contain fragments of a compressor or
turbine that fails at the highest speed that is
obtainable with normal speed control devices
inoperative.
(a)
The fuel system of the engine must be designed and
constructed to supply an appropriate mixture of fuel
to the cylinders throughout the complete operating
range of the engine under all flight and atmospheric
conditions.
(b)
The intake passages of the engine through which air
or fuel in combination with air passes for
combustion purposes must be designed and constructed
to minimize the danger of ice accretion in those
passages. The engine must be designed and
constructed to permit the use of a means for ice
prevention.
(c)
The type and degree of fuel filtering necessary for
protection of the engine fuel system against foreign
particles in the fuel must be specified. The
applicant must show that foreign particles passing
through the prescribed filtering means will not
critically impair engine fuel system functioning.
(d)
Each passage in the induction system that conducts a
mixture of fuel and air must be self-draining, to
prevent a liquid lock in the cylinders, in all
attitudes that the applicant establishes as those
the engine can have when the aircraft in which it is
installed is in the static ground attitude.
(e)
If provided as part of the engine, the applicant
must show for each fluid injection (other than fuel)
system and its controls that the flow of the
injected fluid is adequately controlled.
Each
spark ignition engine must have a dual ignition
system with at least two spark plugs for each
cylinder and two separate electric circuits with
separate sources of electrical energy, or have an
ignition system of equivalent in-flight reliability.
33.39 Lubrication
system.
(a)
The lubrication system of the engine must be
designed and constructed so that it will function
properly in all flight attitudes and atmospheric
conditions in which the airplane is expected to
operate. In wet sump engines, this requirement must
be met when only one-half of the maximum lubricant
supply is in the engine.
(b)
The lubrication system of the engine must be
designed and constructed to allow installing a means
of cooling the lubricant.
(c)
The crankcase must be vented to the atmosphere to
preclude leakage of oil from excessive pressure in
the crankcase.
This
subpart prescribes the block tests and inspections
for reciprocating aircraft engines.
Before each endurance test required by this subpart,
the adjustment setting and functioning
characteristic of each component having an
adjustment setting and a functioning characteristic
that can be established independent of installation
on the engine must be established and recorded.
(a)
Each engine must undergo a vibration survey to
establish the torsional and bending vibration
characteristics of the crankshaft and the propeller
shaft or other output shaft, over the range of
crankshaft speed and engine power, under steady
state and transient conditions, from idling speed to
either 110 percent of the desired maximum continuous
speed rating or 103 percent of the maximum desired
take-off speed rating, whichever is higher. The
survey must be conducted using, for airplane
engines, the same configuration of the propeller
type which is used for the endurance test, and
using, for other engines, the same configuration of
the loading device type which is used for the
endurance test.
(b)
The torsional and bending vibration stresses of the
crankshaft and the propeller shaft or other output
shaft may not exceed the endurance limit stress of
the material from which the shaft is made. If the
maximum stress in the shaft cannot be shown to be
below the endurance limit by measurement, the
vibration frequency and amplitude must be measured.
The peak amplitude must be shown to produce a stress
below the endurance limit; if not, the engine must
be run at the condition producing the peak amplitude
until, for steel shafts, 10 million stress reversals
have been sustained without fatigue failure and, for
other shafts, until it is shown that fatigue will
not occur within the endurance limit stress of the
material.
(c)
Each accessory drive and mounting attachment must be
loaded, with the loads imposed by each accessory
used only for an aircraft service being the limit
load specified by the applicant for the drive or
attachment point.
(d)
The vibration survey described in paragraph (a) of
this section must be repeated with that cylinder not
firing which has the most adverse vibration effect,
in order to establish the conditions under which the
engine can be operated safely in that abnormal
state. However, for this vibration survey, the
engine speed range need only extend from idle to the
maximum desired take-off speed, and compliance with
paragraph (b) of this section need not be shown.
(a)
Each engine must be subjected to the calibration
tests necessary to establish its power
characteristics and the conditions for the endurance
test specified in 33.49. The results of the power
characteristics calibration tests form the basis for
establishing the characteristics of the engine over
its entire operating range of crankshaft rotational
speeds, manifold pressures, fuel/air mixture
settings, and altitudes. Power ratings are based
upon standard atmospheric conditions with only those
accessories installed which are essential for engine
functioning.
(b)
A power check at sea level conditions must be
accomplished on the endurance test engine after the
endurance test. Any change in power characteristics
which occurs during the endurance test must be
determined. Measurements taken during the final
portion of the endurance test may be used in showing
compliance with the requirements of this paragraph.
Each
engine must be tested to establish that the engine
can function without detonation throughout its range
of intended conditions of operation.
(a)
General. Each engine must be subjected to an
endurance test that includes a total of 150 hours of
operation (except as provided in paragraph
(e)(1)(iii) of this section) and, depending upon the
type and contemplated use of the engine, consists of
one of the series of runs specified in paragraphs
(b) through (e) of this section, as applicable. The
runs must be made in the order found appropriate by
the Administrator for the particular engine being
tested. During the endurance test the engine power
and the crankshaft rotational speed must be kept
within ±3 percent of the rated values. During the
runs at rated take-off power and for at least 35
hours at rated maximum continuous power, one
cylinder must be operated at not less than the
limiting temperature, the other cylinders must be
operated at a temperature not lower than 50 degrees
F. below the limiting temperature, and the oil inlet
temperature must be maintained within ±10 degrees F.
of the limiting temperature. An engine that is
equipped with a propeller shaft must be fitted for
the endurance test with a propeller that
thrust-loads the engine to the maximum thrust which
the engine is designed to resist at each applicable
operating condition specified in this section. Each
accessory drive and mounting attachment must be
loaded. During operation at rated take-off power and
rated maximum continuous power, the load imposed by
each accessory used only for an aircraft service
must be the limit load specified by the applicant
for the engine drive or attachment point.
(b)
Unsupercharged engines and engines incorporating
a gear-driven single-speed supercharger. For
engines not incorporating a supercharger and for
engines incorporating a gear-driven single-speed
supercharger the applicant must conduct the
following runs:
(1)
A 30-hour run consisting of alternate periods of 5
minutes at rated take-off power with take-off speed,
and 5 minutes at maximum best economy cruising power
or maximum recommended cruising power.
(2)
A 20-hour run consisting of alternate periods of
11/2hours at rated maximum continuous power with
maximum continuous speed, and1/2hour at 75 percent
rated maximum continuous power and 91 percent
maximum continuous speed.
(3)
A 20-hour run consisting of alternate periods of
11/2hours at rated maximum continuous power with
maximum continuous speed, and1/2hour at 70 percent
rated maximum continuous power and 89 percent
maximum continuous speed.
(4)
A 20-hour run consisting of alternate periods of
11/2hours at rated maximum continuous power with
maximum continuous speed, and1/2hour at 65 percent
rated maximum continuous power and 87 percent
maximum continuous speed.
(5)
A 20-hour run consisting of alternate periods of
11/2hours at rated maximum continuous power with
maximum continuous speed, and1/2hour at 60 percent
rated maximum continuous power and 84.5 percent
maximum continuous speed.
(6)
A 20-hour run consisting of alternate periods of
11/2hours at rated maximum continuous power with
maximum continuous speed, and1/2hour at 50 percent
rated maximum continuous power and 79.5 percent
maximum continuous speed.
(7)
A 20-hour run consisting of alternate periods of
21/2hours at rated maximum continuous power with
maximum continuous speed, and 21/2hours at maximum
best economy cruising power or at maximum
recommended cruising power.
(c)
Engines incorporating a gear-driven two-speed
supercharger. For engines incorporating a
gear-driven two-speed supercharger the applicant
must conduct the following runs:
(1)
A 30-hour run consisting of alternate periods in the
lower gear ratio of 5 minutes at rated take-off
power with take-off speed, and 5 minutes at maximum
best economy cruising power or at maximum
recommended cruising power. If a take-off power
rating is desired in the higher gear ratio, 15 hours
of the 30-hour run must be made in the higher gear
ratio in alternate periods of 5 minutes at the
observed horsepower obtainable with the take-off
critical altitude manifold pressure and take-off
speed, and 5 minutes at 70 percent high ratio rated
maximum continuous power and 89 percent high ratio
maximum continuous speed.
(2)
A 15-hour run consisting of alternate periods in the
lower gear ratio of 1 hour at rated maximum
continuous power with maximum continuous speed,
and1/2hour at 75 percent rated maximum continuous
power and 91 percent maximum continuous speed.
(3)
A 15-hour run consisting of alternate periods in the
lower gear ratio of 1 hour at rated maximum
continuous power with maximum continuous speed,
and1/2hour at 70 percent rated maximum continuous
power and 89 percent maximum continuous speed.
(4)
A 30-hour run in the higher gear ratio at rated
maximum continuous power with maximum continuous
speed.
(5)
A 5-hour run consisting of alternate periods of 5
minutes in each of the supercharger gear ratios. The
first 5 minutes of the test must be made at maximum
continuous speed in the higher gear ratio and the
observed horsepower obtainable with 90 percent of
maximum continuous manifold pressure in the higher
gear ratio under sea level conditions. The condition
for operation for the alternate 5 minutes in the
lower gear ratio must be that obtained by shifting
to the lower gear ratio at constant speed.
(6)
A 10-hour run consisting of alternate periods in the
lower gear ratio of 1 hour at rated maximum
continuous power with maximum continuous speed, and
1 hour at 65 percent rated maximum continuous power
and 87 percent maximum continuous speed.
(7)
A 10-hour run consisting of alternate periods in the
lower gear ratio of 1 hour at rated maximum
continuous power with maximum continuous speed, and
1 hour at 60 percent rated maximum continuous power
and 84.5 percent maximum continuous speed.
(8)
A 10-hour run consisting of alternate periods in the
lower gear ratio of 1 hour at rated maximum
continuous power with maximum continuous speed, and
1 hour at 50 percent rated maximum continuous power
and 79.5 percent maximum continuous speed.
(9)
A 20-hour run consisting of alternate periods in the
lower gear ratio of 2 hours at rated maximum
continuous power with maximum continuous speed, and
2 hours at maximum best economy cruising power and
speed or at maximum recommended cruising power.
(10)
A 5-hour run in the lower gear ratio at maximum best
economy cruising power and speed or at maximum
recommended cruising power and speed.
Where simulated altitude test equipment is not
available when operating in the higher gear ratio,
the runs may be made at the observed horsepower
obtained with the critical altitude manifold
pressure or specified percentages thereof, and the
fuel-air mixtures may be adjusted to be rich enough
to suppress detonation.
(d)
Helicopter engines. To be eligible for use on
a helicopter each engine must either comply with
paragraphs (a) through (j) of 29.923 of this
chapter, or must undergo the following series of
runs:
(1)
A 35-hour run consisting of alternate periods of 30
minutes each at rated take-off power with take-off
speed, and at rated maximum continuous power with
maximum continuous speed.
(2)
A 25-hour run consisting of alternate periods of
21/2hours each at rated maximum continuous power
with maximum continuous speed, and at 70 percent
rated maximum continuous power with maximum
continuous speed.
(3)
A 25-hour run consisting of alternate periods of
21/2hours each at rated maximum continuous power
with maximum continuous speed, and at 70 percent
rated maximum continuous power with 80 to 90 percent
maximum continuous speed.
(4)
A 25-hour run consisting of alternate periods of
21/2hours each at 30 percent rated maximum
continuous power with take-off speed, and at 30
percent rated maximum continuous power with 80 to 90
percent maximum continuous speed.
(5)
A 25-hour run consisting of alternate periods of
21/2hours each at 80 percent rated maximum
continuous power with take-off speed, and at either
rated maximum continuous power with 110 percent
maximum continuous speed or at rated take-off power
with 103 percent take-off speed, whichever results
in the greater speed.
(6)
A 15-hour run at 105 percent rated maximum
continuous power with 105 percent maximum continuous
speed or at full throttle and corresponding speed at
standard sea level carburetor entrance pressure, if
105 percent of the rated maximum continuous power is
not exceeded.
