606 Appendix H capable of operating to the next subsequent depot interval to avoid immediate teardown when damage is detected and determined to be within acceptable limits. REQUIREMENT LESSONS LEARNED (A.3.3.2.2) Blades insufficiently designed for FOD considerations have caused In-Flight Shut Downs (IFSDs) and costly damage to the engine. FOD tolerance can be enhanced by use of more damage tolerant materials and a redistribution of the blade mass. Foreign object damage costs the Navy over $15M each fiscal year based on a con- servative estimate in one source reference. Common culprits are “housekeeping” items, such as nuts, bolts, safety wire, screwdrivers, etc. This paragraph is aimed at this type of FOD problem (environmental factors such as ice, sand, water, and birds are covered in other specification paragraphs). A.4.3.2.2 Foreign object damage (FOD). The requirements of 3.3.2.2 shall be verified by: VERIFICATION RATIONALE (A.4.3.2.2) Analysis, demonstration, and test are required to ensure that the fan and compressor airfoils can meet the operational requirement of 3.3.2.2 and to establish accept or reject criteria for damage that is detected during flight line inspections. VERIFICATION GUIDANCE (A.4.3.2.2) Past verification methods have included analysis, demonstration, and test. The following verification procedure may be used for guidance. “Simulated foreign object damage shall be applied to the (a) critical stage blades at one or more sections of the (b) of the airfoil. The damage applied shall produce at least the stress concentration factor K t of 3.3.2.2. Following the foreign object damage application, the damaged blades shall be tested to the life required in 3.4.1.1. At the completion of the test, there shall be no evidence of blade failure or flaw sizes beyond values allowed by the in-service inspection flaw size of 3.4.1.7.3 as the result of the foreign object damage. Sufficient instrumentation for monitoring the structure of the engine shall be included in the test engine.” Appendix H 607 The following should be used to tailor the preceding guidance paragraph: (a) The first three stages of the compression system are usually considered the critical stages for FOD damage. (b) The leading edge, or at stage(s) and blade location(s) where the highest stresses of 3.3.2.2 occur. Since notched blades are an HCF concern, they should be tested for the HCF life requirements per 4.4.1.5.2. Background: The locations selected for simulated FOD should be those most sensitive to FOD. Typ- ically, the critical location is where a combination of steady state and vibratory stresses combine, and results in the lowest fatigue life. The addition of a stress riser, resulting from FOD, at this critical location further degrades the life of the airfoil. The engine test cycle should be in accordance with the endurance test. A successful engine test will demonstrate that the engine design is robust enough to safely meet the test requirements. It is recommended that LCF and residual life analysis of 4.4.1.5.2 and 4.4.1.7.1, respectively, be reviewed to ensure proper blade and stage locations are used in the test. VERIFICATION LESSONS LEARNED (A.4.3.2.2) Past engine programs have shown that FOD ingestion and subsequent engine repair are a problem aboard aircraft carriers, because it requires engine module rework or extra maintenance to blend out damage. A.3.3.2.3 Ice ingestion. The engine shall operate and perform in accordance with Table IX, during and after ingestion of hailstones and sheet ice at the takeoff, cruise, and descent aircraft speeds. The engine shall not be damaged beyond field repair capability after ingesting the hailstones and ice. REQUIREMENT RATIONALE (A.3.3.2.3) Sufficient structural capability is needed to tolerate ingestion of environmentally generated ice (hailstones), ice shed from the air vehicle, and engine inlet generated ice. Particles or chunks of ice can be dislodged or break off of inlet duct components such as cowl lips, boundary layer bleed wedges, inlet accessory covers, and variable inlet spikes, and cause compressor damage. 608 Appendix H REQUIREMENT GUIDANCE (A.3.3.2.3) The following should be used to tailor table IX: For inlet capture area of 0065 m 2 100 in 2 the engine should be capable of ingesting one 25 mm (1.0 in.) diameter hailstone. For each additional 0065m 2 100 in 2 increase of the initial capture area 0065m 2 100 in 2 , supplement the first hailstone with one 25 mm (1.0 in.) and one 50 mm (2.0 in.) diameter hailstone. The engine should be capable of ingesting pieces of ice of various sizes and shapes including spears, slabs and sheets. Past engine programs have used 5–8 and 7–9 pieces with one piece weighing at least 0.34 kg (0.75 pounds). For turbofan engines the amount of ice has ranged from 2.3 to 4.1 kg (5 to 9 pounds) of total ingested weight. Within 5 seconds after an ice ingestion event, the engine thrust or power should be at least 95% of the thrust or power immediately prior to the event. The hailstones and sheet ice should be between 080 g/cm 3 50 lb/ft 3 and 090 g/cm 3 56lb/ft 3 specific gravity. Hailstones should be ingested at typical takeoff, cruise, and descent conditions. Sheet ice should be ingested at typical takeoff and cruise condition. Background: All-weather aircraft engines should be designed to withstand potential hail conditions. An engine should also have some capability to ingest, without major damage, chunks or sheets of ice which may dislodge or break off of inlet duct components, such as cowl lips, inlet ramps or doors, inlet accessory covers, and air vehicle surfaces. REQUIREMENT LESSONS