АС 27-1A 7/30/97 =°/'FUEL FLOW, (W,) LIMIT - L
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Q LIMIT W¿' TORQUE W¡"
MGT LIMIT OAT -→ FIGURE 450-1. FUEL FLOW NOTES:
(1) Point A on Figure 450-1 is the highest fuel flow within aircraft limitations, but the system pressure drop is not expected to be maximum because of the low kinematic fuel viscosity.
(2) Point B is the maximum flow at cold temperatures but as the fuel temperature is further reduced, the fuel viscosity increases very rapidly.
(3) Point C represents the maximum viscosity of the fuel, but the fuel flow is somewhat reduced from point B. The maximum system pressure drops and, therefore, minimum fuel inlet pressure may occur between points B and C depending on the specific relationship of fuel viscosity to required fuel flow.
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(10) A conservative demonstration would consider the maximum allowable fuel viscosity in combination with the maximum fuel flow. Otherwise, several test points may be required.
(11) For those systems which specify a minimum V/L ratio, the methods provided in Aerospace Recommended Practice (ARP) 492 published by the Society of Automotive Engineers are acceptable in evaluating test results.
(12) Since the lower quantity of fuel in the tank will reduce the hydrostatic head and thus the fuel inlet pressure, § 27.955(a)(2) specifies that the quantity of fuel in the tank should be minimum.
(13) Section 27.955(a)(4) specifies that each main and emergency pump be evaluated. If it can be determined which pump and flow path is critical, only that configuration would be tested. Similarly, for suction fuel systems, the critical flow paths and flow requirements should be evaluated. If pumps are required to supply the necessary fuel, § 27.1305(c) would require a fuel pressure indicator and § 27.1549 would require a red radial at the minimum safe operating fuel pressure for any fuel or fuel usage condition. This pressure limit should be used to determine compliance with § 27.955(a)(1) for all operations.
(14) Section 27.955(b) requires the fuel system to feed promptly when fuel becomes low in one tank and another tank is selected. This requirement is important because momentary fuel flow interruption must be expected to result in complete power failure and, for single engine rotorcraft, an emergency landing.
450A. § 27.955 (Amendment 27-23) FUEL FLOW.
a. Explanation. Amendment 27-23 adds new requirements for test conditions to ensure that adequate fuel flow is available to the engine in critical combinations of adverse conditions that may be expected during operation of the rotorcraft. The amendment also requires a correlation between fuel filter blockage and the fuel filter warning device required by § 27.1305(q). Design and performance standards for auxiliary fuel tank and transfer tank fuel systems are provided. These changes were made to ensure that all parameters associated with fuel supply to the engine are adequately addressed.
b. Procedures. All of the policy material pertaining to this section remains in effect with the following additions:
(1) Section 27.955 is intended to ensure adequate fuel flow to the engine(s) during all operating conditions of the rotorcraft. This includes the fuel flows necessary to operate the engine(s) under the test conditions required by § 27.927. Testing (including bench or rig tests) has been the accepted method of showing compliance with this section although analytical techniques may be used to adjust system test Par 450 601
results to various fuel flow conditions or to account for minor modifications to a system.
Analytical methods that are used to adjust the test results should be verified with limited testing. It should be shown during compliance testing that the fuel pressure, at the engine to airframe interface, will be within the limits specified by the engine manufacturer. The fuel pressure at this point should be maintained within limits specified by the engine manufacturer during all critical maneuvers and accelerations.
All of the following conditions should be met during compliance testing unless it can be shown that combinations of the conditions are not possible.
(i) The fuel quantity in the tank(s) in use during the test may not exceed the unusable fuel quantity established under § 27.959, plus the minimum quantity required to conduct the test.
During the compliance test, the rotorcraft should be maneuvered to create the most critical fuel pressure head between the fuel tank outlet and the engine to airframe interface (engine fuel inlet).
(ili) For boost pump fed systems, it should be determined which pump (primary or secondary) would create the most critical restriction if it failed. The critical pump should then be installed to create the critical restriction, either by actual or simulated failure.
(iv) Various combinations of engine power demand, electrical power available, and motive flow requirements for ejector pumps, will have an effect upon the fuel flow and pressure available at the engine to airframe interface. Adequate fuel pressure should be available to the engine with the most critical combination of these parameters.
(v) Critical values of fuel properties that may adversely affect fuel flow and/or fuel pressure should be applied. This includes alternate types of fuel if certification with alternate fuels is requested. At the minimum, the fuel that will create the highest vapor to liquid ratio should be used during hot fuel tests (§ 27.961). The most viscous fuel should be used during cold fuel tests.