(e)
Turbo-supercharged engines. For engines
incorporating a turbo-supercharger the following
apply except that altitude testing may be simulated
provided the applicant shows that the engine and
supercharger are being subjected to mechanical loads
and operating temperatures no less severe than if
run at actual altitude conditions:
(1)
For engines used in airplanes the applicant must
conduct the runs specified in paragraph (b) of this
section, except—
(i)
The entire run specified in paragraph (b)(1) of this
section must be made at sea level altitude pressure;
(ii)
The portions of the runs specified in paragraphs
(b)(2) through (7) of this section at rated maximum
continuous power must be made at critical altitude
pressure, and the portions of the runs at other
power must be made at 8,000 feet altitude pressure;
and
(iii) The turbo-supercharger used during the
150-hour endurance test must be run on the bench for
an additional 50 hours at the limiting turbine wheel
inlet gas temperature and rotational speed for rated
maximum continuous power operation unless the
limiting temperature and speed are maintained during
50 hours of the rated maximum continuous power
operation.
(2)
For engines used in helicopters the applicant must
conduct the runs specified in paragraph (d) of this
section, except—
(i)
The entire run specified in paragraph (d)(1) of this
section must be made at critical altitude pressure;
(ii)
The portions of the runs specified in paragraphs
(d)(2) and (3) of this section at rated maximum
continuous power must be made at critical altitude
pressure and the portions of the runs at other power
must be made at 8,000 feet altitude pressure;
(iii) The entire run specified in paragraph (d)(4)
of this section must be made at 8,000 feet altitude
pressure;
(iv)
The portion of the runs specified in paragraph
(d)(5) of this section at 80 percent of rated
maximum continuous power must be made at 8,000 feet
altitude pressure and the portions of the runs at
other power must be made at critical altitude
pressure;
(v)
The entire run specified in paragraph (d)(6) of this
section must be made at critical altitude pressure;
and
(vi)
The turbo-supercharger used during the endurance
test must be run on the bench for 50 hours at the
limiting turbine wheel inlet gas temperature and
rotational speed for rated maximum continuous power
operation unless the limiting temperature and speed
are maintained during 50 hours of the rated maximum
continuous power operation.
The
operation test must include the testing found
necessary by the Administrator to demonstrate
backfire characteristics, starting, idling,
acceleration, over-speeding, functioning of
propeller and ignition, and any other operational
characteristic of the engine. If the engine
incorporates a multi-speed supercharger drive, the
design and construction must allow the supercharger
to be shifted from operation at the lower speed
ratio to the higher and the power appropriate to the
manifold pressure and speed settings for rated
maximum continuous power at the higher supercharger
speed ratio must be obtainable within five seconds.
(a)
For each engine that cannot be adequately
substantiated by endurance testing in accordance
with 33.49, the applicant must conduct additional
tests to establish that components are able to
function reliably in all normally anticipated flight
and atmospheric conditions.
(b)
Temperature limits must be established for each
component that requires temperature controlling
provisions in the aircraft installation to assure
satisfactory functioning, reliability, and
durability.
After completing the endurance test—
(a)
Each engine must be completely disassembled;
(b)
Each component having an adjustment setting and a
functioning characteristic that can be established
independent of installation on the engine must
retain each setting and functioning characteristic
within the limits that were established and recorded
at the beginning of the test; and
(c)
Each engine component must conform to the type
design and be eligible for incorporation into an
engine for continued operation, in accordance with
information submitted in compliance with 33.4.
(a)
The applicant may, in conducting the block tests,
use separate engines of identical design and
construction in the vibration, calibration,
detonation, endurance, and operation tests, except
that, if a separate engine is used for the endurance
test it must be subjected to a calibration check
before starting the endurance test.
(b)
The applicant may service and make minor repairs to
the engine during the block tests in accordance with
the service and maintenance instructions submitted
in compliance with 33.4. If the frequency of the
service is excessive, or the number of stops due to
engine malfunction is excessive, or a major repair,
or replacement of a part is found necessary during
the block tests or as the result of findings from
the teardown inspection, the engine or its parts may
be subjected to any additional test the
Administrator finds necessary.
(c)
Each applicant must furnish all testing facilities,
including equipment and competent personnel, to
conduct the block tests.
This
subpart prescribes additional design and
construction requirements for turbine aircraft
engines.
A stress analysis must be performed on each turbine engine
showing the design safety margin of each turbine
engine rotor, spacer, and rotor shaft.
Each
engine must be designed and constructed to function
throughout its declared flight envelope and
operating range of rotational speeds and
power/thrust, without inducing excessive stress in
any engine part because of vibration and without
imparting excessive vibration forces to the aircraft
structure.
When
the engine is operated in accordance with operating
instructions required by 33.5(b), starting, a change
of power or thrust, power or thrust augmentation,
limiting inlet air distortion, or inlet air
temperature may not cause surge or stall to the
extent that flameout, structural failure,
over-temperature, or failure of the engine to
recover power or thrust will occur at any point in
the operating envelope.
The
engine must supply bleed air without adverse effect
on the engine, excluding reduced thrust or power
output, at all conditions up to the discharge flow
conditions established as a limitation under
33.7(c)(11). If bleed air used for engine anti-icing
can be controlled, provision must be made for a
means to indicate the functioning of the engine ice
protection system.
(a)
With fuel supplied to the engine at the flow and
pressure specified by the applicant, the engine must
function properly under each operating condition
required by this part. Each fuel control adjusting
means that may not be manipulated while the fuel
control device is mounted on the engine must be
secured by a locking device and sealed, or otherwise
be inaccessible. All other fuel control adjusting
means must be accessible and marked to indicate the
function of the adjustment unless the function is
obvious.
(b)
There must be a fuel strainer or filter between the
engine fuel inlet opening and the inlet of either
the fuel metering device or the engine-driven
positive displacement pump whichever is nearer the
engine fuel inlet. In addition, the following
provisions apply to each strainer or filter required
by this paragraph (b):
(1)
It must be accessible for draining and cleaning and
must incorporate a screen or element that is easily
removable.
(2)
It must have a sediment trap and drain except that
it need not have a drain if the strainer or filter
is easily removable for drain purposes.
(3)
It must be mounted so that its weight is not
supported by the connecting lines or by the inlet or
outlet connections of the strainer or filter, unless
adequate strength margins under all loading
conditions are provided in the lines and
connections.
(4)
It must have the type and degree of fuel filtering
specified as necessary for protection of the engine
fuel system against foreign particles in the fuel.
The applicant must show:
(i)
That foreign particles passing through the specified
filtering means do not impair the engine fuel system
functioning; and
(ii)
That the fuel system is capable of sustained
operation throughout its flow and pressure range
with the fuel initially saturated with water at 80
°F (27 °C) and having 0.025 fluid ounces per gallon
(0.20 milliliters per liter) of free water added and
cooled to the most critical condition for icing
likely to be encountered in operation. However, this
requirement may be met by demonstrating the
effectiveness of specified approved fuel anti-icing
additives, or that the fuel system incorporates a
fuel heater which maintains the fuel temperature at
the fuel strainer or fuel inlet above 32 °F (0 °C)
under the most critical conditions.
(5)
The applicant must demonstrate that the filtering
means has the capacity (with respect to engine
operating limitations) to ensure that the engine
will continue to operate within approved limits,
with fuel contaminated to the maximum degree of
particle size and density likely to be encountered
in service. Operation under these conditions must be
demonstrated for a period acceptable to the
Administrator, beginning when indication of
impending filter blockage is first given by either:
(i)
Existing engine instrumentation; or
(ii)
Additional means incorporated into the engine fuel
system.
(6)
Any strainer or filter bypass must be designed and
constructed so that the release of collected
contaminants is minimized by appropriate location of
the bypass to ensure that collected contaminants are
not in the bypass flow path.
(c)
If provided as part of the engine, the applicant
must show for each fluid injection (other than fuel)
system and its controls that the flow of the
injected fluid is adequately controlled.
(d)
Engines having a 30-second OEI rating must
incorporate means for automatic availability and
automatic control of a 30-second OEI power.
Each
engine, with all icing protection systems operating,
must—
(a)
Operate throughout its flight power range (including
idling) without the accumulation of ice on the
engine components that adversely affects engine
operation or that causes a serious loss of power or
thrust in continuous maximum and intermittent
maximum icing conditions as defined in appendix C of
Part 25 of this chapter; and
(b)
Idle for 30 minutes on the ground, with the
available air bleed for icing protection at its
critical condition, without adverse effect, in an
atmosphere that is at a temperature between 15° and
30 °F (between −9° and −1 °C) and has a liquid water
content not less than 0.3 grams per cubic meter in
the form of drops having a mean effective diameter
not less than 20 microns, followed by a momentary
operation at take-off power or thrust. During the 30
minutes of idle operation the engine may be run up
periodically to a moderate power or thrust setting
in a manner acceptable to the Administrator.
Each
engine must be equipped with an ignition system for
starting the engine on the ground and in flight. An
electric ignition system must have at least two
igniters and two separate secondary electric
circuits, except that only one igniter is required
for fuel burning augmentation systems.
33.70 Engine
life-limited parts.
By a
procedure approved by the AFRO-CAA, operating
limitations must be established which specify the
maximum allowable number of flight cycles for each
engine life-limited part. Engine life-limited parts
are rotor and major static structural parts whose
primary failure is likely to result in a hazardous
engine effect. Typically, engine life-limited parts
include, but are not limited to disks, spacers,
hubs, shafts, high-pressure casings, and
non-redundant mount components. For the purposes of
this section, a hazardous engine effect is any of
the conditions listed in 33.75 of this part. The
applicant will establish the integrity of each
engine life-limited part by:
(a)
An engineering plan that contains the steps required
to ensure each engine life-limited part is withdrawn
from service at an approved life before hazardous
engine effects can occur. These steps include
validated analysis, test, or service experience
which ensures that the combination of loads,
material properties, environmental influences and
operating conditions, including the effects of other
engine parts influencing these parameters, are
sufficiently well known and predictable so that the
operating limitations can be established and
maintained for each engine life-limited part.
Applicants must perform appropriate damage tolerance
assessments to address the potential for failure
from material, manufacturing, and service induced
anomalies within the approved life of the part.
Applicants must publish a list of the life-limited
engine parts and the approved life for each part in
the Airworthiness Limitations Section of the
Instructions for Continued Airworthiness as required
by 33.4 of this part.
(b)
A manufacturing plan that identifies the specific
manufacturing constraints necessary to consistently
produce each engine life-limited part with the
attributes required by the engineering plan.
(c)
A service management plan that defines in-service
processes for maintenance and the limitations to
repair for each engine life-limited part that will
maintain attributes consistent with those required
by the engineering plan. These processes and
limitations will become part of the Instructions for
Continued Airworthiness.
(a)
General. Each lubrication system must
function properly in the flight attitudes and
atmospheric conditions in which an aircraft is
expected to operate.
(b)
Oil strainer or filter. There must be an oil
strainer or filter through which all of the engine
oil flows. In addition:
(1)
Each strainer or filter required by this paragraph
that has a bypass must be constructed and installed
so that oil will flow at the normal rate through the
rest of the system with the strainer or filter
element completely blocked.
(2)
The type and degree of filtering necessary for
protection of the engine oil system against foreign
particles in the oil must be specified. The
applicant must demonstrate that foreign particles
passing through the specified filtering means do not
impair engine oil system functioning.
(3)
Each strainer or filter required by this paragraph
must have the capacity (with respect to operating
limitations established for the engine) to ensure
that engine oil system functioning is not impaired
with the oil contaminated to a degree (with respect
to particle size and density) that is greater than
that established for the engine in paragraph (b)(2)
of this section.
(4)
For each strainer or filter required by this
paragraph, except the strainer or filter at the oil
tank outlet, there must be means to indicate
contamination before it reaches the capacity
established in accordance with paragraph (b)(3) of
this section.
(5)
Any filter bypass must be designed and constructed
so that the release of collected contaminants is
minimized by appropriate location of the bypass to
ensure that the collected contaminants are not in
the bypass flow path.
(6)
Each strainer or filter required by this paragraph
that has no bypass, except the strainer or filter at
an oil tank outlet or for a scavenge pump, must have
provisions for connection with a warning means to
warn the pilot of the occurence of contamination of
the screen before it reaches the capacity
established in accordance with paragraph (b)(3) of
this section.
