ASTM F3063/F3063M-21

Standard Specification for Aircraft Fuel Storage and Delivery

ABSTRACT
This specification prescribes minimum requirements for the design and integration of Fuel/Energy Storage and Delivery system installations for aeroplanes and is applicable to aeroplanes as defined in the F44 terminology standard. The applicant for a design approval must seek the individual guidance to their respective civil aviation authority (CAA) body concerning the use of this specification as part of a certification plan.
The requirements provided in this specification cover fuel system, fuel tanks, fuel pumps, fuel flow, pressure fueling systems, and fuel jettisoning system.
SCOPE
1.1 This specification covers minimum requirements for the design and integration of Fuel Storage and Delivery system installations for aeroplanes.  
1.2 This specification is applicable to aeroplanes as defined in the F44 terminology standard.  
1.3 The applicant for a design approval must seek the individual guidance to their respective CAA body concerning the use of this standard as part of a certification plan. For information on which CAA regulatory bodies have accepted this standard (in whole or in part) as a means of compliance to their Aeroplane Airworthiness regulations (Hereinafter referred to as “the Rules”), refer to ASTM F44 webpage (www.ASTM.org/COMITTEE/F44.htm), which includes CAA website links. Annex A1 maps the Means of Compliance described in this specification to EASA CS-23, amendment 5, or later, and FAA 14 CFR Part 23, amendment 64, or later.  
1.4 Units—The values stated are SI units followed by imperial units in brackets. The values stated in each system are not necessarily exact equivalents; therefore, to ensure conformance with the standard, each system shall be used independently of the other, and values from the two systems shall not be combined.  
1.5 This standard does not purport to address all of the safety concerns, if any, associated with its use. It is the responsibility of the user of this standard to establish appropriate safety, health, and environmental practices and determine the applicability of regulatory limitations prior to use.  
1.6 This international standard was developed in accordance with internationally recognized principles on standardization established in the Decision on Principles for the Development of International Standards, Guides and Recommendations issued by the World Trade Organization Technical Barriers to Trade (TBT) Committee.

General Information

• Status: Published
• Publication Date: 31-Aug-2021
• ICS: 49.050 - Aerospace engines and propulsion systems
• Technical Committee: F44 - General Aviation Aircraft
• Drafting Committee: F44.40 - Powerplant

Relations

• Refers: ASTM F3116/F3116M-23a - Standard Specification for Design Loads and Conditions
• Effective Date: 01-Oct-2023
• Refers: ASTM F3179/F3179M-23 - Standard Specification for Performance of Aircraft
• Effective Date: 01-Sep-2023
• Refers: ASTM F3060-20 - Standard Terminology for Aircraft
• Effective Date: 01-Jan-2020
• Refers: ASTM F3083/F3083M-18 - Standard Specification for Emergency Conditions, Occupant Safety and Accommodations
• Effective Date: 01-Dec-2018
• Refers: ASTM F3117/F3117M-18c - Standard Specification for Crew Interface in Aircraft
• Effective Date: 01-Nov-2018
• Refers: ASTM F3116/F3116M-18 - Standard Specification for Design Loads and Conditions
• Effective Date: 01-Nov-2018
• Refers: ASTM F3179/F3179M-18 - Standard Specification for Performance of Aircraft
• Effective Date: 01-May-2018
• Refers: ASTM F3060-16a - Standard Terminology for Aircraft
• Effective Date: 01-Nov-2016
• Refers: ASTM F3083/F3083M-16 - Standard Specification for Emergency Conditions, Occupant Safety and Accommodations
• Effective Date: 01-Jun-2016
• Refers: ASTM F3060-16 - Standard Terminology for Aircraft
• Effective Date: 01-Apr-2016
• Refers: ASTM F3179/F3179M-16 - Standard Specification for Performance of Aeroplanes
• Effective Date: 01-Feb-2016
• Refers: ASTM F3060-15b - Standard Terminology for Aircraft
• Effective Date: 15-Sep-2015
• Refers: ASTM F3083/F3083M-15 - Standard Specification for Emergency Conditions, Occupant Safety and Accommodations
• Effective Date: 01-Jun-2015
• Refers: ASTM F3060-15a - Standard Terminology for Aircraft
• Effective Date: 01-May-2015
• Refers: ASTM F3116/F3116M-15 - Standard Specification for Design Loads and Conditions
• Effective Date: 01-May-2015

Overview

ASTM F3063/F3063M-21 is the internationally recognized standard specification for aircraft fuel storage and delivery systems. Developed by ASTM Committee F44, this standard outlines the minimum requirements for the design, integration, installation, and testing of fuel/energy storage and delivery system installations for aeroplanes. It applies specifically to aeroplanes as defined in the F44 terminology standard and serves as a key reference for meeting various national civil aviation authority (CAA) certification requirements.

Proper implementation of this standard helps to ensure the safety, reliability, and compliance of fuel systems in general aviation, commuter, and certain commercial aircraft. All applicants for aircraft design approval should consult relevant CAA authorities for guidance on using ASTM F3063/F3063M-21 as part of their certification plans.

Key Topics

ASTM F3063/F3063M-21 covers an extensive range of critical aspects in the design and integration of aircraft fuel systems, including:

• Fuel Systems: Construction, arrangement, and flow requirements, ensuring proper fuel supply during all operating conditions.
• Fuel Tanks: Design criteria, installation, support, compartmentalization, expansion space, sump requirements, and provisions for inspection and repair.
• Pumps and Flow: Standards for main, emergency, and auxiliary pumps, along with minimum flow rates for gravity and pump-driven systems.
• Fuel System Components: Requirements for vents, drains, filters, strainers, and filler connections to ensure safe, contamination-free fuel delivery.
• Pressure Fueling and Defueling: Safety provisions to prevent overfilling and component damage during fueling and defueling operations.
• Fuel Jettisoning Systems: Criteria for safe and efficient fuel jettisoning in cases where weight reduction is necessary for landing.
• Testing and Performance Verification: Robust procedures to confirm system integrity under operating and environmental stresses.

Applications

This standard provides tangible benefits and practical application across the aviation industry, such as:

• Aircraft Certification: Offers a comprehensive framework for demonstrating compliance with CAA, FAA, and EASA airworthiness regulations regarding fuel systems.
• Aircraft Design and Manufacturing: Guides engineers and manufacturers in developing fuel storage and delivery systems that meet rigorous safety and reliability standards.
• Retrofit and Upgrades: Supports the integration of updated fuel system technologies in legacy aircraft to improve safety and operational efficiency.
• Regulatory Compliance: Enables clear mapping to regulatory requirements, facilitating the approval process with authorities like EASA CS-23 and FAA 14 CFR Part 23.
• Safety Enhancement: Minimizes fire hazards, ensures reliable fuel supply to engines, and addresses critical operational scenarios such as hot weather and emergency conditions.

Related Standards

ASTM F3063/F3063M-21 references and is complemented by several related standards to ensure holistic coverage of aircraft systems, including:

• ASTM F3060: Terminology for Aircraft, providing definitive terms and definitions.
• ASTM F3083/F3083M: Specification for Emergency Conditions, Occupant Safety and Accommodations.
• ASTM F3116/F3116M: Specification for Design Loads and Conditions.
• ASTM F3117/F3117M: Specification for Crew Interface in Aircraft.
• ASTM F3179/F3179M: Specification for Performance of Aircraft.
• ASTM F3397/F3397M: Practice for Aeroplane Turbine Fuel System Hot Weather Operations.
• EASA CS-23: European airworthiness standards for normal, utility, aerobatic, and commuter aeroplanes.
• FAA 14 CFR Part 23 & 34: US airworthiness and emission requirements for aircraft.