(vi) The fuel filter, required by § 27.997, should be partially blocked to simulate the maximum contamination allowable. The blockage should be sufficient to activate the impending bypass indicator that is required by § 27.1305(q).
(2) Unique Conditions. The phrase," '...Provide the engine with at least 100 percent of the fuel required under all operating and maneuvering conditions...
(§ 27.955(a)), includes unique flight conditions within the operational envelope of the rotorcraft. Critical conditions of fuel flow to the engine(s) may exist under the following conditions (and others identified by the applicant); therefore, they should be evaluated and tested if applicable:
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In a single engine rotorcraft, a rapid acceleration to maximum power (torque) that will be requested for certification may be a critical condition. In this case the fuel flow required during the transient may exceed the fuel flow required for steady state at the maximum power condition.
(ii) In multiengine rotorcraft, a rapid acceleration to the maximum OEI power rating that will be requested may be a critical condition. The fuel flow during the transient may be higher than that required at the steady state OEI condition.
(3) If auxiliary fuel pumps (boost pumps) are used to supply fuel to the engines, and ejector pumps are used for cross-feed or other inter-tank fuel distribution systems, a test should be run that will place the maximum fuel demand on the auxiliary pump(s).
(4) In some multiengine rotorcraft, a single pump may be required to provide fuel flow to all engines in the event of an auxiliary pump failure. If this is the case, a test should be conducted with a simulated (or actual) failed auxiliary pump. If the functional auxiliary pump is designed to provide motive flow for cross-feed systems, the most critical condition of fuel flow demand should be tested.
(5) Transient and steady state maneuver loads (g-loading) may affect the fuel pressure at the engine to airframe interface. This effect should be considered and then tested, if appropriate.
(6) The methods of specifying the engine inlet fuel pressure requirements are sometimes related to fuel temperature and altitude. Therefore, it is necessary to explore the extremes of the envelope to assure compliance rather than attempting to select one critical condition. For instance, the increase in fuel viscosity at cold temperatures may increase system pressure drop and offset a slight drop in required fuel flow. In this case, critical tuel inlet conditions may not be experienced at maximum engine fuel flow.
(7) A conservative demonstration would consider the maximum allowable fuel viscosity in combination with the maximum fuel flow. Otherwise, several test points may be required.
(8) Euel Transfer Systems. Section 27.955(b) specifies that if normal operation of the rotorcraft fuel transfer system continually delivers fuel to an engine feed tank, and maintains a specific fuel level in the feed tank, then the specified fuel level in the feed tank should be maintained automatically during all flight or surface operating conditions expected with the rotorcraft.
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451. RESERVED.
Intentionally Left Blank Par 451 604
452. § 27.959 UNUSABLE FUEL SUPPLY.
a. Explanation. This rule requires the applicant to establish a value for unusable fuel for each tank. This value for unusable fuel may be selected by the applicant to facilitate compliance with § 27.1337(b)(1) provided the amount is equal to or greater than the actual unusable fuel. The actual unusable fuel is the amount of fuel in the tank when, in the critical flight attitude, evidence of system or engine malfunction occurs or, in the case of transfer tanks, when flow to the receiving tank is interrupted.
b. Procedures.
(1) The unusable fuel for each tank can be determined by flight tests which involve flight in the critical stable attitude and during maneuvers until indication of a malfunction. Maneuvers should be conducted to be critical or conservative with respect to unusable fuel. For boosted systems, the "first evidence of malfunction" may be a pressure fluctuation to below the fuel pressure minimum redline, engine power fluctuation, or boost pump failure warning indication. For suction lift systems, the indication may be engine power interruption. Since an accurate measurement of the remaining fuel in the tank should be obtained, a method to close off flow from that tank would be needed. For transfer tanks, or tanks which are limited to use only during cruise flight, the flight regimes usually can be limited to level flight at the CG condition which, by inspection, would create the maximum unusable fuel. For tanks for general use, the flight regimes should also include takeoff and landing using pitch attitudes to be expected, as well as hover and level flight conditions. The possible adverse effects of extreme lateral CG should be considered.