(7)
Each strainer or filter required by this paragraph
must be accessible for draining and cleaning.
(c)
Oil tanks. (1) Each oil tank must have an
expansion space of not less than 10 percent of the
tank capacity.
(2)
It must be impossible to inadvertently fill the oil
tank expansion space.
(3)
Each recessed oil tank filler connection that can
retain any appreciable quantity of oil must have
provision for fitting a drain.
(4)
Each oil tank cap must provide an oil-tight seal.
For an applicant seeking eligibility for an engine
to be installed on an airplane approved for ETOPS,
the oil tank must be designed to prevent a hazardous
loss of oil due to an incorrectly installed oil tank
cap.
(5)
Each oil tank filler must be marked with the word
“oil.”
(6)
Each oil tank must be vented from the top part of
the expansion space, with the vent so arranged that
condensed water vapor that might freeze and obstruct
the line cannot accumulate at any point.
(7)
There must be means to prevent entrance into the oil
tank or into any oil tank outlet, of any object that
might obstruct the flow of oil through the system.
(8)
There must be a shutoff valve at the outlet of each
oil tank, unless the external portion of the oil
system (including oil tank supports) is fireproof.
(9)
Each unpressurized oil tank may not leak when
subjected to a maximum operating temperature and an
internal pressure of 5 p.s.i., and each pressurized
oil tank may not leak when subjected to maximum
operating temperature and an internal pressure that
is not less than 5 p.s.i. plus the maximum operating
pressure of the tank.
(10)
Leaked or spilled oil may not accumulate between the
tank and the remainder of the engine.
(11)
Each oil tank must have an oil quantity indicator or
provisions for one.
(12)
If the propeller feathering system depends on engine
oil—
(i)
There must be means to trap an amount of oil in the
tank if the supply becomes depleted due to failure
of any part of the lubricating system other than the
tank itself;
(ii)
The amount of trapped oil must be enough to
accomplish the feathering operation and must be
available only to the feathering pump; and
(iii) Provision must be made to prevent sludge or
other foreign matter from affecting the safe
operation of the propeller feathering system.
(d)
Oil drains. A drain (or drains) must be
provided to allow safe drainage of the oil system.
Each drain must—
(1)
Be accessible; and
(2)
Have manual or automatic means for positive locking
in the closed position.
(e)
Oil radiators. Each oil radiator must
withstand, without failure, any vibration, inertia,
and oil pressure load to which it is subjected
during the block tests.
Each hydraulic actuating system must function properly under
all conditions in which the engine is expected to
operate. Each filter or screen must be accessible
for servicing and each tank must meet the design
criteria of 33.71.
The
design and construction of the engine must enable an
increase—
(a)
From minimum to rated take-off power or thrust with
the maximum bleed air and power extraction to be
permitted in an aircraft, without over temperature,
surge, stall, or other detrimental factors occurring
to the engine whenever the power control lever is
moved from the minimum to the maximum position in
not more than 1 second, except that the
Administrator may allow additional time increments
for different regimes of control operation requiring
control scheduling; and
(b)
From the fixed minimum flight idle power lever
position when provided, or if not provided, from not
more than 15 percent of the rated take-off power or
thrust available to 95 percent rated take-off power
or thrust in not over 5 seconds. The 5-second power
or thrust response must occur from a stabilized
static condition using only the bleed air and
accessories loads necessary to run the engine. This
take-off rating is specified by the applicant and
need not include thrust augmentation.
If
any of the engine main rotating systems continue to
rotate after the engine is shutdown for any reason
while in flight, and if means to prevent that
continued rotation are not provided, then any
continued rotation during the maximum period of
flight, and in the flight conditions expected to
occur with that engine inoperative, may not result
in any condition described in 33.75(g)(2)(i) through
(vi) of this part.
(a)
(1) The applicant must analyze the engine, including
the control system, to assess the likely
consequences of all failures that can reasonably be
expected to occur. This analysis will take into
account, if applicable:
(i)
Aircraft-level devices and procedures assumed to be
associated with a typical installation. Such
assumptions must be stated in the analysis.
(ii)
Consequential secondary failures and latent
failures.
(iii) Multiple failures referred to in paragraph (d)
of this section or that result in the hazardous
engine effects defined in paragraph (g)(2) of this
section.
(2)
The applicant must summarize those failures that
could result in major engine effects or hazardous
engine effects, as defined in paragraph (g) of this
section, and estimate the probability of occurrence
of those effects. Any engine part the failure of
which could reasonably result in a hazardous engine
effect must be clearly identified in this summary.
(3)
The applicant must show that hazardous engine
effects are predicted to occur at a rate not in
excess of that defined as extremely remote
(probability range of 10−7to 10−9per
engine flight hour). Since the estimated probability
for individual failures may be insufficiently
precise to enable the applicant to assess the total
rate for hazardous engine effects, compliance may be
shown by demonstrating that the probability of a
hazardous engine effect arising from an individual
failure can be predicted to be not greater than 10−8per
engine flight hour. In dealing with probabilities of
this low order of magnitude, absolute proof is not
possible, and compliance may be shown by reliance on
engineering judgment and previous experience
combined with sound design and test philosophies.
(4)
The applicant must show that major engine effects
are predicted to occur at a rate not in excess of
that defined as remote (probability range of 10−5to
10−7per engine flight hour).
(b)
The AFRO-CAA may require that any assumption as to
the effects of failures and likely combination of
failures be verified by test.
(c)
The primary failure of certain single elements
cannot be sensibly estimated in numerical terms. If
the failure of such elements is likely to result in
hazardous engine effects, then compliance may be
shown by reliance on the prescribed integrity
requirements of 33.15, 33.27, and 33.70 as
applicable. These instances must be stated in the
safety analysis.
(d)
If reliance is placed on a safety system to prevent
a failure from progressing to hazardous engine
effects, the possibility of a safety system failure
in combination with a basic engine failure must be
included in the analysis. Such a safety system may
include safety devices, instrumentation, early
warning devices, maintenance checks, and other
similar equipment or procedures. If items of a
safety system are outside the control of the engine
manufacturer, the assumptions of the safety analysis
with respect to the reliability of these parts must
be clearly stated in the analysis and identified in
the installation instructions under §33.5 of this
part.
(e)
If the safety analysis depends on one or more of the
following items, those items must be identified in
the analysis and appropriately substantiated.
(1)
Maintenance actions being carried out at stated
intervals. This includes the verification of the
serviceability of items that could fail in a latent
manner. When necessary to prevent hazardous engine
effects, these maintenance actions and intervals
must be published in the instructions for continued
airworthiness required under §33.4 of this part.
Additionally, if errors in maintenance of the
engine, including the control system, could lead to
hazardous engine effects, the appropriate procedures
must be included in the relevant engine manuals.
(2)
Verification of the satisfactory functioning of
safety or other devices at pre-flight or other
stated periods. The details of this satisfactory
functioning must be published in the appropriate
manual.
(3)
The provisions of specific instrumentation not
otherwise required.
(4)
Flight crew actions to be specified in the operating
instructions established under 33.5.
(f)
If applicable, the safety analysis must also
include, but not be limited to, investigation of the
following:
(1)
Indicating equipment;
(2)
Manual and automatic controls;
(3)
Compressor bleed systems;
(4)
Refrigerant injection systems;
(5)
Gas temperature control systems;
(6)
Engine speed, power, or thrust governors and fuel
control systems;
(7)
Engine overspeed, overtemperature, or topping
limiters;
(8)
Propeller control systems; and
(9)
Engine or propeller thrust reversal systems.
(g)
Unless otherwise approved by the AFRO-CAA and stated
in the safety analysis, for compliance with part 33,
the following failure definitions apply to the
engine:
(1)
An engine failure in which the only consequence is
partial or complete loss of thrust or power (and
associated engine services) from the engine will be
regarded as a minor engine effect.
(2)
The following effects will be regarded as hazardous
engine effects:
(i)
Non-containment of high-energy debris;
(ii)
Concentration of toxic products in the engine bleed
air intended for the cabin sufficient to
incapacitate crew or passengers;
(iii) Significant thrust in the opposite direction
to that commanded by the pilot;
(iv)
Uncontrolled fire;
(v)
Failure of the engine mount system leading to
inadvertent engine separation;
(vi)
Release of the propeller by the engine, if
applicable; and
(vii) Complete inability to shut the engine down.
(3)
An effect whose severity falls between those effects
covered in paragraphs (g)(1) and (g)(2) of this
section will be regarded as a major engine effect.
(a)
General. Compliance with paragraphs (b), (c),
and (d) of this section shall be in accordance with
the following:
(1)
Except as specified in paragraph (d) of this
section, all ingestion tests must be conducted with
the engine stabilized at no less than 100-percent
take-off power or thrust, for test day ambient
conditions prior to the ingestion. In addition, the
demonstration of compliance must account for engine
operation at sea level take-off conditions on the
hottest day that a minimum engine can achieve
maximum rated take-off thrust or power.
(2)
The engine inlet throat area as used in this section
to determine the bird quantity and weights will be
established by the applicant and identified as a
limitation in the installation instructions required
under 33.5.
(3)
The impact to the front of the engine from the large
single bird, the single largest medium bird which
can enter the inlet, and the large flocking bird
must be evaluated. Applicants must show that the
associated components when struck under the
conditions prescribed in paragraphs (b), (c) or (d)
of this section, as applicable, will not affect the
engine to the extent that the engine cannot comply
with the requirements of paragraphs (b)(3), (c)(6)
and (d)(4) of this section.
(4)
For an engine that incorporates an inlet protection
device, compliance with this section shall be
established with the device functioning. The engine
approval will be endorsed to show that compliance
with the requirements has been established with the
device functioning.
(5)
Objects that are accepted by the Administrator may
be substituted for birds when conducting the bird
ingestion tests required by paragraphs (b), (c) and
(d) of this section.
(6)
If compliance with the requirements of this section
is not established, the engine type certification
documentation will show that the engine shall be
limited to aircraft installations in which it is
shown that a bird cannot strike the engine, or be
ingested into the engine, or adversely restrict
airflow into the engine.
(b)
Large single bird. Compliance with the large
bird ingestion requirements shall be in accordance
with the following:
(1)
The large bird ingestion test shall be conducted
using one bird of a weight determined from Table 1
aimed at the most critical exposed location on the
first stage rotor blades and ingested at a bird
speed of 200-knots for engines to be installed on
airplanes, or the maximum airspeed for normal
rotorcraft flight operations for engines to be
installed on rotorcraft.
(2)
Power lever movement is not permitted within 15
seconds following ingestion of the large bird.
(3)
Ingestion of a single large bird tested under the
conditions prescribed in this section may not result
in any condition described in 33.75(g)(2) of this
part.
(4)
Compliance with the large bird ingestion
requirements of this paragraph may be shown by
demonstrating that the requirements of 33.94(a)
constitute a more severe demonstration of blade
containment and rotor unbalance than the
requirements of this paragraph.
Table 1 to 33.76—Large Bird Weight Requirements
|
Engine Inlet Throat Area (A)—Square-meters
(square-inches) |
Bird weight kg. (lb.) |
|
1.35 (2,092)> A |
1.85 (4.07) minimum, unless a smaller bird is
determined to be a more severe demonstration. |
|
1.35 (2,092)≤ A< 3.90 (6,045) |
2.75 (6.05) |
|
3.90 (6,045)≤ A |
3.65 (8.03) |
(c)
Small and medium flocking bird . Compliance
with the small and medium bird ingestion
requirements shall be in accordance with the
following:
(1)
Analysis or component test, or both, acceptable to
the Administrator, shall be conducted to determine
the critical ingestion parameters affecting power
loss and damage. Critical ingestion parameters shall
include, but are not limited to, the effects of bird
speed, critical target location, and first stage
rotor speed. The critical bird ingestion speed
should reflect the most critical condition within
the range of airspeeds used for normal flight
operations up to 1,500 feet above ground level, but
not less than V1minimum for airplanes.