By following ASTM F3063/F3063M-21 in conjunction with these related standards, organizations can achieve a robust, standardized approach to aircraft fuel storage and delivery system safety and certification.

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Keywords: aircraft fuel system standard, ASTM aircraft fuel storage, aircraft fuel delivery system, aviation fuel tank requirements, aviation pressure fueling, aircraft fuel jettisoning, FAA fuel system compliance, EASA CS-23 fuel system

Standards Content (Sample)

This international standard was developed in accordance with internationally recognized principles on standardization established in the Decision on Principles for the
Development of International Standards, Guides and Recommendations issued by the World Trade Organization Technical Barriers to Trade (TBT) Committee.
Designation: F3063/F3063M −21
Standard Specification for
Aircraft Fuel Storage and Delivery
ThisstandardisissuedunderthefixeddesignationF3063/F3063M;thenumberimmediatelyfollowingthedesignationindicatestheyear
of original adoption or, in the case of revision, the year of last revision. A number in parentheses indicates the year of last reapproval.
A superscript epsilon (´) indicates an editorial change since the last revision or reapproval.
1. Scope 2. Referenced Documents
2.1 ASTM Standards:
1.1 This specification covers minimum requirements for the
F3060 Terminology for Aircraft
design and integration of Fuel Storage and Delivery system
F3083/F3083M Specification for Emergency Conditions,
installations for aeroplanes.
Occupant Safety and Accommodations
1.2 This specification is applicable to aeroplanes as defined
F3116/F3116M Specification for Design Loads and Condi-
in the F44 terminology standard.
tions
F3117/F3117M Specification for Crew Interface in Aircraft
1.3 The applicant for a design approval must seek the
F3179/F3179M Specification for Performance of Aircraft
individual guidance to their respective CAA body concerning
F3397/F3397M Practice forAeroplane Turbine Fuel System
the use of this standard as part of a certification plan. For
Hot Weather Operations
information on which CAA regulatory bodies have accepted
2.2 EASA Standards:
this standard (in whole or in part) as a means of compliance to
CS-23 Normal, Utility,Aerobatic and CommuterAeroplanes
theirAeroplaneAirworthinessregulations(Hereinafterreferred
CS-34 EasyAccess Rules forAircraft Engine Emissions and
to as “the Rules”), refer to ASTM F44 webpage
Fuel Venting
(www.ASTM.org/COMITTEE/F44.htm), which includes CAA
2.3 FAA Standards:
website links. Annex A1 maps the Means of Compliance
14 CFR Part 23 Airworthiness Standards: Normal Category
described in this specification to EASA CS-23, amendment 5,
Airplanes
or later, and FAA 14 CFR Part 23, amendment 64, or later.
14 CFR Part 34 Fuel Venting and Exhaust Emission Re-
1.4 Units—The values stated are SI units followed by
quirements for Turbine Engine Powered Airplanes
imperial units in brackets.The values stated in each system are
not necessarily exact equivalents; therefore, to ensure confor-
3. Terminology
mance with the standard, each system shall be used indepen-
3.1 The following are a selection of relevant terms. See
dently of the other, and values from the two systems shall not
Terminology F3060 for more definitions and abbreviations.
be combined.
3.2 Definitions:
1.5 This standard does not purport to address all of the
3.2.1 main pump, n—a pump that supplies sufficient fuel to
safety concerns, if any, associated with its use. It is the
sustain the engine during normal operations.
responsibility of the user of this standard to establish appro-
3.2.2 emergency pump, n—a pump that can sustain engine
priate safety, health, and environmental practices and deter-
operation at full power in the event the main pump fails.
mine the applicability of regulatory limitations prior to use.
3.2.3 auxiliary pump, n—a pump that can provide some fuel
1.6 This international standard was developed in accor-
flow during emergency operations but not enough to sustain
dance with internationally recognized principles on standard-
engine operation at full power. Its function is to aid in priming
ization established in the Decision on Principles for the
the engine and suppressing fuel vapors.
Development of International Standards, Guides and Recom-
mendations issued by the World Trade Organization Technical 3.3 Abbreviations:
Barriers to Trade (TBT) Committee.
For referenced ASTM standards, visit the ASTM website, www.astm.org, or
contact ASTM Customer Service at [e-mail]. For Annual Book of ASTM
ThisspecificationisunderthejurisdictionofASTMCommitteeF44onGeneral Standards volume information, refer to the standard’s Document Summary page on
Aviation Aircraft and is the direct responsibility of Subcommittee F44.40 on the ASTM website.
Powerplant. Available from European Union Aviation Safety Agency (EASA), Konrad-
Current edition approved Sept. 1, 2021. Published October 2021. Originally Adenauer-Ufer 3, D-50668 Cologne, Germany, https://www.easa.europa.eu.
approved in 2015. Last previous edition approved in 2020 as F3063/F3063M–20. Available from Federal Aviation Administration (FAA), 800 Independence
DOI: 10.1520/F3063_F3063M-21. Ave., SW, Washington, DC 20591, http://www.faa.gov.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
F3063/F3063M − 21
3.3.1 cc—cubic centimetre 4.2.2.4 A fuel system in which those parts of the system
from each tank outlet to any engine are independent of each
3.3.2 CFR—Code of Federal Regulations
part of the system supplying fuel to any other engine must be
3.3.3 RPM—rotation per minute
provided.
4.3 Drains:
4. Fuel System
4.3.1 There must be at least one drain to allow safe drainage
4.1 General:
of the entire fuel system with the aeroplane in its normal
4.1.1 Each fuel system must be constructed and arranged to
ground attitude.
ensure fuel flow at a rate and pressure established for proper
4.3.2 Each drain installed for the purpose of draining
engine and auxiliary power unit functioning under each likely
hazardousquantitiesofwater,asrequiredby5.6,mustmeetthe
operating condition, including any maneuver for which certi-
provision of 4.3.2.1 through 4.3.2.2.
fication is requested and during which the engine or auxiliary
4.3.2.1 Discharge clear of all parts of the aeroplane.
power unit is permitted to be in operation.
4.3.2.2 Have a drain valve that meets 4.3.2.2(1) through (6).
4.1.2 Eachfuelsystemforaturbineengineandcompression
(1) That has manual or automatic means for positive
ignition engine must be capable of sustained operation
locking in the closed position.
throughout its flow and pressure range with fuel initially
(2) That is readily accessible.
saturatedwithwaterat27 °C[80 °F]andhaving0.75ccoffree
(3) That can be easily opened and closed.
waterper3.8L[1USgal]addedandcooledtothemostcritical
(4) That is either located or protected to prevent fuel
condition for icing likely to be encountered in operation.
spillage in the event of a landing with landing gear retracted.
4.1.3 Each fuel system for a turbine engine powered aero-