(2) Normally, these tests are conducted with all equipment (pumps, ejectors, etc.) operating as prescribed by the design. However, values for unusable fuel with pump failures, if significantly different, should also be determined and listed in the flight manual. The value for unusable fuel to be considered in the empty weight of the aircraft should be that value determined with the pumps) operating normally; i.e., pump failure need not be considered.
c. While the procedures of Paragraph (b)(1) are acceptable, fuel exhaustion during critical flight test conditions must be expected. To minimize this possible flight test hazard, the applicant may, in many cases, utilize analysis and/or ground tests involving normally available flight test data on aircraft attitudes, tank configuration studies, and critical flight condition studies to determine unusable fuel. Any questionable results, however, should be resolved by actual flight test or introduction of conservatism into the finding.
453. § 27.961 FUEL SYSTEM HOT WEATHER OPERATION.
a. Explanation.
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(1) Section 27.961 specifies that a hot fuel test be conducted on suction lift systems, and on other fuel systems conducive to vapor formation, to ensure that the system is free from vapor lock at a fuel temperature of 110° F under critical operating conditions.
(2) Pressure boosted systems would not ordinarily require hot fuel tests unless- (i> There are high points in the fuel system which would allow accumulation of vapor; or (II) The engine fuel inlet pressure is negative relative to tank pressure because of low boost pump pressure or high fuel system pressure losses (but still within fuel pressure limits).
(3) The requirement to use 110° F fuel is a carryover from the recodification of CAR Part 6, although the use of hotter fuel would tend more toward vapor formation.
(4) The term "vapor lock" means a change in normal engine operation as a result of the formation of fuel vapor-air mixtures in the fuel feed system.
b. Procedures.
(1) The fuel type to be used should be that with the highest true vapor pressure (TVP) at the 110° F condition.
(2) The fuel should be heated as rapidly as possible since the longer fuel is heated the more vaporization occurs resulting in unconservative test results.
(3) If the test is performed at cool ambients, the fuel lines, tanks, etc., may have to be insulated to ensure that the fuel inlet temperature is approximately the same as would be experienced on a hot day.
(4) The fuel level should be the lowest consistent with test safety.
(5) The flight tests to the service ceiling should include maximum power climbs to selected intermediate altitudes where various maneuvers including the following are performed:
(i> Low power descent with rapid transition to takeoff power.
(ii) Turns and cyclic pull-ups with load factors comparable to the flight strain survey.
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(ili) For multiengine rotorcraft with 30-minute and/or 2.5-minute OEI power ratings, conduct a rapid single-engine acceleration from low power to engine topping power followed by cruise at the maximum allowable OEI power.
(6) The flight test maneuvers should be repeated at the service ceiling.
(7) Except for transients and descents, the power available used should correspond to a 100° F sea level day lapsed 3.6° F/1,000 foot pressure altitude.
(8) Engine operation throughout the test should be normal; i.e., no surge, stall, flameout, etc., and the engine fuel inlet requirements should not be exceeded.
(9) Alternative tests on appropriate test rigs may be conducted ensuring proper simulation of altitude, ambient temperature, fuel temperature, fuel flow, and load factors.
453A. § 27.961 (Amendment 27-23) FUEL SYSTEM HOT WEATHER OPERATION
a. Explanation. Amendment 27-23 simplifies and restates the fuel system hot weather certification requirements and adds a requirement for the system to be capable of providing adequate fuel during probable transients. These changes clarified the existing wording to assure adequate qualification testing.
b. Procedures. This paragraph specifies that all suction lift systems and any other fuel system that may be conducive to vapor formation show satisfactory engine fuel inlet conditions (within criteria established by the engine manufacturer) when using the fuel with the highest true vapor pressure (TVP) at 110° F fuel temperature. Engine operating conditions should include those defined by §§ 27.927(b)(1) and 27.927(b)(2).
Compliance can be shown by analysis, testing, or a combination of both.
454. § 27.963 FUEL TANKS: GENERAL.
a. Explanation.
(1) Paragraph (a) sets forth general requirements for fuel tank structural aspects.
(2) Paragraph (b) requires design features to react forces to be expected from fuel surging due to accelerations of the rotorcraft.
(3) Paragraph (c) requires design features to ensure heat transfer from an engine compartment fire will not jeopardize the fuel tank integrity.
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AC 27-1A 7130/97
(4) Paragraph (d) requires design features to minimize the hazards of a leaking fuel tank and also requires design features to ensure that unwanted transfer of fuel from one tank to another does not occur due to differences of pressure in the tanks.
b. Procedures.
(1) For Paragraph (a), the tests of § 27.965 are normally adequate if performed in conjunction with the reliability test of § 21.35 or other service simulation tests.