(2)
Medium bird engine tests shall be conducted so as to
simulate a flock encounter, and will use the bird
weights and quantities specified in Table 2. When
only one bird is specified, that bird will be aimed
at the engine core primary flow path; the other
critical locations on the engine face area must be
addressed, as necessary, by appropriate tests or
analysis, or both. When two or more birds are
specified in Table 2, the largest of those birds
must be aimed at the engine core primary flow path,
and a second bird must be aimed at the most critical
exposed location on the first stage rotor blades.
Any remaining birds must be evenly distributed over
the engine face area.
(3)
In addition, except for rotorcraft engines, it must
also be substantiated by appropriate tests or
analysis or both, that when the full fan assembly is
subjected to the ingestion of the quantity and
weights of bird from Table 3, aimed at the fan
assembly's most critical location outboard of the
primary core flow-path, and in accordance with the
applicable test conditions of this paragraph, that
the engine can comply with the acceptance criteria
of this paragraph.
(4)
A small bird ingestion test is not required if the
prescribed number of medium birds pass into the
engine rotor blades during the medium bird test.
(5)
Small bird ingestion tests shall be conducted so as
to simulate a flock encounter using one 85 gram
(0.187 lb.) bird for each 0.032 square-meter (49.6
square-inches) of inlet area, or fraction thereof,
up to a maximum of 16 birds. The birds will be aimed
so as to account for any critical exposed locations
on the first stage rotor blades, with any remaining
birds evenly distributed over the engine face area.
(6)
Ingestion of small and medium birds tested under the
conditions prescribed in this paragraph may not
cause any of the following:
(i)
More than a sustained 25-percent power or thrust
loss;
(ii)
The engine to be shut down during the required
run-on demonstration prescribed in paragraphs (c)(7)
or (c)(8) of this section;
(iii) The conditions defined in paragraph (b)(3) of
this section.
(iv)
Unacceptable deterioration of engine handling
characteristics.
(7)
Except for rotorcraft engines, the following test
schedule shall be used:
(i)
Ingestion so as to simulate a flock encounter, with
approximately 1 second elapsed time from the moment
of the first bird ingestion to the last.
(ii)
Followed by 2 minutes without power lever movement
after the ingestion.
(iii) Followed by 3 minutes at 75-percent of the
test condition.
(iv)
Followed by 6 minutes at 60-percent of the test
condition.
(v)
Followed by 6 minutes at 40-percent of the test
condition.
(vi)
Followed by 1 minute at approach idle.
(vii) Followed by 2 minutes at 75-percent of the
test condition.
(viii) Followed by stabilizing at idle and engine
shut down.
(ix)
The durations specified are times at the defined
conditions with the power being changed between each
condition in less than 10 seconds.
(8)
For rotorcraft engines, the following test schedule
shall be used:
(i)
Ingestion so as to simulate a flock encounter within
approximately 1 second elapsed time between the
first ingestion and the last.
(ii)
Followed by 3 minutes at 75-percent of the test
condition.
(iii) Followed by 90 seconds at descent flight idle.
(iv)
Followed by 30 seconds at 75-percent of the test
condition.
(v)
Followed by stabilizing at idle and engine shut
down.
(vi)
The durations specified are times at the defined
conditions with the power being changed between each
condition in less than 10 seconds.
(9)
Engines intended for use in multi-engine rotorcraft
are not required to comply with the medium bird
ingestion portion of this section, providing that
the appropriate type certificate documentation is so
endorsed.
(10)
If any engine operating limit(s) is exceeded during
the initial 2 minutes without power lever movement,
as provided by paragraph (c)(7)(ii) of this section,
then it shall be established that the limit
exceedence will not result in an unsafe condition.
Table 2 to 33.76—Medium Flocking Bird Weight and Quantity
Requirements
|
Engine Inlet Throat Area (A)—
Square-meters (square-inches) |
Bird quantity |
Bird weight kg. (lb.) |
|
0.05 (77.5)> A |
none |
|
|
0.05 (77.5)≤ A <0.10 (155) |
1 |
0.35 (0.77) |
|
0.10 (155)≤ A <0.20 (310) |
1 |
0.45 (0.99) |
|
0.20 (310)≤ A <0.40 (620) |
2 |
0.45 (0.99) |
|
0.40 (620)≤ A <0.60 (930) |
2 |
0.70 (1.54) |
|
0.60 (930)≤ A <1.00 (1,550) |
3 |
0.70 (1.54) |
|
1.00 (1,550)≤ A <1.35 (2,092) |
4 |
0.70 (1.54) |
|
1.35 (2,092)≤ A <1.70 (2,635) |
1 |
1.15 (2.53) |
|
|
plus 3 |
0.70 (1.54) |
|
1.70 (2,635)≤ A <2.10 (3,255) |
1 |
1.15 (2.53) |
|
|
plus 4 |
0.70 (1.54) |
|
2.10 (3,255)≤ A <2.50 (3,875) |
1 |
1.15 (2.53) |
|
|
plus 5 |
0.70 (1.54) |
|
2.50 (3,875)≤ A <3.90 (6045) |
1 |
1.15 (2.53) |
|
|
plus 6 |
0.70 (1.54) |
|
3.90 (6045)≤ A <4.50 (6975) |
3 |
1.15 (2.53) |
|
4.50 (6975)≤ A |
4 |
1.15 (2.53) |
Table 3 to 33.76—Additional Integrity Assessment
|
Engine Inlet Throat Area (A)—
square-meters (square-inches) |
Bird quantity |
Bird weight kg. (lb.) |
|
1.35 (2,092)> A |
none |
|
|
1.35 (2,092)≤ A <2.90 (4,495) |
1 |
1.15 (2.53) |
|
2.90 (4,495)≤ A <3.90 (6,045) |
2 |
1.15 (2.53) |
|
3.90 (6,045)≤ A |
1 |
1.15 (2.53) |
|
|
plus 6 |
0.70 (1.54) |
(d)
Large flocking bird . An engine test will be
performed as follows:
(1)
Large flocking bird engine tests will be performed
using the bird mass and weights in Table 4, and
ingested at a bird speed of 200 knots.
(2)
Prior to the ingestion, the engine must be
stabilized at no less than the mechanical rotor
speed of the first exposed stage or stages that, on
a standard day, would produce 90 percent of the sea
level static maximum rated take-off power or thrust.
(3)
The bird must be targeted on the first exposed
rotating stage or stages at a blade airfoil height
of not less than 50 percent measured at the leading
edge.
(4)
Ingestion of a large flocking bird under the
conditions prescribed in this paragraph must not
cause any of the following:
(i)
A sustained reduction of power or thrust to less
than 50 percent of maximum rated take-off power or
thrust during the run-on segment specified under
paragraph (d)(5)(i) of this section.
(ii)
Engine shutdown during the required run-on
demonstration specified in paragraph (d)(5) of this
section.
(iii) The conditions specified in paragraph (b)(3)
of this section.
(5)
The following test schedule must be used:
(i)
Ingestion followed by 1 minute without power lever
movement.
(ii)
Followed by 13 minutes at not less than 50 percent
of maximum rated take-off power or thrust.
(iii) Followed by 2 minutes between 30 and 35
percent of maximum rated take-off power or thrust.
(iv)
Followed by 1 minute with power or thrust increased
from that set in paragraph (d)(5)(iii) of this
section, by between 5 and 10 percent of maximum
rated take-off power or thrust.
(v)
Followed by 2 minutes with power or thrust reduced
from that set in paragraph (d)(5)(iv) of this
section, by between 5 and 10 percent of maximum
rated take-off power or thrust.
(vi)
Followed by a minimum of 1 minute at ground idle
then engine shutdown. The durations specified are
times at the defined conditions. Power lever
movement between each condition will be 10 seconds
or less, except that power lever movements allowed
within paragraph (d)(5)(ii) of this section are not
limited, and for setting power under paragraph
(d)(5)(iii) of this section will be 30 seconds or
less.
(6)
Compliance with the large flocking bird ingestion
requirements of this paragraph (d) may also be
demonstrated by:
(i)
Incorporating the requirements of paragraph (d)(4)
and (d)(5) of this section, into the large single
bird test demonstration specified in paragraph
(b)(1) of this section; or
(ii)
Use of an engine subassembly test at the ingestion
conditions specified in paragraph (b)(1) of this
section if:
(A)
All components critical to complying with the
requirements of paragraph (d) of this section are
included in the subassembly test;
(B)
The components of paragraph (d)(6)(ii)(A) of this
section are installed in a representative engine for
a run-on demonstration in accordance with paragraphs
(d)(4) and (d)(5) of this section; except that
section (d)(5)(i) is deleted and section (d)(5)(ii)
must be 14 minutes in duration after the engine is
started and stabilized; and
(C)
The dynamic effects that would have been experienced
during a full engine ingestion test can be shown to
be negligible with respect to meeting the
requirements of paragraphs (d)(4) and (d)(5) of this
section.
(7)
Applicants must show that an unsafe condition will
not result if any engine operating limit is exceeded
during the run-on period.
Table 4 to 33.76—Large Flocking Bird Mass and Weight
|
Engine inlet throat area
(square meters/square inches) |
Bird quantity |
Bird mass and weight
(kg (lbs)) |
|
A < 2.50 (3875) |
none |
|
|
2.50 (3875) ≤ A < 3.50 (5425) |
1 |
1.85 (4.08) |
|
3.50 (5425) ≤ A < 3.90 (6045) |
1 |
2.10 (4.63) |
|
3.90 (6045) ≤ A |
1 |
2.50 (5.51) |
(a)–(b) [Reserved]
(c)
Ingestion of ice under the conditions of paragraph
(e) of this section may not—
(1)
Cause a sustained power or thrust loss; or
(2)
Require the engine to be shut down.
(d)
For an engine that incorporates a protection device,
compliance with this section need not be
demonstrated with respect to foreign objects to be
ingested under the conditions prescribed in
paragraph (e) of this section if it is shown that—
(1)
Such foreign objects are of a size that will not
pass through the protective device;
(2)
The protective device will withstand the impact of
the foreign objects; and
(3)
The foreign object, or objects, stopped by the
protective device will not obstruct the flow of
induction air into the engine with a resultant
sustained reduction in power or thrust greater than
those values required by paragraph (c) of this
section.
(e)
Compliance with paragraph (c) of this section must
be shown by engine test under the following
ingestion conditions:
(1)
Ice quantity will be the maximum accumulation on a
typical inlet cowl and engine face resulting from a
2-minute delay in actuating the anti-icing system;
or a slab of ice which is comparable in weight or
thickness for that size engine.
(2)
The ingestion velocity will simulate ice being
sucked into the engine inlet.
(3)
Engine operation will be maximum cruise power or
thrust.
(4)
The ingestion will simulate a continuous maximum
icing encounter at 25 degrees Fahrenheit.
(a)
All engines. (1) The ingestion of large
hailstones (0.8 to 0.9 specific gravity) at the
maximum true air speed, up to 15,000 feet (4,500
meters), associated with a representative aircraft
operating in rough air, with the engine at maximum
continuous power, may not cause unacceptable
mechanical damage or unacceptable power or thrust
loss after the ingestion, or require the engine to
be shut down. One-half the number of hailstones
shall be aimed randomly over the inlet face area and
the other half aimed at the critical inlet face
area. The hailstones shall be ingested in a rapid
sequence to simulate a hailstone encounter and the
number and size of the hailstones shall be
determined as follows:
(i)
One 1-inch (25 millimeters) diameter hailstone for
engines with inlet areas of not more than 100 square
inches (0.0645 square meters).
(ii)
One 1-inch (25 millimeters) diameter and one 2-inch
(50 millimeters) diameter hailstone for each 150
square inches (0.0968 square meters) of inlet area,
or fraction thereof, for engines with inlet areas of
more than 100 square inches (0.0645 square meters).
(2)
In addition to complying with paragraph (a)(1) of
this section and except as provided in paragraph (b)
of this section, it must be shown that each engine
is capable of acceptable operation throughout its
specified operating envelope when subjected to
sudden encounters with the certification standard
concentrations of rain and hail, as defined in
appendix B to this part. Acceptable engine operation
precludes flameout, run down, continued or
non-recoverable surge or stall, or loss of
acceleration and deceleration capability, during any
three minute continuous period in rain and during
any 30 second continuous period in hail. It must
also be shown after the ingestion that there is no
unacceptable mechanical damage, unacceptable power
or thrust loss, or other adverse engine anomalies.