(5) That allows the fuel to be caught for examination.
plane must meet the applicable fuel venting requirements of 14
(6) That can be observed for proper closing.
CFRPart34fortheUnitedStatesortheapplicablefuelventing
requirements as contained in the rules specified by the appli-
5. Fuel Tanks
cant’s local civil aviation authority.
5.1 General:
4.1.4 Each fuel system must be arranged so that it meets the
5.1.1 Each fuel tank must be able to withstand, without
requirement of 4.1.4.1 or 4.1.4.2.
failure, the vibration, inertia, fluid, and structural loads that it
4.1.4.1 No fuel pump can draw fuel from more than one
may be subjected to in operation.
tank at a time.
5.1.2 Each flexible fuel tank liner must be shown to be
4.1.4.2 There are means to prevent introducing air into the
suitable for the particular application.
system.
5.1.3 The total usable capacity of the fuel tanks must be
4.1.5 Fuel system components in an engine nacelle or in the
enoughforatleast30minofoperationatmaximumcontinuous
fuselage must be protected from damage which could result in
power.
spillage of enough fuel to constitute a fire hazard as a result of
a wheels-up landing on a paved runway.
5.2 Installation:
4.1.6 Each check valve must be constructed, or otherwise
5.2.1 Eachfueltankmustbesupportedsothattankloadsare
incorporate provisions, to preclude incorrect assembly or
not concentrated.
connection of the valve.
5.2.1.1 There must be pads, if necessary, to prevent chafing
between each tank and its supports.
4.2 Independence:
5.2.1.2 Padding must be nonabsorbent or treated to prevent
4.2.1 Each fuel system for a multiengine aeroplane must be
the absorption of fuel.
arranged so that, in at least one system configuration, the
failure of any one component will not result in the loss of 5.2.1.3 If a flexible tank liner is used, it must be supported
power of more than one engine or require immediate action by so that it is not required to withstand fluid loads.
the pilot to prevent the loss of power of more than one engine. 5.2.1.4 Interior surfaces adjacent to the liner must be
4.2.2 If a single fuel tank (or series of fuel tanks intercon- smooth and free from projections that could cause wear, unless
provisions are made for protection of the liner at those points;
nected to function as a single fuel tank) is used on a
multiengine aeroplane, the provision of 4.2.2.1 through 4.2.2.4 or the construction of the liner itself provides such protection.
and 5.7.7 must be met. 5.2.1.5 A positive pressure must be maintained within the
vapor space of each bladder cell under any condition of
4.2.2.1 There shall be an independent tank outlet for each
engine. operation, except for a particular condition for which it is
shown that a zero or negative pressure will not cause the
4.2.2.2 There shall be a shut-off valve at each tank outlet.
This shutoff valve may also serve as the fire wall shutoff valve bladder cell to collapse.
required if the line between the valve and the engine compart- 5.2.1.6 Siphoning of fuel (other than minor spillage) or
ment does not contain more than 1 L [1 US qt] of fuel (or any collapse of bladder fuel cells may not result from improper
greater amount shown to be safe) that can escape into the securing or loss of the fuel filler cap.
engine compartment.
5.2.2 Fuel tanks must be designed, located, and installed so
4.2.2.3 At least two vents arranged to minimize the prob- as to retain fuel when subjected to the inertia loads resulting
ability of both vents becoming obstructed simultaneously must from the ultimate static load factors prescribed in Specification
be provided. F3083/F3083M; and under conditions likely to occur when the
F3063/F3063M − 21
aeroplane lands on a paved runway at a normal landing speed 5.5.2.1 Each vapor elimination connections and each vapor
under the conditions specified in 5.2.2.1 through 5.2.2.3. return provisions must have a separate vent line to lead vapors
5.2.2.1 The aeroplane is in a normal landing attitude and its back to the top of one of the fuel tanks.
landing gear is retracted.
5.5.2.2 If there is more than one tank and it is necessary to
5.2.2.2 The most critical landing gear leg is collapsed and
use these tanks in a definite sequence for any reason, the vapor
the other landing gear legs are extended.
vent line must lead back to the fuel tank to be used first, unless
5.2.2.3 In showing compliance with 5.2.2.1 and 5.2.2.2, the
the relative capacities of the tanks are such that return to
tearing away of an engine mount must be considered unless all
another tank is preferable.
the engines are installed above the wing or on the tail or
5.5.3 For aeroplanes approved for aerobatics, the require-
fuselage of the aeroplane.
ments in 5.5.3.1 through 5.5.3.2 must be prevented for each
5.2.3 Each integral fuel tank must have adequate facilities
acrobatic maneuver for which certification is requested.
for interior inspection and repair.
5.5.3.1 Excessive loss of fuel, including short periods of
5.2.4 For Level 4 aeroplanes, fuel tanks within the fuselage
inverted flight.
contour must be able to resist rupture and be in a protected
5.5.3.2 Siphoning of fuel from the vent when normal flight
position so that exposure of the tanks to scraping action with
has been resumed.
the ground is unlikely.
5.6 Sump:
5.3 Compartments:
5.6.1 Each fuel tank must have a drainable sump with an
5.3.1 Each tank compartment must be ventilated and
effective capacity, in the normal ground and flight attitudes, of
drained to prevent the accumulation of flammable fluids or
0.25 % of the tank capacity, or 0.24 L [ ⁄16 US gal], whichever
vapors.
is greater.
5.3.2 Eachcompartmentadjacenttoatankthatisanintegral
5.6.2 Each fuel tank must allow drainage of any hazardous
part of the aeroplane structure must also be ventilated and
quantity of water from any part of the tank to its sump with the
drained.
aeroplane in the normal ground attitude.
5.4 Expansion Space:
5.6.3 Each reciprocating engine fuel system must have a
5.4.1 Each fuel tank must have an expansion space of not
sump that meets the requirements of 5.6.3.1 through 5.6.3.3.
less than2%ofthe tank capacity, unless the tank vent
5.6.3.1 Have a sediment bowl or chamber that is accessible
discharges clear of the aeroplane (in which case no expansion
for drainage.
space is required).
5.6.3.2 Have a capacity of 30 cm [1 oz] for every 75.7 L
5.4.2 It must be impossible to fill the expansion space
[20 US gal] of fuel tank capacity.
inadvertently with the aeroplane in the normal ground attitude.
5.6.3.3 Each fuel tank outlet must be located so that, in the
5.5 Vents and Carburetor Vapor Vents:
normalflightattitude,waterwilldrainfromallpartsofthetank
5.5.1 Each fuel tank must be vented from the top part of the
except the sump to the sediment bowl or chamber.
expansion space.
5.5.1.1 Each vent outlet must be located and constructed in
5.7 Filler Connection:
a manner that minimizes the possibility of its being obstructed
5.7.1 Each fuel tank filler connection must be marked as
by ice or other foreign matter.
