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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.
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