(b)
Engines for rotorcraft. As an alternative to
the requirements specified in paragraph (a)(2) of
this section, for rotorcraft turbine engines only,
it must be shown that each engine is capable of
acceptable operation during and after the ingestion
of rain with an overall ratio of water droplet flow
to airflow, by weight, with a uniform distribution
at the inlet plane, of at least four percent.
Acceptable engine operation precludes flameout, run
down, continued or non-recoverable surge or stall,
or loss of acceleration and deceleration capability.
It must also be shown after the ingestion that there
is no unacceptable mechanical damage, unacceptable
power loss, or other adverse engine anomalies. The
rain ingestion must occur under the following static
ground level conditions:
(1)
A normal stabilization period at take-off power
without rain ingestion, followed immediately by the
suddenly commencing ingestion of rain for three
minutes at take-off power, then
(2)
Continuation of the rain ingestion during subsequent
rapid deceleration to minimum idle, then
(3)
Continuation of the rain ingestion during three
minutes at minimum idle power to be certified for
flight operation, then
(4)
Continuation of the rain ingestion during subsequent
rapid acceleration to take-off power.
(c)
Engines for supersonic airplanes. In addition
to complying with paragraphs (a)(1) and (a)(2) of
this section, a separate test for supersonic
airplane engines only, shall be conducted with three
hailstones ingested at supersonic cruise velocity.
These hailstones shall be aimed at the engine's
critical face area, and their ingestion must not
cause unacceptable mechanical damage or unacceptable
power or thrust loss after the ingestion or require
the engine to be shut down. The size of these
hailstones shall be determined from the linear
variation in diameter from 1-inch (25 millimeters)
at 35,000 feet (10,500 meters) to1/4-inch (6
millimeters) at 60,000 feet (18,000 meters) using
the diameter corresponding to the lowest expected
supersonic cruise altitude. Alternatively, three
larger hailstones may be ingested at subsonic
velocities such that the kinetic energy of these
larger hailstones is equivalent to the applicable
supersonic ingestion conditions.
(d)
For an engine that incorporates or requires the use
of a protection device, demonstration of the rain
and hail ingestion capabilities of the engine, as
required in paragraphs (a), (b), and (c) of this
section, may be waived wholly or in part by the
Administrator if the applicant shows that:
(1)
The subject rain and hail constituents are of a size
that will not pass through the protection device;
(2)
The protection device will withstand the impact of
the subject rain and hail constituents; and
(3)
The subject of rain and hail constituents, stopped
by the protection device, will not obstruct the flow
of induction air into the engine, resulting in
damage, power or thrust loss, or other adverse
engine anomalies in excess of what would be accepted
in paragraphs (a), (b), and (c) of this section.
Each
fuel burning thrust augmentor, including the nozzle,
must—
(a)
Provide cutoff of the fuel burning thrust augmentor;
(b)
Permit on-off cycling;
(c)
Be controllable within the intended range of
operation;
(d)
Upon a failure or malfunction of augmentor
combustion, not cause the engine to lose thrust
other than that provided by the augmentor; and
(e)
Have controls that function compatibly with the
other engine controls and automatically shut off
augmentor fuel flow if the engine rotor speed drops
below the minimum rotational speed at which the
augmentor is intended to function.
This
subpart prescribes the block tests and inspections
for turbine engines.
33.82 General.
Before each endurance test required by this subpart,
the adjustment setting and functioning
characteristic of each component having an
adjustment setting and a functioning characteristic
that can be established independent of installation
on the engine must be established and recorded.
(a)
Each engine must undergo vibration surveys to
establish that the vibration characteristics of
those components that may be subject to mechanically
or aerodynamically induced vibratory excitations are
acceptable throughout the declared flight envelope.
The engine surveys shall be based upon an
appropriate combination of experience, analysis, and
component test and shall address, as a minimum,
blades, vanes, rotor discs, spacers, and rotor
shafts.
(b)
The surveys shall cover the ranges of power or
thrust, and both the physical and corrected
rotational speeds for each rotor system,
corresponding to operations throughout the range of
ambient conditions in the declared flight envelope,
from the minimum rotational speed up to 103 percent
of the maximum physical and corrected rotational
speed permitted for rating periods of two minutes or
longer, and up to 100 percent of all other permitted
physical and corrected rotational speeds, including
those that are overspeeds. If there is any
indication of a stress peak arising at the highest
of those required physical or corrected rotational
speeds, the surveys shall be extended sufficiently
to reveal the maximum stress values present, except
that the extension need not cover more than a
further 2 percentage points increase beyond those
speeds.
(c)
Evaluations shall be made of the following:
(1)
The effects on vibration characteristics of
operating with scheduled changes (including
tolerances) to variable vane angles, compressor
bleeds, accessory loading, the most adverse inlet
air flow distortion pattern declared by the
manufacturer, and the most adverse conditions in the
exhaust duct(s); and
(2)
The aerodynamic and aeromechanical factors which
might induce or influence flutter in those systems
susceptible to that form of vibration.
(d)
Except as provided by paragraph (e) of this section,
the vibration stresses associated with the vibration
characteristics determined under this section, when
combined with the appropriate steady stresses, must
be less than the endurance limits of the materials
concerned, after making due allowances for operating
conditions for the permitted variations in
properties of the materials. The suitability of
these stress margins must be justified for each part
evaluated. If it is determined that certain
operating conditions, or ranges, need to be limited,
operating and installation limitations shall be
established.
(e)
The effects on vibration characteristics of
excitation forces caused by fault conditions (such
as, but not limited to, out-of balance, local
blockage or enlargement of stator vane passages,
fuel nozzle blockage, incorrectly schedule
compressor variables, etc.) shall be evaluated by
test or analysis, or by reference to previous
experience and shall be shown not to create a
hazardous condition.
(f)
Compliance with this section shall be substantiated
for each specific installation configuration that
can affect the vibration characteristics of the
engine. If these vibration effects cannot be fully
investigated during engine certification, the
methods by which they can be evaluated and methods
by which compliance can be shown shall be
substantiated and defined in the installation
instructions required by 33.5.
(a)
Each engine must be subjected to those calibration
tests necessary to establish its power
characteristics and the conditions for the endurance
test specified 33.87. The results of the power
characteristics calibration tests form the basis for
establishing the characteristics of the engine over
its entire operating range of speeds, pressures,
temperatures, and altitudes. Power ratings are based
upon standard atmospheric conditions with no
airbleed for aircraft services and with only those
accessories installed which are essential for engine
functioning.
(b)
A power check at sea level conditions must be
accomplished on the endurance test engine after the
endurance test and any change in power
characteristics which occurs during the endurance
test must be determined. Measurements taken during
the final portion of the endurance test may be used
in showing compliance with the requirements of this
paragraph.
(c)
In showing compliance with this section, each
condition must stabilize before measurements are
taken, except as permitted by paragraph (d) of this
section.
(d)
In the case of engines having 30-second OEI, and
2-minute OEI ratings, measurements taken during the
applicable endurance test prescribed in 33.87(f) (1)
through (8) may be used in showing compliance with
the requirements of this section for these OEI
ratings.
(a)
General. Each engine must be subjected to an
endurance test that includes a total of at least 150
hours of operation and, depending upon the type and
contemplated use of the engine, consists of one of
the series of runs specified in paragraphs (b)
through (g) of this section, as applicable. For
engines tested under paragraphs (b), (c), (d), (e)
or (g) of this section, the prescribed 6-hour test
sequence must be conducted 25 times to complete the
required 150 hours of operation. Engines for which
the 30-second OEI and 2-minute OEI ratings are
desired must be further tested under paragraph (f)
of this section. The following test requirements
apply:
(1)
The runs must be made in the order found appropriate
by the Administrator for the particular engine being
tested.
(2)
Any automatic engine control that is part of the
engine must control the engine during the endurance
test except for operations where automatic control
is normally overridden by manual control or where
manual control is otherwise specified for a
particular test run.
(3)
Except as provided in paragraph (a)(5) of this
section, power or thrust, gas temperature, rotor
shaft rotational speed, and, if limited, temperature
of external surfaces of the engine must be at least
100 percent of the value associated with the
particular engine operation being tested. More than
one test may be run if all parameters cannot be held
at the 100 percent level simultaneously.
(4)
The runs must be made using fuel, lubricants and
hydraulic fluid which conform to the specifications
specified in complying with 33.7(c).
(5)
Maximum air bleed for engine and aircraft services
must be used during at least one-fifth of the runs.
However, for these runs, the power or thrust or the
rotor shaft rotational speed may be less than 100
percent of the value associated with the particular
operation being tested if the Administrator finds
that the validity of the endurance test is not
compromised.
(6)
Each accessory drive and mounting attachment must be
loaded. The load imposed by each accessory used only
for aircraft service must be the limit load
specified by the applicant for the engine drive and
attachment point during rated maximum continuous
power or thrust and higher output. The endurance
test of any accessory drive and mounting attachment
under load may be accomplished on a separate rig if
the validity of the test is confirmed by an approved
analysis.
(7)
During the runs at any rated power or thrust the gas
temperature and the oil inlet temperature must be
maintained at the limiting temperature except where
the test periods are not longer than 5 minutes and
do not allow stabilization. At least one run must be
made with fuel, oil, and hydraulic fluid at the
minimum pressure limit and at least one run must be
made with fuel, oil, and hydraulic fluid at the
maximum pressure limit with fluid temperature
reduced as necessary to allow maximum pressure to be
attained.
(8)
If the number of occurrences of either transient
rotor shaft overspeed or transient gas
overtemperature is limited, that number of the
accelerations required by paragraphs (b) through (g)
of this section must be made at the limiting
overspeed or overtemperature. If the number of
occurrences is not limited, half the required
accelerations must be made at the limiting overspeed
or overtemperature.
(9)
For each engine type certificated for use on
supersonic aircraft the following additional test
requirements apply:
(i)
To change the thrust setting, the power control
lever must be moved from the initial position to the
final position in not more than one second except
for movements into the fuel burning thrust augmentor
augmentation position if additional time to confirm
ignition is necessary.
(ii)
During the runs at any rated augmented thrust the
hydraulic fluid temperature must be maintained at
the limiting temperature except where the test
periods are not long enough to allow stabilization.
(iii) During the simulated supersonic runs the fuel
temperature and induction air temperature may not be
less than the limiting temperature.
(iv)
The endurance test must be conducted with the fuel
burning thrust augmentor installed, with the primary
and secondary exhaust nozzles installed, and with
the variable area exhaust nozzles operated during
each run according to the methods specified in
complying with 33.5(b).
(v)
During the runs at thrust settings for maximum
continuous thrust and percentages thereof, the
engine must be operated with the inlet air
distortion at the limit for those thrust settings.
(b)
Engines other than certain rotorcraft engines.
For each engine except a rotorcraft engine for
which a rating is desired under paragraph (c), (d),
or (e) of this section, the applicant must conduct
the following runs:
(1)
Take-off and idling. One hour of alternate
five-minute periods at rated take-off power and
thrust and at idling power and thrust. The developed
powers and thrusts at take-off and idling conditions
and their corresponding rotor speed and gas
temperature conditions must be as established by the
power control in accordance with the schedule
established by the manufacturer. The applicant may,
during any one period, manually control the rotor
speed, power, and thrust while taking data to check
performance. For engines with augmented take-off
power ratings that involve increases in turbine
inlet temperature, rotor speed, or shaft power, this
period of running at take-off must be at the
augmented rating. For engines with augmented
take-off power ratings that do not materially
increase operating severity, the amount of running
conducted at the augmented rating is determined by
the Administrator. In changing the power setting
after each period, the power-control lever must be
moved in the manner prescribed in paragraph (b)(5)
of this section.
(2)
Rated maximum continuous and take-off power and
thrust. Thirty minutes at—
(i)
Rated maximum continuous power and thrust during
fifteen of the twenty-five 6-hour endurance test
cycles; and
(ii)
Rated take-off power and thrust during ten of the
twenty-five 6-hour endurance test cycles.