prescribed in Specification F3117/F3117M.
5.5.1.2 Each vent must be constructed to prevent siphoning
5.7.2 Fuel tank filler connections must be located outside
of fuel during normal operation.
the personnel compartment.
5.5.1.3 The venting capacity must allow the rapid relief of
5.7.3 Spilled fuel must be prevented from entering the fuel
excessive differences of pressure between the interior and
tank compartment or any part of the aeroplane other than the
exterior of the tank.
tank itself.
5.5.1.4 Airspaces of tanks with interconnected outlets must
5.7.4 Each filler cap must provide a fuel-tight seal for the
be interconnected.
main filler opening. However, there may be small openings in
5.5.1.5 There may be no point in any vent line where
the fuel tank cap for venting purposes or for the purpose of
moisture can accumulate with the aeroplane in either the
allowingpassageofafuelgaugethroughthecapprovidedsuch
ground or level flight attitudes, unless drainage is provided.
openings comply with the requirements of 5.5.1.
Any drain valve installed must be accessible for drainage.
5.7.5 Each fuel filling point, except pressure fueling con-
5.5.1.6 No vent may terminate at a point where the dis-
nection points, must have a provision for electrically bonding
charge of fuel from the vent outlet will constitute a fire hazard
the aeroplane to ground fueling equipment.
or from which fumes may enter personnel compartments.
5.7.6 Fuelfilleropeningsshouldbedesignedtoprecludethe
5.5.1.7 Vents must be arranged to prevent the loss of fuel,
use of fuels other than those approved for use.
except fuel discharged because of thermal expansion, when the
5.7.6.1 Fuel filler openings no larger than 60 mm [2.36 in.]
aeroplane is parked in any direction on a ramp havinga1%
are appropriate for aeroplanes with engines requiring gasoline
slope.
as the only permissible fuel.
5.5.2 Each carburetor with vapor elimination connections
and each fuel injection engine employing vapor return provi- 5.7.6.2 Fuelfilleropeningsnosmallerthan75mm[2.95in.]
sions must meet the conditions specified in 5.5.2.1 through are appropriate for aeroplanes with engines requiring turbine
5.5.2.2. fuel as the only permissible fuel.
F3063/F3063M − 21
5.7.7 For single fuel tanks on multiengine aeroplanes the 5.9.1.3 For each nonmetallic tank with walls supported by
filler caps should be designed to prevent inflight loss, incorrect the aeroplane structure and constructed in an acceptable
installation, or have means to indicate that the cap is not mannerusingacceptablebasictankmaterial,andwithactualor
properly installed. simulated support conditions, a pressure of 14 kPa [2 psi] for
thefirsttankofaspecificdesign.Thesupportingstructuremust
5.8 Strainers & Filters:
be designed for the critical loads occurring in the flight or
5.8.1 There must be a fuel strainer for the fuel tank outlet or
landing strength conditions combined with the fuel pressure
for the booster pump to prevent the passage of any object that
loads resulting from the corresponding accelerations.
could restrict fuel flow or damage any fuel system component.
5.9.2 For aeroplane with more than one engine or with more
5.8.1.1 For spark ignition engine powered aeroplanes, the
than two seats or with a maximum takeoff weight of more than
fuel strainer must have 3 meshes to 6 meshes per centimeter
750 kg [1650 lb] or a stall speed above 83 km/h [45 knots],
[8 meshes to 16 meshes per inch].
each fuel tank with large, unsupported, or unstiffened flat
5.8.1.2 The fuel strainer must either:
surfaces whose failure or deformation could cause fuel
(1) have a length of at least twice the diameter of the fuel
leakage, must be able to withstand the test defined in 5.9.2.1
tank outlet, or
through 5.9.2.2 without leakage, failure, or excessive deforma-
(2) haveaclearareaofeachfueltankoutletstraineratleast
tion of the tank walls:
five times the area of the outlet line.
5.9.2.1 Each complete tank assembly and its support must
5.8.1.3 The diameter of each strainer must be at least that of
be vibration tested under conditions that simulate the actual
the fuel tank outlet.
installation.
5.8.1.4 Each strainer must be accessible for inspection and
5.9.2.2 Except as specified in 5.9.3.4, the tank assembly
cleaning.
must be vibrated for 25 h at a total displacement of not less
5.8.2 There must be a fuel strainer or filter between the fuel
than 0.8 mm [ ⁄32 in.] (unless another displacement is substan-
tank outlet and the inlet of either the fuel metering device or an
tiated) while ⁄3 filled with water or other suitable test fluid.
engine driven positive displacement pump, whichever is nearer
5.9.3 For aeroplane with more than one engine or with more
the fuel tank outlet.
than two seats or with a maximum takeoff weight of more than
5.8.2.1 The fuel strainer or filter must be accessible for
750 kg [1650 lb] or a stall speed above 83 km/h [45 knots],
drainingandcleaningandmustincorporateascreenorelement
aeroplane the test frequency of vibration must meet the
which is easily removable.
requirements defined in 5.9.3.1 through 5.9.3.7:
5.8.2.2 The fuel strainer or filter must have a sediment trap
5.9.3.1 If no frequency of vibration resulting from any rpm
and drain except that it need not have a drain if the strainer or
withinthenormaloperatingrangeofengineorpropellerspeeds
filter is easily removable for drain purposes.
is critical, the test frequency of vibration is the number of
5.8.2.3 The fuel strainer or filter must be mounted so that its
cycles per minute obtained by multiplying the maximum
weight is not supported by the connecting lines or by the inlet
continuous propeller speed in rpm by 0.9 for propeller-driven
or outlet connections of the strainer or filter itself, unless
aeroplanes, and for non-propeller driven aeroplanes the test
adequate strength margins under all loading conditions are
frequency of vibration is 2000 cycles per minute.
provided in the lines and connections.
5.9.3.2 Ifonlyonefrequencyofvibrationresultingfromany
5.8.2.4 The fuel strainer or filter must have the capacity
rpm within the normal operating range of engine or propeller
(with respect to operating limitations established for the
speeds is critical, that frequency of vibration must be the test
engine) to ensure that engine fuel system functioning is not
frequency.
impaired, with the fuel contaminated to a degree (with respect
5.9.3.3 If more than one frequency of vibration resulting
to particle size and density) that is greater than that established
from any rpm within the normal operating range of engine or
for the engine during its type certification.
propeller speeds is critical, the most critical of these frequen-
5.8.3 For Level 4 aeroplanes, unless means are provided in
cies must be the test frequency.
the fuel system to prevent the accumulation of ice on the filter,
5.9.3.4 Under 5.9.3.2 and 5.9.3.3, the time of test must be
a means must be provided to automatically maintain the fuel
adjusted to accomplish the same number of vibration cycles
flow if ice clogging of the filter occurs.
that would be accomplished in 25 h at the frequency specified