(2) For Paragraph (b), internal or external stiffening may be required for surge resistance. If the analysis provided to show the adequacy of the surge resistance is questionable, the slosh and vibration tests of § 27.965 may be accepted as substantiation of this requirement.
(3) The fuel tank clearance required by Paragraph (c) may be determined by inspection of the design.
(4) The ventilation and interconnect requirements of Paragraph (d) may usually be determined by flight tests which explore maximum rates of climb and descent with sensitive pressure measuring equipment installed inside tanks and in the ventilation airspaces provided to comply with this rule.
454A. § 27.963 (Amendment 27-23) FUEL TANKS: GENERAL.
a. Explanation. Amendment 27-23 added new subsections (e) and (f) that require designs and tests to ensure that no exposed surface inside a fuel tank would, under normal or malfunction conditions, constitute an ignition source. They also set forth standards for the design and qualification of fuel tanks located in personnel compartments. These requirements are needed to ensure freedom from the hazards of fuel tank internal explosions and to ensure that fuel tanks, installed in passenger compartments, present no hazards to the personnel or to the rotorcraft.
b. Procedures. Section 27.963(e) requires the temperature of any exposed surface inside a fuel tank to be at least 50° F lower than the lowest auto-ignition temperature of the fuel or fuel vapors in the tank (reference AC Paragraph 586b(3), § 27.1185). For compliance with § 27.963(e), the internal component surface temperatures can be determined by flight or laboratory tests. The most critical flight conditions are established with sensitive temperature and pressure measuring equipment. This equipment is installed inside the tanks and in the ventilation air spaces.
454B. § 27.963 (Amendment 27-30) FUEL TANKS: GENERAL.
a. Explanation. Amendment 27-30 adds a new requirement to Paragraph (g) that (in addition to the current requirements) requires that the fuel tank bladder or liner be 608 Par 454
7130/97 AC 27-1A puncture resistant by meeting the TSO C-80, Paragraph 16.0, screwdriver test requirements, using a new crash resistance based minimum puncture force of 370 Ibs.
A new requirement is also added to Paragraph (f). Paragraph (f) now additionally requires that each fuel tank installed in a personnel compartment be crash resistant by meeting the applicable criteria of the new Crash Resistant Fuel System requirements of § 29.952 (Re: Paragraph 447).
b. Procedures.
(1) Paragraph (g). The procedures for old Paragraph (g) still apply under new
(g). In addition, to comply with the added puncture resistance requirement under new
(g), the requirements of § 27.952(h) must be met. Paragraph 447 of this document gives the detailed compliance procedures for § 27.952(h). The compliance procedures for § 27.952(h) also provide compliance for puncture resistance under § 27.963(g).
(2) Paragraph (f). The procedures for old Paragraph (f) still apply under new (f).
Compliance with the added crash resistance requirement of new (t) can be shown by conducting a thorough design review of each fuel tank compartment to ensure that all the regulatory criteria are met. (All fuel drains and vents should also be reviewed to ensure that they meet applicable § 27.952 requirements.) A basic static loads analysis followed by a stress analysis is typically used to determine that the enclosure protects the fuel tank and provides the crash resistance level necessary for occupant survival in an otherwise survivable impact. The applicable emergency load factors are typically used to design the enclosure. (Section 27.952 contains the corresponding load factors for fuel cells and their attachments.) The emergency load factors are typically adequate for all loading conditions encountered by the enclosure in service. The typical design approach is to design the enclosure to crush at a rate approximately the same as the crush rate of the fuel tank and to ensure that all puncture hazards (such as sharp projections either enhanced or created by impact that would penetrate the fuel tank) are minimized in design. (See Paragraph 447 guidance material for details.)
455. § 27.965 (Amendment 27-12) FUEL TANK TESTS.
a. Explanation. This regulation defines the tests that must be accomplished to show compliance for rotorcraft fuel tanks.
(1) Four basic types of fuel tanks are: (1) a metal tank installed in the aircraft or at the wing tip; (2) an integral tank; (3) a nonmetallic self-supporting tank (fiberglass);
and (4) nonmetallic flexible bladder-type tanks.
(2) There are two basic tests required by the regulations. One test procedure substantiates the design by tests and analysis by applying applicable pressure to the tank. The other procedure substantiates the design by vibration and slosh tests of the tanks.
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b. Procedures.