(3)
Rated maximum continuous power and thrust.
One hour and 30 minutes at rated maximum continuous
power and thrust.
(4)
Incremental cruise power and thrust. Two
hours and 30 minutes at the successive power lever
positions corresponding to at least 15 approximately
equal speed and time increments between maximum
continuous engine rotational speed and ground or
minimum idle rotational speed. For engines operating
at constant speed, the thrust and power may be
varied in place of speed. If there is significant
peak vibration anywhere between ground idle and
maximum continuous conditions, the number of
increments chosen may be changed to increase the
amount of running made while subject to the peak
vibrations up to not more than 50 percent of the
total time spent in incremental running.
(5)
Acceleration and deceleration runs. 30
minutes of accelerations and decelerations,
consisting of six cycles from idling power and
thrust to rated take-off power and thrust and
maintained at the take-off power lever position for
30 seconds and at the idling power lever position
for approximately four and one-half minutes. In
complying with this paragraph, the power-control
lever must be moved from one extreme poition to the
other in not more than one second, except that, if
different regimes of control operations are
incorporated necessitating scheduling of the
power-control lever motion in going from one extreme
position to the other, a longer period of time is
acceptable, but not more than two seconds.
(6)
Starts. One hundred starts must be made, of
which 25 starts must be preceded by at least a
two-hour engine shutdown. There must be at least 10
false engine starts, pausing for the applicant's
specified minimum fuel drainage time, before
attempting a normal start. There must be at least 10
normal restarts with not longer than 15 minutes
since engine shutdown. The remaining starts may be
made after completing the 150 hours of endurance
testing.
(c)
Rotorcraft engines for which a 30-minute OEI
power rating is desired. For each rotorcraft
engine for which a 30-minute OEI power rating is
desired, the applicant must conduct the following
series of tests:
(1)
Take-off and idling. One hour of alternate
5-minute periods at rated take-off power and at
idling power. The developed powers at take-off and
idling conditions and their corresponding rotor
speed and gas temperature conditions must be as
established by the power control in accordance with
the schedule established by the manufacturer. During
any one period, the rotor speed and power may be
controlled manually while taking data to check
performance. For engines with augmented take-off
power ratings that involve increases in turbine
inlet temperature, rotor speed, or shaft power, this
period of running at rated take-off power must be at
the augmented power rating. In changing the power
setting after each period, the power control lever
must be moved in the manner prescribed in paragraph
(c)(5) of this section.
(2)
Rated 30-minute OEI power. Thirty minutes at
rated 30-minute OEI power.
(3)
Rated maximum continuous power. Two hours at
rated maximum continuous power.
(4)
Incremental cruise power. Two hours at the
successive power lever positions corresponding with
not less than 12 approximately equal speed and time
increments between maximum continuous engine
rotational speed and ground or minimum idle
rotational speed. For engines operating at constant
speed, power may be varied in place of speed. If
there are significant peak vibrations anywhere
between ground idle and maximum continuous
conditions, the number of increments chosen must be
changed to increase the amount of running conducted
while being subjected to the peak vibrations up to
not more than 50 percent of the total time spent in
incremental running.
(5)
Acceleration and deceleration runs. Thirty
minutes of accelerations and decelerations,
consisting of six cycles from idling power to rated
take-off power and maintained at the take-off power
lever position for 30 seconds and at the idling
power lever position for approximately 41/2minutes.
In complying with this paragraph, the power control
lever must be moved from one extreme position to the
other in not more than 1 second, except that if
different regimes of control operations are
incorporated necessitating scheduling of the power
control lever motion in going from one extreme
position to the other, a longer period of time is
acceptable, but not more than 2 seconds.
(6)
Starts. One hundred starts, of which 25
starts must be preceded by at least a two-hour
engine shutdown. There must be at least 10 false
engine starts, pausing for the applicant's specified
minimum fuel drainage time, before attempting a
normal start. There must be at least 10 normal
restarts with not longer than 15 minutes since
engine shutdown. The remaining starts may be made
after completing the 150 hours of endurance testing.
(d)
Rotorcraft engines for which a continuous OEI
rating is desired. For each rotorcraft engine
for which a continuous OEI power rating is desired,
the applicant must conduct the following series of
tests:
(1)
Take-off and idling. One hour of alternate
5-minute periods at rated take-off power and at
idling power. The developed powers at take-off and
idling conditions and their corresponding rotor
speed and gas temperature conditions must be as
established by the power control in accordance with
the schedule established by the manufacturer. During
any one period the rotor speed and power may be
controlled manually while taking data to check
performance. For engines with augmented take-off
power ratings that involve increases in turbine
inlet temperature, rotor speed, or shaft power, this
period of running at rated take-off power must be at
the augmented power rating. In changing the power
setting after each period, the power control lever
must be moved in the manner prescribed in paragraph
(c)(5) of this section.
(2)
Rated maximum continuous and take-off power.
Thirty minutes at—
(i)
Rated maximum continuous power during fifteen of the
twenty-five 6-hour endurance test cycles; and
(ii)
Rated take-off power during ten of the twenty-five
6-hour endurance test cycles.
(3)
Rated continuous OEI power. One hour at rated
continuous OEI power.
(4)
Rated maximum continuous power. One hour at
rated maximum continuous power.
(5)
Incremental cruise power. Two hours at the
successive power lever positions corresponding with
not less than 12 approximately equal speed and time
increments between maximum continuous engine
rotational speed and ground or minimum idle
rotational speed. For engines operating at constant
speed, power may be varied in place of speed. If
there are significant peak vibrations anywhere
between ground idle and maximum continuous
conditions, the number of increments chosen must be
changed to increase the amount of running conducted
while being subjected to the peak vibrations up to
not more than 50 percent of the total time spent in
incremental running.
(6)
Acceleration and deceleration runs. Thirty
minutes of accelerations and decelerations,
consisting of six cycles from idling power to rated
take-off power and maintained at the take-off power
lever position for 30 seconds and at the idling
power lever position for approximately 41/2minutes.
In complying with this paragraph, the power control
lever must be moved from one extreme position to the
other in not more than 1 second, except that if
different regimes of control operations are
incorporated necessitating scheduling of the power
control lever motion in going from one extreme
position to the other, a longer period of time is
acceptable, but not more than 2 seconds.
(7)
Starts. One hundred starts, of which 25
starts must be preceded by at least a 2-hour engine
shutdown. There must be at least 10 false engine
starts, pausing for the applicant's specified
minimum fuel drainage time, before attempting a
normal start. There must be at least 10 normal
restarts with not longer than 15 minutes since
engine shutdown. The remaining starts may be made
after completing the 150 hours of endurance testing.
(e)
Rotorcraft engines for which a 21/2-minute OEI
power rating is desired. For each rotorcraft
engine for which a 21/2-minute OEI power rating is
desired, the applicant must conduct the following
series of tests:
(1)
Take-off, 21/2-minute OEI, and idling. One
hour of alternate 5-minute periods at rated take-off
power and at idling power except that, during the
third and sixth take-off power periods, only
21/2minutes need be conducted at rated take-off
power, and the remaining 21/2minutes must be
conducted at rated 21/2-minute OEI power. The
developed powers at take-off, 21/2-minute OEI, and
idling conditions and their corresponding rotor
speed and gas temperature conditions must be as
established by the power control in accordance with
the schedule established by the manufacturer. The
applicant may, during any one period, control
manually the rotor speed and power while taking data
to check performance. For engines with augmented
take-off power ratings that involve increases in
turbine inlet temperature, rotor speed, or shaft
power, this period of running at rated take-off
power must be at the augmented rating. In changing
the power setting after or during each period, the
power control lever must be moved in the manner
prescribed in paragraph (d)(6) of this section.
(2)
The tests required in paragraphs (b)(2) through
(b)(6), or (c)(2) through (c)(6), or (d)(2) through
(d)(7) of this section, as applicable, except that
in one of the 6-hour test sequences, the last 5
minutes of the 30 minutes at take-off power test
period of paragraph (b)(2) of this section, or of
the 30 minutes at 30-minute OEI power test period of
paragraph (c)(2) of this section, or of the l hour
at continuous OEI power test period of paragraph
(d)(3) of this section, must be run at 21/2-minute
OEI power.
(f)
Rotorcraft engines for which 30-second OEI and
2-minute OEI ratings are desired. For each
rotorcraft engine for which 30-second OEI and
2-minute OEI power ratings are desired, and
following completion of the tests under paragraphs
(b), (c), (d), or (e) of this section, the applicant
may disassemble the tested engine to the extent
necessary to show compliance with the requirements
of 33.93(a). The tested engine must then be
reassembled using the same parts used during the
test runs of paragraphs (b), (c), (d), or (e) of
this section, except those parts described as
consumables in the Instructions for Continued
Airworthiness. The applicant must then conduct the
following test sequence four times, for a total time
of not less than 120 minutes:
(1)
Take-off power. Three minutes at rated
take-off power.
(2)
30-second OEI power. Thirty seconds at rated
30-second OEI power.
(3)
2-minute OEI power. Two minutes at rated
2-minute OEI power.
(4)
30-minute OEI power, continuous OEI power, or
maximum continuous power. Five minutes at rated
30-minute OEI power, rated continuous OEI power, or
rated maximum continuous power, whichever is
greatest, except that, during the first test
sequence, this period shall be 65 minutes.
(5)
50 percent take-off power. One minute at 50
percent take-off power.
(6)
30-second OEI power. Thirty seconds at rated
30-second OEI power.
(7)
2-minute OEI power. Two minutes at rated
2-minute OEI power.
(8)
Idle. One minute at idle.
(g)
Supersonic aircraft engines. For each engine
type certificated for use on supersonic aircraft the
applicant must conduct the following:
(1)
Subsonic test under sea level ambient atmospheric
conditions. Thirty runs of one hour each must be
made, consisting of—
(i)
Two periods of 5 minutes at rated take-off augmented
thrust each followed by 5 minutes at idle thrust;
(ii)
One period of 5 minutes at rated take-off thrust
followed by 5 minutes at not more than 15 percent of
rated take-off thrust;
(iii) One period of 10 minutes at rated take-off
augmented thrust followed by 2 minutes at idle
thrust, except that if rated maximum continuous
augmented thrust is lower than rated take-off
augmented thrust, 5 of the 10-minute periods must be
at rated maximum continuous augmented thrust; and
(iv)
Six periods of 1 minute at rated take-off augmented
thrust each followed by 2 minutes, including
acceleration and deceleration time, at idle thrust.
(2)
Simulated supersonic test. Each run of the
simulated supersonic test must be preceded by
changing the inlet air temperature and pressure from
that attained at subsonic condition to the
temperature and pressure attained at supersonic
velocity, and must be followed by a return to the
temperature attained at subsonic condition. Thirty
runs of 4 hours each must be made, consisting of—
(i)
One period of 30 minutes at the thrust obtained with
the power control lever set at the position for
rated maximum continuous augmented thrust followed
by 10 minutes at the thrust obtained with the power
control lever set at the position for 90 percent of
rated maximum continuous augmented thrust. The end
of this period in the first five runs must be made
with the induction air temperature at the limiting
condition of transient overtemperature, but need not
be repeated during the periods specified in
paragraphs (g)(2)(ii) through (iv) of this section;
(ii)
One period repeating the run specified in paragraph
(g)(2)(i) of this section, except that it must be
followed by 10 minutes at the thrust obtained with
the power control lever set at the position for 80
percent of rated maximum continuous augmented
thrust;
(iii) One period repeating the run specified in
paragraph (g)(2)(i) of this section, except that it
must be followed by 10 minutes at the thrust
obtained with the power control lever set at the
position for 60 percent of rated maximum continuous
augmented thrust and then 10 minutes at not more
than 15 percent of rated take-off thrust;
(iv)
One period repeating the runs specified in
paragraphs (g)(2)(i) and (ii) of this section; and
(v)
One period of 30 minutes with 25 of the runs made at
the thrust obtained with the power control lever set
at the position for rated maximum continuous
augmented thrust, each followed by idle thrust and
with the remaining 5 runs at the thrust obtained
with the power control lever set at the position for
rated maximum continuous augmented thrust for 25
minutes each, followed by subsonic operation at not
more than 15 percent or rated take-off thrust and
accelerated to rated take-off thrust for 5 minutes
using hot fuel.