5.9 Tests: in 5.9.3.1.
5.9.1 Each fuel tank must be able to withstand the pressures 5.9.3.5 During the test, the tank assembly must be rocked at
defined in 5.9.1.1 through 5.9.1.3 without failure or leakage. a rate of 16 to 20 complete cycles per minute, through an angle
5.9.1.1 For each conventional metal tank and nonmetallic of 15° on either side of the horizontal (30° total), about an axis
parallel to the axis of the fuselage, for 25 h.
tank with walls not supported by the aeroplane structure, a
pressure of 24 kPa [3.5 psi], or that pressure developed during 5.9.3.6 Each integral tank using methods of construction
maximum ultimate acceleration with a full tank, whichever is and sealing not previously proven to be adequate by test data
greater. or service experience must be able to withstand the vibration
test specified in 5.9.3.1 through 5.9.3.4.
5.9.1.2 For each integral tank, the pressure developed dur-
ing the maximum limit acceleration of the aeroplane with a full 5.9.3.7 Eachtankwithanonmetalliclinermustbesubjected
tank, with simultaneous application of the critical limit struc- to the sloshing test outlined in 5.9.3.5, with the fuel at room
tural loads. temperature. In addition, a specimen liner of the same basic
F3063/F3063M − 21
construction as that to be used in the aeroplane must, when most critical with respect to fuel feed and quantity of unusable
installed in a suitable test tank, withstand the sloshing test with fuel. These conditions may be simulated in a suitable mockup.
fuel at a temperature of 43 °C [110 °F].
7.1.1.1 The quantity of fuel in the tank may not exceed the
amount established as the unusable fuel supply for that tank
5.10 Unusable Fuel Supply:
under 5.10.1 plus that quantity necessary to show compliance
5.10.1 The unusable fuel supply for each tank must be
with this section.
established as not less than that quantity at which the first
7.1.1.2 If there is a fuel flowmeter, it must be blocked
evidence of malfunctioning occurs under the most adverse fuel
during the flow test and the fuel must flow through the meter
feed condition occurring under each intended operation and
or its bypass.
flight maneuver involving that tank. Fuel system component
7.1.1.3 If there is a flowmeter without a bypass, it must not
failures need not be considered.
have any probable failure mode that would restrict fuel flow
5.10.2 The effect on the usable fuel quantity as a result of a
below the level required for this fuel demonstration.
failure of any pump shall be determined.
7.1.1.4 The fuel flow must include that flow necessary for
5.11 Flow Between Interconnected Tanks:
vapor return flow, jet pump drive flow, and for all other
5.11.1 It must be impossible, in a gravity feed system with
purposes for which fuel is used.
interconnected tank outlets, for enough fuel to flow between
the tanks to cause an overflow of fuel from any tank vent under 7.2 Gravity Systems—The fuel flow rate for gravity systems
the conditions in 5.10, except that full tanks must be used. (main and reserve supply) must be 150 % of the takeoff fuel
consumption of the engine.
5.11.2 If fuel can be pumped from one tank to another in
flight, the fuel tank vents and the fuel transfer system must be
7.3 Pump Systems:
designed so that no structural damage to any aeroplane
7.3.1 The fuel flow rate for each pump system (main and
component can occur because of overfilling of any tank.
reserve supply) for each spark ignition reciprocating engine
mustbe125%andforeachcompressionignitionreciprocating
6. Fuel Pumps
engine 100 % of the fuel flow required by the engine at the
6.1 Main Pumps:
maximum takeoff power approved under this part.
6.1.1 For reciprocating engine installations, the main pump
7.3.2 The flow rate established in 7.3.1 is required for each
must be directly driven by the engine or meet the turbine
main pump and each emergency pump, and must be available
engine requirements found in 6.1.2.
when the pump is operating as it would during takeoff.
6.1.2 For turbine engine installations, the main pump must
7.3.3 For each hand-operated pump, the flow rate estab-
meet the requirements of 6.1.2.1 through 6.1.2.3.
lished in 7.3.1 must occur at not more than 60 complete cycles
6.1.2.1 There must be at least one main pump.
(120 single strokes) per minute.
6.1.2.2 The power supply for the main pump for each
7.3.4 The fuel pressure, with main and emergency pumps
engine must be independent of the power supply for each main
operating simultaneously, must not exceed the fuel inlet
pump for any other engine.
pressure limits of the engine unless it can be shown that no
6.1.2.3 Provision must be made to allow the bypass of each
adverse effect occurs.
positive displacement fuel pump other than a fuel injection
7.4 Auxiliary Fuel Systems and Fuel Transfer Systems:
pump approved as part of the engine.
7.4.1 Subsections 7.2, 7.3, and 7.6 apply to each auxiliary
6.2 Emergency Pump:
andtransfersystem,withtheexceptionlistedin7.4.1.1through
6.2.1 There must be an emergency pump immediately
7.4.1.2.
available to supply fuel to the engine if any main pump (other
7.4.1.1 The required fuel flow rate must be established upon
than a fuel injection pump approved as part of an engine) fails.
the basis of maximum continuous power and engine rotational
6.2.2 The power supply for each emergency pump must be
speed, instead of takeoff power and fuel consumption.
independent of the power supply for each corresponding main
7.4.1.2 If there is a placard providing operating instructions,
pump.
a lesser flow rate may be used for transferring fuel from any
6.3 If both the main pump and emergency pump operate
auxiliary tank into a larger main tank.
continuously, there must be a means to indicate to the appro-
(1) This lesser flow rate must be adequate to maintain
priate flight crewmembers a malfunction of either pump.
engine maximum continuous power but the flow rate must not
overfill the main tank at lower engine powers.
6.4 Operation of any fuel pump may not affect engine
operation so as to create a hazard, regardless of the engine
7.5 Multiple Fuel Tanks:
power or thrust setting or the functional status of any other fuel
7.5.1 For reciprocating engines that are supplied with fuel
pump.
from more than one tank, if engine power loss becomes
apparentduetofueldepletionfromthetankselected,itmustbe
7. Fuel Flow
possible after switching to any full tank, in level flight, to
7.1 General: obtain75%maximumcontinuouspowerorfullpowerandfuel
pressure for on that engine in not more than:
7.1.1 The ability of the fuel system to provide fuel at the
rates specified in this section and at a pressure sufficient for 7.5.1.1 Ten seconds (10 s) for naturally aspirated single-
proper engine operation must be shown in the attitude that is engine aeroplanes.
F3063/F3063M − 21
7.5.1.2 Twenty seconds (20 s) for turbocharged single- 8.2.1 Allow checking for proper shutoff operation before
engine aeroplanes, provided that 75 % maximum continuous each fueling of the tank.
naturally aspirated power is regained within 10 s.
8.2.2 Fo
...