(1) Pressure Test. The 3.5 or 2.0 PSI pressure test listed in the regulations should be conducted unless the pressure with a full tank for maximum limit acceleration or emergency acceleration is greater. Section 27.337 gives the value for the limit acceleration.
(2) Vibration and Slosh Tests.
There is not an absolute value of what constitutes "large" unsupported or unstiffened flat areas. However, it has generally been considered that any fuel tank with less than 10 gallons capacity, constructed with simple, wide, flat geometric shape and using metal (in metal tanks) of 0.05-inch thickness or greater would not require tests in accordance with § 27.965(d). Using this basis, a 14- by 14-inch properly constructed tank would not require vibration and slosh tests.
(it) If the tank construction is of a metal or integral design which can be shown to be similar to previously approved tanks with acceptable service history, the vibration and slosh tests may not be required. Similarity would entail comparing the construction technique; i.e., similar panel size, similar sealing methods, skin and angle thickness, loads being similar, etc.
For fuel tanks located in the sponson or stub wing, the entire sponson or wing should be rocked and vibrated unless it can be determined that a certain portion of the tanks is critical. In this case a fixture should be developed such that the portion of the tank being tested is rocked about a pivot point which would produce the same amplitudes of motion for the portion of the tank being tested, as if the whole sponson or wing was being tested. Structure loads in conjunction with these tests have not been required.
(iv) The amplitude of vibration specified in the regulation is double amplitude (peak to peak). Vibration amplitudes less than one thirty-second of an inch must be justified by instrumented tests of the tank installed in the aircraft.
(v) The vibration and slosh procedures listed in Military Specification, MIL-T-6396, have been accepted to show compliance with § 27.965(d).
(3) After all tests have been conducted, the tanks should be leak checked using test fluid conforming to Federal Specification TT-S-735 type III or equivalent.
456. § 27.967 FUEL TANK INSTALLATION.
a. $ 27,967 (a):
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(1) Explanation. This paragraph was added by Amendment 27-30 to create parallelism of both regulatory structure and level of safety between Parts 27 and 29 by the Crash Resistant Fuel System final rule on October 3, 1994. This paragraph sets forth a series of detail requirements for fuel tanks intended to ensure that tank leakage or failure is unlikely. These regulatory requirements pertain primarily to proper support of the tank and protection against chafing.
(2) Procedures. For conventional metal tanks, the support devices, commonly called "cradles," should be designed with wide flanges or cap strips at the contact area with the tank to distribute the loads in the tank material. To prevent chafing, install nonmetallic padding, treated to eliminate absorption of fuel between the tank and the support structure. Cork strips sealed with shellac and bonded to the support structure have been found suitable. Fuel cell sealant material should be applied over rivet heads and in corners. Bladder cells must be designed to fit accurately in the cell cavity in order to avoid fluid loads in the bladder itself. The interior of the cavity should be smooth to avoid damage to the bladder cells.
b. § 27.967(Q):
(1) Explanation. This paragraph requires the design to provide ventilation and drainage of spaces adjacent to fuel tanks to avoid accumulation of fuel or fumes to be expected from minor leakage of fuel tanks. This is needed to minimize the possibility of fire or explosion in these spaces. An exception to this requirement is allowed for bladder cells installed in a closed compartment. For this configuration, ventilation may be limited to that provided by compartment drains if the ventilation is adequate to maintain proper pressure relationship between the bladder cell and cell compartment air spaces.
(2) Procedures. With the assumption that fuel tank leakage will occur, require the tank compartments to be provided with drains at any low point. These drains should conduct fuel clear of the rotorcraft and should be three-eighths of an inch or larger in diameter to minimize clogging. As with any drain intended to function in flight, verification that reverse flow will not occur due to pressure differentials at each end of the drain is appropriate. Ventilation for these tanks should involve openings in the compartment such that in-flight slipstream and/or rotor downwash will rapidly and continuously purge the tank compartment of fuel fumes. Openings should not be located so the fumes or fuel can reenter the rotorcraft. For flexible tank liner configurations (bladder cells), no specific ventilation is required if the cell is located in a compartment which is closed, except for drain holes. Note that a cell leak may be expected to produce fumes in the compartment airspace which are flammable; thus, items installed in bladder tank cavities shall not create a hazard during either normal or malfunction conditions. The vent system for the interior of the cell must be adequate to ensure that the bladder cell interior pressure is always positive or at least neutral with respect to any other airspace in the cell compartment to prevent collapse of the bladder cell. Drainage of the cell compartment should meet the criteria discussed above.
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