(3)
Starts. One hundred starts must be made, of
which 25 starts must be preceded by an engine
shutdown of at least 2 hours. There must be at least
10 false engine starts, pausing for the applicant's
specified minimum fuel drainage time before
attempting a normal start. At least 10 starts must
be normal restarts, each made no later than 15
minutes after engine shutdown. The starts may be
made at any time, including the period of endurance
testing.
(a)
Each engine must run for 5 minutes at maximum
permissible rpm with the gas temperature at least 75
°F (42 °C) higher than the maximum rating's
steady-state operating limit, excluding maximum
values of rpm and gas temperature associated with
the 30-second OEI and 2-minute OEI ratings.
Following this run, the turbine assembly must be
within serviceable limits.
(b)
Each engine for which 30-second OEI and 2-minute OEI
ratings are desired, that does not incorporate a
means to limit temperature, must be run for a period
of 5 minutes at the maximum power-on rpm with the
gas temperature at least 75 °F (42 °C) higher than
the 30-second OEI rating operating limit. Following
this run, the turbine assembly may exhibit distress
beyond the limits for an overtemperature condition
provided the engine is shown by analysis or test, as
found necessary by the Administrator, to maintain
the integrity of the turbine assembly.
(c)
Each engine for which 30-second OEI and 2-minute OEI
ratings are desired, that incorporates a means to
limit temperature, must be run for a period of 4
minutes at the maximum power-on rpm with the gas
temperature at least 35 °F (20 °C) higher than the
maximum operating limit. Following this run, the
turbine assembly may exhibit distress beyond the
limits for an overtemperature condition provided the
engine is shown by analysis or test, as found
necessary by the Administrator, to maintain the
integrity of the turbine assembly.
(d)
A separate test vehicle may be used for each test
condition.
(a)
The operation test must include testing found
necessary by the Administrator to demonstrate—
(1)
Starting, idling, acceleration, overspeeding,
ignition, functioning of the propeller (if the
engine is designated to operate with a propeller);
(2)
Compliance with the engine response requirements of
33.73; and
(3)
The minimum power or thrust response time to 95
percent rated take-off power or thrust, from power
lever positions representative of minimum idle and
of minimum flight idle, starting from stabilized
idle operation, under the following engine load
conditions:
(i)
No bleed air and power extraction for aircraft use.
(ii)
Maximum allowable bleed air and power extraction for
aircraft use.
(iii) An intermediate value for bleed air and power
extraction representative of that which might be
used as a maximum for aircraft during approach to a
landing.
(4)
If testing facilities are not available, the
determination of power extraction required in
paragraph (a)(3)(ii) and (iii) of this section may
be accomplished through appropriate analytical
means.
(b)
The operation test must include all testing found
necessary by the Administrator to demonstrate that
the engine has safe operating characteristics
throughout its specified operating envelope.
Each
applicant, except an applicant for an engine being
type certificated through amendment of an existing
type certificate or through supplemental type
certification procedures, must complete one of the
following tests on an engine that substantially
conforms to the type design to establish when the
initial maintenance inspection is required:
(a)
An approved engine test that simulates the
conditions in which the engine is expected to
operate in service, including typical start-stop
cycles.
(b)
An approved engine test conducted in accordance with
33.201 (c) through (f).
(a)
For those systems that cannot be adequately
substantiated by endurance testing in accordance
with the provisions of 33.87, additional tests must
be made to establish that components are able to
function reliably in all normally anticipated flight
and atmospheric conditions.
(b)
Temperature limits must be established for those
components that require temperature controlling
provisions in the aircraft installation to assure
satisfactory functioning, reliability, and
durability.
(c)
Each unpressurized hydraulic fluid tank may not fail
or leak when subjected to maximum operating
temperature and an internal pressure of 5 p.s.i.,
and each pressurized hydraulic fluid tank may not
fail or leak when subjected to maximum operating
temperature and an internal pressure not less than 5
p.s.i. plus the maximum operating pressure of the
tank.
(d)
For an engine type certificated for use in
supersonic aircraft, the systems, safety devices,
and external components that may fail because of
operation at maximum and minimum operating
temperatures must be identified and tested at
maximum and minimum operating temperatures and while
temperature and other operating conditions are
cycled between maximum and minimum operating values.
If
continued rotation is prevented by a means to lock
the rotor(s), the engine must be subjected to a test
that includes 25 operations of this means under the
following conditions:
(a)
The engine must be shut down from rated maximum
continuous thrust or power; and
(b)
The means for stopping and locking the rotor(s) must
be operated as specified in the engine operating
instructions while being subjected to the maximum
torque that could result from continued flight in
this condition; and
(c)
Following rotor locking, the rotor(s) must be held
stationary under these conditions for five minutes
for each of the 25 operations.
(a)
After completing the endurance testing of 33.87 (b),
(c), (d), (e), or (g) of this part, each engine must
be completely disassembled, and
(1)
Each component having an adjustment setting and a
functioning characteristic that can be established
independent of installation on the engine must
retain each setting and functioning characteristic
within the limits that were established and recorded
at the beginning of the test; and
(2)
Each engine part must conform to the type design and
be eligible for incorporation into an engine for
continued operation, in accordance with information
submitted in compliance with 33.4.
(b)
After completing the endurance testing of 33.87(f),
each engine must be completely disassembled, and
(1)
Each component having an adjustment setting and a
functioning characteristic that can be established
independent of installation on the engine must
retain each setting and functioning characteristic
within the limits that were established and recorded
at the beginning of the test; and
(2)
Each engine may exhibit deterioration in excess of
that permitted in paragraph (a)(2) of this section
including some engine parts or components that may
be unsuitable for further use. The applicant must
show by analysis and/or test, as found necessary by
the Administrator, that structural integrity of the
engine including mounts, cases, bearing supports,
shafts, and rotors, is maintained; or
(c)
In lieu of compliance with paragraph (b) of this
section, each engine for which the 30-second OEI and
2-minute OEI ratings are desired, may be subjected
to the endurance testing of 33.87 (b), (c), (d), or
(e) of this part, and followed by the testing of
33.87(f) without intervening disassembly and
inspection. However, the engine must comply with
paragraph (a) of this section after completing the
endurance testing of 33.87(f).
(a)
Except as provided in paragraph (b) of this section,
it must be demonstrated by engine tests that the
engine is capable of containing damage without
catching fire and without failure of its mounting
attachments when operated for at least 15 seconds,
unless the resulting engine damage induces a self
shutdown, after each of the following events:
(1)
Failure of the most critical compressor or fan blade
while operating at maximum permissible r.p.m. The
blade failure must occur at the outermost retention
groove or, for integrally-bladed rotor discs, at
least 80 percent of the blade must fail.
(2)
Failure of the most critical turbine blade while
operating at maximum permissible r.p.m. The blade
failure must occur at the outermost retention groove
or, for integrally-bladed rotor discs, at least 80
percent of the blade must fail. The most critical
turbine blade must be determined by considering
turbine blade weight and the strength of the
adjacent turbine case at case temperatures and
pressures associated with operation at maximum
permissible r.p.m.
(b)
Analysis based on rig testing, component testing, or
service experience may be substitute for one of the
engine tests prescribed in paragraphs (a)(1) and
(a)(2) of this section if—
(1)
That test, of the two prescribed, produces the least
rotor unbalance; and
(2)
The analysis is shown to be equivalent to the test.
33.95 Engine-propeller systems tests.
If
the engine is designed to operate with a propeller,
the following tests must be made with a
representative propeller installed by either
including the tests in the endurance run or
otherwise performing them in a manner acceptable to
the Administrator:
(a)
Feathering operation: 25 cycles.
(b)
Negative torque and thrust system operation: 25
cycles from rated maximum continuous power.
(c)
Automatic de-coupler operation: 25 cycles from rated
maximum continuous power (if repeated decoupling and
re-coupling in service is the intended function of
the device).
(d)
Reverse thrust operation: 175 cycles from the
flight-idle position to full reverse and 25 cycles
at rated maximum continuous power from full forward
to full reverse thrust. At the end of each cycle the
propeller must be operated in reverse pitch for a
period of 30 seconds at the maximum rotational speed
and power specified by the applicant for reverse
pitch operation.
If
the engine is designed with a propeller brake which
will allow the propeller to be brought to a stop
while the gas generator portion of the engine
remains in operation, and remain stopped during
operation of the engine as an auxiliary power unit
(“APU mode”), in addition to the requirements of
33.87, the applicant must conduct the following
tests:
(a)
Ground locking: A total of 45 hours with the
propeller brake engaged in a manner which clearly
demonstrates its ability to function without adverse
effects on the complete engine while the engine is
operating in the APU mode under the maximum
conditions of engine speed, torque, temperature, air
bleed, and power extraction as specified by the
applicant.
(b)
Dynamic braking: A total of 400 application-release
cycles of brake engagements must be made in a manner
which clearly demonstrates its ability to function
without adverse effects on the complete engine under
the maximum conditions of engine
acceleration/deceleration rate, speed, torque, and
temperature as specified by the applicant. The
propeller must be stopped prior to brake release.
(c)
One hundred engine starts and stops with the
propeller brake engaged.
(d)
The tests required by paragraphs (a), (b), and (c)
of this section must be performed on the same
engine, but this engine need not be the same engine
used for the tests required by 33.87.
(e)
The tests required by paragraphs (a), (b), and (c)
of this section must be followed by engine
disassembly to the extent necessary to show
compliance with the requirements of 33.93(a) and
33.93(b).
(a)
If the engine incorporates a reverser, the endurance
calibration, operation, and vibration tests
prescribed in this subpart must be run with the
reverser installed. In complying with this section,
the power control lever must be moved from one
extreme position to the other in not more than one
second except, if regimes of control operations are
incorporated necessitating scheduling of the
power-control lever motion in going from one extreme
position to the other, a longer period of time is
acceptable but not more than three seconds. In
addition, the test prescribed in paragraph (b) of
this section must be made. This test may be
scheduled as part of the endurance run.
(b)
175 reversals must be made from flight-idle forward
thrust to maximum reverse thrust and 25 reversals
must be made from rated take-off thrust to maximum
reverse thrust. After each reversal the reverser
must be operated at full reverse thrust for a period
of one minute, except that, in the case of a
reverser intended for use only as a braking means on
the ground, the reverser need only be operated at
full reverse thrust for 30 seconds.
(a)
Each applicant may, in making a block test, use
separate engines of identical design and
construction in the vibration, calibration,
endurance, and operation tests, except that, if a
separate engine is used for the endurance test it
must be subjected to a calibration check before
starting the endurance test.
(b)
Each applicant may service and make minor repairs to
the engine during the block tests in accordance with
the service and maintenance instructions submitted
in compliance with 33.4. If the frequency of the
service is excessive, or the number of stops due to
engine malfunction is excessive, or a major repair,
or replacement of a part is found necessary during
the block tests or as the result of findings from
the teardown inspection, the engine or its parts
must be subjected to any additional tests the
Administrator finds necessary.
(c)
Each applicant must furnish all testing facilities,
including equipment and competent personnel, to
conduct the block tests.
An
applicant seeking type design approval for an engine
to be installed on a two-engine airplane approved
for ETOPS without the service experience specified
in part 25, Appendix K, K25.2.1 of this chapter,
must comply with the following:
(a)
The engine must be designed using a design quality
process acceptable to the AFRO-CAA, that ensures the
design features of the engine minimize the
occurrence of failures, malfunctions, defects, and
maintenance errors that could result in an IFSD,
loss of thrust control, or other power loss.
(b)
The design features of the engine must address
problems shown to result in an IFSD, loss of thrust
control, or other power loss in the applicant's
other relevant type designs approved within the past
10 years, to the extent that adequate service data
is available within that 10-year period. An
applicant without adequate service data must show
experience with and knowledge of problem mitigating
design practices equivalent to that gained from
actual service experience in a manner acceptable to
the AFRO-CAA.