This document is not an ASTM standard and is intended only to provide the user of an ASTM standard an indication of what changes have been made to the previous version. Because
it may not be technically possible to adequately depict all changes accurately, ASTM recommends that users consult prior editions as appropriate. In all cases only the current version
of the standard as published by ASTM is to be considered the official document.
Designation: F3063/F3063M − 20 F3063/F3063M − 21
Standard Specification for
Aircraft Fuel Storage and Delivery
This standard is issued under the fixed designation F3063/F3063M; the number immediately following the designation indicates the year
of original adoption or, in the case of revision, the year of last revision. A number in parentheses indicates the year of last reapproval.
A superscript epsilon (´) indicates an editorial change since the last revision or reapproval.
1. Scope
1.1 This specification covers minimum requirements for the design and integration of Fuel Storage and Delivery system
installations for aeroplanes.
1.2 This specification is applicable to aeroplanes as defined in the F44 terminology standard.
1.3 The applicant for a design approval must seek the individual guidance to their respective CAA body concerning the use of this
standard as part of a certification plan. For information on which CAA regulatory bodies have accepted this standard (in whole
or in part) as a means of compliance to their Aeroplane Airworthiness regulations (Hereinafter referred to as “the Rules”), refer
to ASTM F44 webpage (www.ASTM.org/COMITTEE/F44.htm)(www.ASTM.org/COMITTEE/F44.htm), which includes CAA
website links. Annex A1 maps the Means of Compliance described in this specification to EASA CS-23, amendment 5, or later,
and FAA 14 CFR Part 23, amendment 64, or later.
1.4 Units—The values stated are SI units followed by imperial units in brackets. The values stated in each system are not
necessarily exact equivalents; therefore, to ensure conformance with the standard, each system shall be used independently of the
other, and values from the two systems shall not be combined.
1.5 This standard does not purport to address all of the safety concerns, if any, associated with its use. It is the responsibility
of the user of this standard to establish appropriate safety, health, and environmental practices and determine the applicability of
regulatory limitations prior to use.
1.6 This international standard was developed in accordance with internationally recognized principles on standardization
established in the Decision on Principles for the Development of International Standards, Guides and Recommendations issued
by the World Trade Organization Technical Barriers to Trade (TBT) Committee.
2. Referenced Documents
2.1 ASTM Standards:
F3060 Terminology for Aircraft
F3083/F3083M Specification for Emergency Conditions, Occupant Safety and Accommodations
F3116/F3116M Specification for Design Loads and Conditions
F3117/F3117M Specification for Crew Interface in Aircraft
F3179/F3179M Specification for Performance of Aircraft
F3397/F3397M Practice for Aeroplane Turbine Fuel System Hot Weather Operations
This specification is under the jurisdiction of ASTM Committee F44 on General Aviation Aircraft and is the direct responsibility of Subcommittee F44.40 on Powerplant.
Current edition approved March 1, 2020Sept. 1, 2021. Published March 2020October 2021. Originally approved in 2015. Last previous edition approved in 20182020 as
F3063/F3063M–18a.–20. DOI: 10.1520/F3063_F3063M–20.10.1520/F3063_F3063M-21.
For referenced ASTM standards, visit the ASTM website, www.astm.org, or contact ASTM Customer Service at [e-mail]. For Annual Book of ASTM Standards
volume information, refer to the standard’s Document Summary page on the ASTM website.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
F3063/F3063M − 21
2.2 EASA Standards:
CS-23 Normal, Utility, Aerobatic and Commuter Aeroplanes
CS-34 Easy Access Rules for Aircraft Engine Emissions and Fuel Venting
2.3 Other Standard:FAA Standards:
US 14 CFR Part 23 Airworthiness Standards: Normal, Utility, Aerobatic and Commuter Normal Category Airplanes
(Amendment 62)
US 14 CFR Part 34 Fuel Venting and Exhaust Emission Requirements for Turbine Engine Powered Airplanes
3. Terminology
3.1 The following are a selection of relevant terms. See Terminology F3060 for more definitions and abbreviations.
3.2 Definitions:
3.2.1 main pump, n—a pump that supplies sufficient fuel to sustain the engine during normal operations.
3.2.2 emergency pump, n—a pump that can sustain engine operation at full power in the event the main pump fails.
3.2.3 auxiliary pump, n—a pump that can provide some fuel flow during emergency operations but not enough to sustain engine
operation at full power. Its function is to aid in priming the engine and suppressing fuel vapors.
3.3 Abbreviations:
3.3.1 cc—cubic centimetre
3.3.2 CFR—Code of Federal Regulations
3.3.3 RPM—rotation per minute
4. Fuel System
4.1 General:
4.1.1 Each fuel system must be constructed and arranged to ensure fuel flow at a rate and pressure established for proper engine
and auxiliary power unit functioning under each likely operating condition, including any maneuver for which certification is
requested and during which the engine or auxiliary power unit is permitted to be in operation.
4.1.2 Each fuel system for a turbine engine and compression ignition engine must be capable of sustained operation throughout
its flow and pressure range with fuel initially saturated with water at 27 °C [80 °F] and having 0.75 cc of free water per 3.8 L [1
US gal] added and cooled to the most critical condition for icing likely to be encountered in operation.
4.1.3 Each fuel system for a turbine engine powered aeroplane must meet the applicable fuel venting requirements of 14 CFR Part
34 for the US United States or the applicable fuel venting requirements as contained in the rules specified by the applicant’s local
civil aviation authority.
4.1.4 Each fuel system must be arranged so that it meets the requirement of 4.1.4.1 or 4.1.4.2.
4.1.4.1 No fuel pump can draw fuel from more than one tank at a time.
4.1.4.2 There are means to prevent introducing air into the system.
4.1.5 Fuel system components in an engine nacelle or in the fuselage must be protected from damage which could result in spillage
of enough fuel to constitute a fire hazard as a result of a wheels-up landing on a paved runway.
Available from U.S. Government Publishing Office (GPO), 732 N. Capitol St., NW, Washington, DC 20401, http://www.gpo.gov.European Union Aviation Safety Agency
(EASA), Konrad-Adenauer-Ufer 3, D-50668 Cologne, Germany, https://www.easa.europa.eu.
Available from Federal Aviation Administration (FAA), 800 Independence Ave., SW, Washington, DC 20591, http://www.faa.gov.
F3063/F3063M − 21
4.1.6 Each check valve must be constructed, or otherwise incorporate provisions, to preclude incorrect assembly or connection of
the valve.
4.2 Independence:
4.2.1 Each fuel system for a multiengine aeroplane must be arranged so that, in at least one system configuration, the failure of
any one component will not result in the loss of power of more than one engine or require immediate action by the pilot to prevent
the loss of power of more than one engine.
4.2.2 If a single fuel tank (or series of fuel tanks interconnected to function as a single fuel tank) is used on a multiengine
aeroplane, the provision of 4.2.2.1 through 4.2.2.4 and 5.7.7 must be met.
4.2.2.1 There shall be an independent tank outlet for each engine.
4.2.2.2 There shall be a shut-off valve at each tank outlet. This shutoff valve may also serve as the fire wall shutoff valve required
if the line between the valve and the engine compartment does not contain more than 1 L [1 US qt] of fuel (or any greater amount
shown to be safe) that can escape into the engine compartment.
4.2.2.3 At least two vents arranged to minimize the probability of both vents becoming obstructed simultaneously must be
provided.
4.2.2.4 A fuel system in which those parts of the system from each tank outlet to any engine are independent of each part of the
system supplying fuel to any other engine must be provided.
4.3 Drains:
4.3.1 There must be at least one drain to allow safe drainage of the entire fuel system with the aeroplane in its normal ground
attitude.
4.3.2 Each drain installed for the purpose of draining hazardous quantities of water, as required by 5.6, must meet the provision
of 4.3.2.1 through 4.3.2.2.
4.3.2.1 Discharge clear of all parts of the aeroplane.
4.3.2.2 Have a drain valve that meets 4.3.2.2(1) through (6).
(1) That has manual or automatic means for positive locking in the closed position.
(2) That is readily accessible.
(3) That can be easily opened and closed.
(4) That is either located or protected to prevent fuel spillage in the event of a landing with landing gear retracted.
(5) That allows the fuel to be caught for examination.
(6) That can be observed for proper closing.
5. Fuel Tanks
5.1 General:
5.1.1 Each fuel tank must be able to withstand, without failure, the vibration, inertia, fluid, and structural loads that it may be
subjected to in operation.
5.1.2 Each flexible fuel tank liner must be shown to be suitable for the particular application.
5.1.3 The total usable capacity of the fuel tanks must be enough for at least 30 min of operation at maximum continuous power.
5.2 Installation:
5.2.1 Each fuel tank must be supported so that tank loads are not concentrated.
F3063/F3063M − 21
5.2.1.1 There must be pads, if necessary, to prevent chafing between each tank and its supports.
5.2.1.2 Padding must be nonabsorbent or treated to prevent the absorption of fuel.
5.2.1.3 If a flexible tank liner is used, it must be supported so that it is not required to withstand fluid loads.
5.2.1.4 Interior surfaces adjacent to the liner must be smooth and free from projections that could cause wear, unless provisions
are made for protection of the liner at those points; or the construction of the liner itself provides such protection.
5.2.1.5 A positive pressure must be maintained within the vapor space of each bladder cell under any condition of operation,
except for a particular condition for which it is shown that a zero or negative pressure will not cause the bladder cell to collapse.
5.2.1.6 Siphoning of fuel (other than minor spillage) or collapse of bladder fuel cells may not result from improper securing or
loss of the fuel filler cap.
5.2.2 Fuel tanks must be designed, located, and installed so as to retain fuel when subjected to the inertia loads resulting from the
ultimate static load factors prescribed in Specification F3083/F3083M; and under conditions likely to occur when the aeroplane
lands on a paved runway at a normal landing speed under the conditions specified in 5.2.2.1 through 5.2.2.3.
5.2.2.1 The aeroplane is in a normal landing attitude and its landing gear is retracted.
5.2.2.2 The most critical landing gear leg is collapsed and the other landing gear legs are extended.
5.2.2.3 In showing compliance with 5.2.2.1 and 5.2.2.2, the tearing away of an engine mount must be considered unless all the
engines are installed above the wing or on the tail or fuselage of the aeroplane.
5.2.3 Each integral fuel tank must have adequate facilities for interior inspection and repair.
5.2.4 For Level 4 aeroplanes, fuel tanks within the fuselage contour must be able to resist rupture and be in a protected position
so that exposure of the tanks to scraping action with the ground is unlikely.
5.3 Compartments:
5.3.1 Each tank compartment must be ventilated and drained to prevent the accumulation of flammable fluids or vapors.
5.3.2 Each compartment adjacent to a tank that is an integral part of the aeroplane structure must also be ventilated and drained.
5.4 Expansion Space:
5.4.1 Each fuel tank must have an expansion space of not less than 2 % of the tank capacity, unless the tank vent discharges clear
of the aeroplane (in which case no expansion space is required).
5.4.2 It must be impossible to fill the expansion space inadvertently with the aeroplane in the normal ground attitude.
5.5 Vents and Carburetor Vapor Vents:
5.5.1 Each fuel tank must be vented from the top part of the expansion space.
5.5.1.1 Each vent outlet must be located and constructed in a manner that minimizes the possibility of its being obstructed by ice
or other foreign matter.
5.5.1.2 Each vent must be constructed to prevent siphoning of fuel during normal operation.
5.5.1.3 The venting capacity must allow the rapid relief of excessive differences of pressure between the interior and exterior of
the tank.
F3063/F3063M − 21
5.5.1.4 Airspaces of tanks with interconnected outlets must be interconnected.
5.5.1.5 There may be no point in any vent line where moisture can accumulate with the aeroplane in either the ground or level
flight attitudes, unless drainage is provided. Any drain valve installed must be accessible for drainage.
5.5.1.6 No vent may terminate at a point where the discharge of fuel from the vent outlet will constitute a fire hazard or from which
fumes may enter personnel compartments.
5.5.1.7 Vents must be arranged to prevent the loss of fuel, except fuel discharged because of thermal expansion, when the
aeroplane is parked in any direction on a ramp having a 1 % slope.
5.5.2 Each carburetor with vapor elimination connections and each fuel injection engine employing vapor return provisions must
meet the conditions specified in 5.5.2.1 through 5.5.2.2.
5.5.2.1 Each vapor elimination connections and each vapor return provisions must have a separate vent line to lead vapors back
to the top of one of the fuel tanks.
5.5.2.2 If there is more than one tank and it is necessary to use these tanks in a definite sequence for any reason, the vapor vent
line must lead back to the fuel tank to be used first, unless the relative capacities of the tanks are such that return to another tank
is preferable.
5.5.3 For aeroplanes approved for aerobatics, the requirements in 5.5.3.1 through 5.5.3.2 must be prevented for each acrobatic
maneuver for which certification is requested.
5.5.3.1 Excessive loss of fuel, including short periods of inverted flight.
5.5.3.2 Siphoning of fuel from the vent when normal flight has been resumed.
5.6 Sump:
5.6.1 Each fuel tank must have a drainable sump with an effective capacity, in the normal ground and flight attitudes, of 0.25 %
of the tank capacity, or 0.24 L [ ⁄16 US gal], whichever is greater.
5.6.2 Each fuel tank must allow drainage of any hazardous quantity of water from any part of the tank to its sump with the
aeroplane in the normal ground attitude.
5.6.3 Each reciprocating engine fuel system must have a sump that meets the requirements of 5.6.3.1 through 5.6.3.3.
5.6.3.1 Have a sediment bowl or chamber that is accessible for drainage.
5.6.3.2 Have a capacity of 30 cm [1 oz] for every 75.7 L [20 US gal] of fuel tank capacity.
5.6.3.3 Each fuel tank outlet must be located so that, in the normal flight attitude, water will drain from all parts of the tank except
the sump to the sediment bowl or chamber.
5.7 Filler Connection:
5.7.1 Each fuel tank filler connection must be marked as prescribed in Specification F3117/F3117M.
5.7.2 Fuel tank filler connections must be located outside the personnel compartment.
5.7.3 Spilled fuel must be prevented from entering the fuel tank compartment or any part of the aeroplane other than the tank itself.
5.7.4 Each filler cap must provide a fuel-tight seal for the main filler opening. However, there may be small openings in the fuel
tank cap for venting purposes or for the purpose of allowing passage of a fuel gauge through the cap provided such openings
comply with the requirements of 5.5.1.
F3063/F3063M − 21
5.7.5 Each fuel filling point, except pressure fueling connection points, must have a provision for electrically bonding the
aeroplane to ground fueling equipment.
5.7.6 Fuel filler openings should be designed to preclude the use of fuels other than those approved for use.
5.7.6.1 Fuel filler openings no larger than 60 mm [2.36 in.] are appropriate for aeroplanes with engines requiring gasoline as the
only permissible fuel.
5.7.6.2 Fuel filler openings no smaller than 75 mm [2.95 in.] are appropriate for aeroplanes with engines requiring turbine fuel
as the only permissible fuel.
5.7.7 For single fuel tanks on multiengine aeroplanes the filler caps should be designed to prevent inflight loss, incorrect
installation, or have means to indicate that the cap is not properly installed.
5.8 Strainers & Filters:
5.8.1 There must be a fuel strainer for the fuel tank outlet or for the booster pump to prevent the passage of any object that could
restrict fuel flow or damage any fuel system component.
5.8.1.1 For spark ignition engine powered aeroplanes, the fuel strainer must have 3 meshes to 6 meshes per centimeter [8 meshes
[8 meshes to 16 meshes per inch].
5.8.1.2 The fuel strainer must either:
(1) have a length of at least twice the diameter of the fuel tank outlet, or
(2) have a clear area of each fuel tank outlet strainer at least five times the area of the outlet line.
5.8.1.3 The diameter of each strainer must be at least that of the fuel tank outlet.
5.8.1.4 Each strainer must be accessible for inspection and cleaning.
5.8.2 There must be a fuel strainer or filter between the fuel tank outlet and the inlet of either the fuel metering device or an engine
driven positive displacement pump, whichever is nearer the fuel tank outlet.
5.8.2.1 The fuel strainer or filter must be accessible for draining and cleaning and must incorporate a screen or element which is
easily removable.
5.8.2.2 The fuel strainer or filter 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.
5.8.2.3 The fuel strainer or filter 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 itself, unless adequate strength margins under all loading conditions are provided in the
lines and connections.
5.8.2.4 The fuel strainer or filter must have the capacity (with respect to operating limitations established for the engine) to ensure
that engine fuel system functioning is not impaired, with the fuel contaminated to a degree (with respect to particle size and
density) that is greater than that established for the engine during its type certification.
5.8.3 For Level 4 aeroplanes, unless means are provided in the fuel system to prevent the accumulation of ice on the filter, a means
must be provided to automatically maintain the fuel flow if ice clogging of the filter occurs.
5.9 Tests:
5.9.1 Each fuel tank must be able to withstand the pressures defined in 5.9.1.1 through 5.9.1.3 without failure or leakage.