(c)
Except as specified in paragraph (f) of this
section, the applicant must conduct a simulated
ETOPS mission cyclic endurance test in accordance
with an approved test plan on an engine that
substantially conforms to the type design. The test
must:
(1)
Include a minimum of 3,000 representative service
start-stop mission cycles and three simulated
diversion cycles at maximum continuous thrust or
power for the maximum diversion time for which ETOPS
eligibility is sought. Each start-stop mission cycle
must include the use of take-off, climb, cruise,
descent, approach, and landing thrust or power and
the use of thrust reverse (if applicable). The
diversions must be evenly distributed over the
duration of the test. The last diversion must be
conducted within 100 cycles of the completion of the
test.
(2)
Be performed with the high speed and low speed main
engine rotors independently unbalanced to obtain a
minimum of 90 percent of the recommended field
service maintenance vibration levels. For engines
with three main engine rotors, the intermediate
speed rotor must be independently unbalanced to
obtain a minimum of 90 percent of the recommended
production acceptance vibration level. The required
peak vibration levels must be verified during a slow
acceleration and deceleration run of the test engine
covering the main engine rotor operating speed
ranges.
(3)
Include a minimum of three million vibration cycles
for each 60 rpm incremental step of the typical
high-speed rotor start-stop mission cycle. The test
may be conducted using any rotor speed step
increment from 60 to 200 rpm provided the test
encompasses the typical service start-stop cycle
speed range. For incremental steps greater than 60
rpm, the minimum number of vibration cycles must be
linearly increased up to ten million cycles for a
200 rpm incremental step.
(4)
Include a minimum of 300,000 vibration cycles for
each 60 rpm incremental step of the high-speed rotor
approved operational speed range between minimum
flight idle and cruise power not covered by
paragraph (c)(3) of this section. The test may be
conducted using any rotor speed step increment from
60 to 200 rpm provided the test encompasses the
applicable speed range. For incremental steps
greater than 60 rpm the minimum number of vibration
cycles must be linearly increased up to 1 million
for a 200 rpm incremental step.
(5)
Include vibration surveys at periodic intervals
throughout the test. The equivalent value of the
peak vibration level observed during the surveys
must meet the minimum vibration requirement of
33.201(c)(2).
(d)
Prior to the test required by paragraph (c) of this
section, the engine must be subjected to a
calibration test to document power and thrust
characteristics.
(e)
At the conclusion of the testing required by
paragraph (c) of this section, the engine must:
(1)
Be subjected to a calibration test at sea-level
conditions. Any change in power or thrust
characteristics must be within approved limits.
(2)
Be visually inspected in accordance with the on-wing
inspection recommendations and limits contained in
the Instructions for Continued Airworthiness
submitted in compliance with 33.4.
(3)
Be completely disassembled and inspected—
(i)
In accordance with the applicable inspection
recommendations and limits contained in the
Instructions for Continued Airworthiness submitted
in compliance with 33.4;
(ii)
With consideration of the causes of IFSD, loss of
thrust control, or other power loss identified by
paragraph (b) of this section; and
(iii) In a manner to identify wear or distress
conditions that could result in an IFSD, loss of
thrust control, or other power loss not specifically
identified by paragraph (b) of this section or
addressed within the Instructions for Continued
Airworthiness.
(4)
Not show wear or distress to the extent that could
result in an IFSD, loss of thrust control, or other
power loss within a period of operation before the
component, assembly, or system would likely have
been inspected or functionally tested for integrity
while in service. Such wear or distress must have
corrective action implemented through a design
change, a change to maintenance instructions, or
operational procedures before ETOPS eligibility is
granted. The type and frequency of wear and distress
that occurs during the engine test must be
consistent with the type and frequency of wear and
distress that would be expected to occur on ETOPS
eligible engines.
(f)
An alternative mission cycle endurance test that
provides an equivalent demonstration of the
unbalance and vibration specified in paragraph (c)
of this section may be used when approved by the
AFRO-CAA.
(g)
For an applicant using the simulated ETOPS mission
cyclic endurance test to comply with 33.90, the test
may be interrupted so that the engine may be
inspected by an on-wing or other method, using
criteria acceptable to the AFRO-CAA, after
completion of the test cycles required to comply
with 33.90(a). Following the inspection, the ETOPS
test must be resumed to complete the requirements of
this section.
Appendix A to Part
33—Instructions for Continued Airworthiness
a33.1 general
(a)
This appendix specifies requirements for the
preparation of Instructions for Continued
Airworthiness as required by 33.4.
(b)
The Instructions for Continued Airworthiness for
each engine must include the Instructions for
Continued Airworthiness for all engine parts. If
Instructions for Continued Airworthiness are not
supplied by the engine part manufacturer for an
engine part, the Instructions for Continued
Airworthiness for the engine must include the
information essential to the continued airworthiness
of the engine.
(c)
The applicant must submit to the AFRO-CAA a program
to show how changes to the Instructions for
Continued Airworthiness made by the applicant or by
the manufacturers of engine parts will be
distributed.
a33.2 format
(a)
The Instructions for Continued Airworthiness must be
in the form of a manual or manuals as appropriate
for the quantity of data to be provided.
(b)
The format of the manual or manuals must provide for
a practical arrangement.
a33.3 content
The
contents of the manual or manuals must be prepared
in the English language. The Instructions for
Continued Airworthiness must contain the following
manuals or sections, as appropriate, and
information:
(a)
Engine Maintenance Manual or Section. (1)
Introduction information that includes an
explanation of the engine's features and data to the
extent necessary for maintenance or preventive
maintenance.
(2)
A detailed description of the engine and its
components, systems, and installations.
(3)
Installation instructions, including proper
procedures for uncrating, de-inhibiting, acceptance
checking, lifting, and attaching accessories, with
any necessary checks.
(4)
Basic control and operating information describing
how the engine components, systems, and
installations operate, and information describing
the methods of starting, running, testing, and
stopping the engine and its parts including any
special procedures and limitations that apply.
(5)
Servicing information that covers details regarding
servicing points, capacities of tanks, reservoirs,
types of fluids to be used, pressures applicable to
the various systems, locations of lubrication
points, lubricants to be used, and equipment
required for servicing.
(6)
Scheduling information for each part of the engine
that provides the recommended periods at which it
should be cleaned, inspected, adjusted, tested, and
lubricated, and the degree of inspection the
applicable wear tolerances, and work recommended at
these periods. However, the applicant may refer to
an accessory, instrument, or equipment manufacturer
as the source of this information if the applicant
shows that the item has an exceptionally high degree
of complexity requiring specialized maintenance
techniques, test equipment, or expertise. The
recommended overhaul periods and necessary cross
references to the Airworthiness Limitations section
of the manual must also be included. In addition,
the applicant must include an inspection program
that includes the frequency and extent of the
inspections necessary to provide for the continued
airworthiness of the engine.
(7)
Troubleshooting information describing probable
malfunctions, how to recognize those malfunctions,
and the remedial action for those malfunctions.
(8)
Information describing the order and method of
removing the engine and its parts and replacing
parts, with any necessary precautions to be taken.
Instructions for proper ground handling, crating,
and shipping must also be included.
(9)
A list of the tools and equipment necessary for
maintenance and directions as to their method of
use.
(b)
Engine Overhaul Manual or Section. (1)
Disassembly information including the order and
method of disassembly for overhaul.
(2)
Cleaning and inspection instructions that cover the
materials and apparatus to be used and methods and
precautions to be taken during overhaul. Methods of
overhaul inspection must also be included.
(3)
Details of all fits and clearances relevant to
overhaul.
(4)
Details of repair methods for worn or otherwise
substandard parts and components along with the
information necessary to determine when replacement
is necessary.
(5)
The order and method of assembly at overhaul.
(6)
Instructions for testing after overhaul.
(7)
Instructions for storage preparation, including any
storage limits.
(8)
A list of tools needed for overhaul.
(c)
ETOPS Requirements. For an applicant seeking
eligibility for an engine to be installed on an
airplane approved for ETOPS, the Instructions for
Continued Airworthiness must include procedures for
engine condition monitoring. The engine condition
monitoring procedures must be able to determine
prior to flight, whether an engine is capable of
providing, within approved engine operating limits,
maximum continuous power or thrust, bleed air, and
power extraction required for a relevant engine
inoperative diversion. For an engine to be installed
on a two-engine airplane approved for ETOPS, the
engine condition monitoring procedures must be
validated before ETOPS eligibility is granted.
a33.4 airworthiness limitations section
The
Instructions for Continued Airworthiness must
contain a section titled Airworthiness Limitations
that is segregated and clearly distinguishable from
the rest of the document. This section must set
forth each mandatory replacement time, inspection
interval, and related procedure required for type
certification. If the Instructions for Continued
Airworthiness consist of multiple documents, the
section required by this paragraph must be included
in the principal manual. This section must contain a
legible statement in a prominent location that
reads: “The Airworthiness Limitations section is
AFRO-CAA approved and specifies maintenance required
under 43.16 and 91.403 of the African Civil Aviation
Agency Regulations unless an alternative program has
been AFRO-CAA approved.”
Figure B1, Table B1, Table B2, Table B3, and Table
B4 specify the atmospheric concentrations and size
distributions of rain and hail for establishing
certification, in accordance with the requirements
of 33.78(a)(2). In conducting tests, normally by
spraying liquid water to simulate rain conditions
and by delivering hail fabricated from ice to
simulate hail conditions, the use of water droplets
and hail having shapes, sizes and distributions of
sizes other than those defined in this appendix B,
or the use of a single size or shape for each water
droplet or hail, can be accepted, provided that
applicant shows that the substitution does not
reduce the severity of the test.
Table B1—Certification Standard Atmospheric Rain
Concentrations
|
Altitude (feet) |
Rain water content (RWC)
(grams water/meter3air) |
|
0 |
20.0 |
|
20,000 |
20.0 |
|
26,300 |
15.2 |
|
32,700 |
10.8 |
|
39,300 |
7.7 |
|
46,000 |
5.2 |
RWC values at other altitudes may be determined by linear
interpolation.
Table B2—Certification Standard Atmospheric Hail
Concentrations
|
Altitude (feet) |
Hail water content (HWC)
(grams water/meter3air) |
|
0 |
6.0 |
|
7,300 |
8.9 |
|
8,500 |
9.4 |
|
10,000 |
9.9 |
|
12,000 |
10.0 |
|
15,000 |
10.0 |
|
16,000 |
8.9 |
|
17,700 |
7.8 |
|
19,300 |
6.6 |
|
21,500 |
5.6 |
|
24,300 |
4.4 |
|
29,000 |
3.3 |
|
46,000 |
0.2 |
HWC values at other altitudes may be determined by linear
interpolation. The hail threat below 7,300 feet and
above 29,000 feet is based on linearly extrapolated
data.
Table B3—Certification Standard Atmospheric Rain Droplet Size
Distribution
|
Rain droplet diameter (mm) |
Contribution total RWC (%) |
|
0–0.49 |
0 |
|
0.50–0.99 |
2.25 |
|
1.00–1.49 |
8.75 |
|
1.50–1.99 |
16.25 |
|
2.00–2.49 |
19.00 |
|
2.50–2.99 |
17.75 |
|
3.00–3.49 |
13.50 |
|
3.50–3.99 |
9.50 |
|
4.00–4.49 |
6.00 |
|
4.50–4.99 |
3.00 |
|
5.00–5.49 |
2.00 |
|
5.50–5.99 |
1.25 |
|
6.00–6.49 |
0.50 |
|
6.50–7.00 |
0.25 |
|
Total |
100.00 |
Median diameter of rain droplets in 2.66 mm
Table B4—Certification Standard Atmospheric Hail Size
Distribution
|
Hail diameter (mm) |
Contribution total HWC (%) |
|
0–4.9 |
0 |
|
5.0–9.9 |
17.00 |
|
10.0–14.9 |
25.00 |
|
15.0–19.9 |
22.50 |
|
20.0–24.9 |
16.00 |
|
25.0–29.9 |
9.75 |
|
30.0–34.9 |
4.75 |
|
35.0–39.9 |
2.50 |
|
40.0–44.9 |
1.50 |
|
45.0–49.9 |
0.75 |
|
50.0–55.0 |
0.25 |
|
Total |
100.00 |
Median diameter of hail is 16 mm
|