5.9.1.1 For each conventional metal tank and nonmetallic tank with walls not supported by the aeroplane structure, a pressure of
24 kPa [3.5 psi], or that pressure developed during maximum ultimate acceleration with a full tank, whichever is greater.
F3063/F3063M − 21
5.9.1.2 For each integral tank, the pressure developed during the maximum limit acceleration of the aeroplane with a full tank,
with simultaneous application of the critical limit structural loads.
5.9.1.3 For each nonmetallic tank with walls supported by the aeroplane structure and constructed in an acceptable manner using
acceptable basic tank material, and with actual or simulated support conditions, a pressure of 14 kPa [2 psi] for the first tank of
a specific design. The supporting structure must be designed for the critical loads occurring in the flight or landing strength
conditions combined with the fuel pressure loads resulting from the corresponding accelerations.
5.9.2 For aeroplane with more than one engine or with more than two seats or with a maximum takeoff weight of more than 750
kg [1650 lb] or a stall speed above 83 km/h [45 knots], each fuel tank with large, unsupported, or unstiffened flat surfaces whose
failure or deformation could cause fuel leakage, must be able to withstand the test defined in 5.9.2.1 through 5.9.2.2 without
leakage, failure, or excessive deformation of the tank walls:
5.9.2.1 Each complete tank assembly and its support must be vibration tested under conditions that simulate the actual installation.
5.9.2.2 Except as specified in 5.9.3.4, the tank assembly must be vibrated for 25 h at a total displacement of not less than 0.8 mm
1 2
[ ⁄32 in.] (unless another displacement is substantiated) while ⁄3 filled with water or other suitable test fluid.
5.9.3 For aeroplane with more than one engine or with more than two seats or with a maximum takeoff weight of more than 750
kg [1650 lb] or a stall speed above 83 km/h [45 knots], aeroplane the test frequency of vibration must meet the requirements
defined in 5.9.3.1 through 5.9.3.7:
5.9.3.1 If no frequency of vibration resulting from any rpm within the normal operating range of engine or propeller speeds is
critical, the test frequency of vibration is the number of cycles per minute obtained by multiplying the maximum continuous
propeller speed in rpm by 0.9 for propeller-driven aeroplanes, and for non-propeller driven aeroplanes the test frequency of
vibration is 2000 cycles per minute.
5.9.3.2 If only one frequency of vibration resulting from any rpm within the normal operating range of engine or propeller speeds
is critical, that frequency of vibration must be the test frequency.
5.9.3.3 If more than one frequency of vibration resulting from any rpm within the normal operating range of engine or propeller
speeds is critical, the most critical of these frequencies must be the test frequency.
5.9.3.4 Under 5.9.3.2 and 5.9.3.3, the time of test must be adjusted to accomplish the same number of vibration cycles that would
be accomplished in 25 h at the frequency specified in 5.9.3.1.
5.9.3.5 During the test, the tank assembly must be rocked at a rate of 16 to 20 complete cycles per minute, through an angle of
15° on either side of the horizontal (30° total), about an axis parallel to the axis of the fuselage, for 25 h.
5.9.3.6 Each integral tank using methods of construction and sealing not previously proven to be adequate by test data or service
experience must be able to withstand the vibration test specified in 5.9.3.1 through 5.9.3.4.
5.9.3.7 Each tank with a nonmetallic liner must be subjected to the sloshing test outlined in 5.9.3.5, with the fuel at room
temperature. In addition, a specimen liner of the same basic construction as that to be used in the aeroplane must, when installed
in a suitable test tank, withstand the sloshing test with fuel at a temperature of 43 °C [110 °F].
5.10 Unusable Fuel Supply:
5.10.1 The unusable fuel supply for each tank must be established as not less than that quantity at which the first evidence of
malfunctioning occurs under the most adverse fuel feed condition occurring under each intended operation and flight maneuver
involving that tank. Fuel system component failures need not be considered.
5.10.2 The effect on the usable fuel quantity as a result of a failure of any pump shall be determined.
5.11 Flow Between Interconnected Tanks:
F3063/F3063M − 21
5.11.1 It must be impossible, in a gravity feed system with interconnected tank outlets, for enough fuel to flow between the tanks
to cause an overflow of fuel from any tank vent under the conditions in 5.10, except that full tanks must be used.
5.11.2 If fuel can be pumped from one tank to another in flight, the fuel tank vents and the fuel transfer system must be designed
so that no structural damage to any aeroplane component can occur because of overfilling of any tank.
6. Fuel Pumps
6.1 Main Pumps:
6.1.1 For reciprocating engine installations, the main pump must be directly driven by the engine or meet the turbine engine
requirements found in 6.1.2.
6.1.2 For turbine engine installations, the main pump must meet the requirements of 6.1.2.1 through 6.1.2.3.
6.1.2.1 There must be at least one main pump.
6.1.2.2 The power supply for the main pump for each engine must be independent of the power supply for each main pump for
any other engine.
6.1.2.3 Provision must be made to allow the bypass of each positive displacement fuel pump other than a fuel injection pump
approved as part of the engine.
6.2 Emergency Pump:
6.2.1 There must be an emergency pump immediately available to supply fuel to the engine if any main pump (other than a fuel
injection pump approved as part of an engine) fails.
6.2.2 The power supply for each emergency pump must be independent of the power supply for each corresponding main pump.
6.3 If both the main pump and emergency pump operate continuously, there must be a means to indicate to the appropriate flight
crewmembers a malfunction of either pump.
6.4 Operation of any fuel pump may not affect engine operation so as to create a hazard, regardless of the engine power or thrust
setting or the functional status of any other fuel pump.
7. Fuel Flow
7.1 General:
7.1.1 The ability of the fuel system to provide fuel at the rates specified in this section and at a pressure sufficient for proper engine
operation must be shown in the attitude that is most critical with respect to fuel feed and quantity of unusable fuel. These conditions
may be simulated in a suitable mockup.
7.1.1.1 The quantity of fuel in the tank may not exceed the amount established as the unusable fuel supply for that tank under
5.10.1 plus that quantity necessary to show compliance with this section.
7.1.1.2 If there is a fuel flowmeter, it must be blocked during the flow test and the fuel must flow through the meter or its bypass.
7.1.1.3 If there is a flowmeter without a bypass, it must not have any probable failure mode that would restrict fuel flow below
the level required for this fuel demonstration.
7.1.1.4 The fuel flow must include that flow necessary for vapor return flow, jet pump drive flow, and for all other purposes for
which fuel is used.
F3063/F3063M − 21
7.2 Gravity Systems—The fuel flow rate for gravity systems (main and reserve supply) must be 150 % of the takeoff fuel
consumption of the engine.
7.3 Pump Systems:
7.3.1 The fuel flow rate for each pump system (main and reserve supply) for each spark ignition reciprocating engine must be 125
% and for each compression ignition reciprocating engine 100 % of the fuel flow required by the engine at the maximum takeoff
power approved under this part.
7.3.2 The flow rate established in 7.3.1 is required for each main pump and each emergency pump, and must be available when
the pump is operating as it would during takeoff.
7.3.3 For each hand-operated pump, the flow rate established in 7.3.1 must occur at not more than 60 complete cycles (120 single
strokes) per minute.
7.3.4 The fuel pressure, with main and emergency pumps operating simultaneously, must not exceed the fuel inlet pressure limits
of the engine unless it can be shown that no adverse effect occurs.
7.4 Auxiliary Fuel Systems and Fuel Transfer Systems:
7.4.1 Subsections 7.2, 7.3, and 7.6 apply to each auxiliary and transfer system, with the exception listed in 7.4.1.1 through 7.4.1.2.
7.4.1.1 The required fuel flow rate must be established upon the basis of maximum continuous power and engine rotational speed,
instead of takeoff power and fuel consumption.
7.4.1.2 If there is a placard providing operating instructions, a lesser flow rate may be used for transferring fuel from any auxiliary
tank into a larger main tank.
(1) This lesser flow rate must be adequate to maintain engine maximum continuous power but the flow rate must not overfill
the main tank at lower engine powers.
7.5 Multiple Fuel Tanks:
7.5.1 For reciprocating engines that are supplied with fuel from more than one tank, if engine power loss becomes apparent due
to fuel depletion from the tank selected, it must be possible after switching to any full tank, in level flight, to obtain 75 % maximum
continuous power or full power and fuel pressure for on that engine in not more than:
7.5.1.1 Ten seconds (10 s) for naturally aspirated single-engine
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