LEARNED (A.3.3.2.3) Past engine programs have used these ice ingestion requirements: Airflow kg/s (pps) No. of Pieces Weight Kg (lb.) Velocity km/hr (ft/sec) Miscellaneous Information 112 (246) 7–9 2.3 (5) 54 (50) no piece weighing greater than 0.34 kg (0.75) lb. 118 (260) 7–9 2.3 (5) 54 (50) no piece weighing greater than 0.34 kg (0.75 lb.) 161 (355) 5–8 4.1 (9) not specified no piece less than 0.4 kg (1 lb.), one 0.9 kg (2 lb.) spear 114 cm (45 in.) long Engines that have inlet guide vanes and lack adequate anti-icing or deicing provisions have had operational restrictions and increased maintenance workloads when exposed to Appendix H 609 icing conditions. The problems were eliminated by providing adequate anti-icing systems that reduced inlet case ice accumulation and resultant fan blade damage. Whereas the damage from inlet case ice shedding is a maintenance problem (blending and replacement of fan blades), ice shed from the air vehicle or environmentally generated ice can be a safety of flight concern if sufficient structural capability in the engine design is not provided. Numerous incidents of axial-flow compressor damage caused by ice ingestion are on record. Some of these resulted in complete engine failure and disintegration of the engine. The Engine L, on the Navy’s Aircraft B, passed its Full Scale Development (FSD) icing condition requirement by similarity. However, during fleet operations, engines experienced damage from ice ingestion events. The damage occurred when ice, formed on the inlet lip and duct walls during flight, was dislodged and ingested during descent in warm weather and during hard carrier landings. Naval Research Laboratory Report 9025, dated 30 December 1986, states that about 525 mishaps related to ice ingestion have occurred on U.S. Navy aircraft from 1964 through 1984. It mentions that the Engine J (Aircraft I) often sustained compressor blade damage that was substantial or required an overhaul. Engine J had not been tested for ice ingestion during qualification tests so the effects of ice ingestion were unknown when the engine entered service. It became apparent in field deployment that ice ingestion was a concern and this may have been determined earlier if an ice ingestion test had been conducted during development of the engine. A.4.3.2.3 Ice ingestion. The requirements of 3.3.2.3 shall be verified by: VERIFICATION RATIONALE (A.4.3.2.3) Engine ice ingestion verification is needed to ensure satisfactory engine performance is maintained in icing conditions throughout the aircraft flight envelope. VERIFICATION GUIDANCE (A.4.3.2.3) Past verification methods have included analysis, demonstration, and test. The test procedure should require the engine to run for at least 5 minutes following ice ingestion, before it is shut down for inspection. During the ingestion test, high speed photographic coverage should be taken. Sufficient instrumentation for monitoring the structure of the engine should be included in the test engine. 610 Appendix H The procedures to be used for introduction of the hailstones and sheet ice at the engine inlet and the engine power settings and speed at which the hailstones and sheet ice are to be ingested should be specified in the pretest data. The temperature of the ice should be between −22 and 0 C (28−32 F). The test procedure for sheet ice ingestion should require the most severe ice velocities representing ice shedding off the aircraft inlet lip. VERIFICATION LESSONS LEARNED (A.4.3.2.3) After undergoing icing tests behind a tanker, Aircraft I with Engine J sustained severe damage from ice when the aircraft descended to a warmer altitude and inlet ice was ingested by the engine. For turboshaft engines the most severe ice ingestion velocities may occur at low velocities reflecting the ingestion of ice shed from the inlet. A.3.3.2.4 Sand and dust ingestion. The engine shall meet all requirements of the specification during and after the sand and dust ingestion event specified herein. The engine shall ingest air-containing sand and dust particles in a concentration of (a) mg sand/m 3 . The engine shall ingest the specified course and fine contaminant distribution for (b) and (c) hours, respectively. The engine shall operate at intermediate thrust for TJ/TFs or maximum continuous power for TP/TSs with the specified concentration of sand and dust particles, with no greater than (d) percent loss in thrust or power, and (e) percent gain in specific fuel consumption. Helicopter engines shall ingest the 0–80 micron (0–315×10 −3 in) sand and dust of 4.11.2.1.3 in a concentration of 53 mg/m 3 (33×10 −6 lb/ft 3 ) air for 54 hours and inspection shall reveal no impending failure. REQUIREMENT RATIONALE (A.3.3.2.4) The operation of aircraft in sand and dust environments can result in serious erosion damage to engine parts. Sand and dust particles are highly abrasive and tend to erode the thin metal tips and trailing edges of gas turbine engine compressor blades and vanes. Sand and dust ingestion is also needed to determine the effect on surface coatings. Engine power loss, surge margin loss, and specific fuel consumption gain may adversely affect system safety. Appendix H 611 REQUIREMENT GUIDANCE (A.3.3.2.4) Table XXXV should be used to tailor the specification paragraph. Background: This requirement is based upon a severe but realistic potential service ground environment. It is recognized, however, that the time the engine is subjected to the concentration of sand and dust particles is quite dependent upon the particular engine and air vehicle contribution. The coarse and fine sand and dust particle size is representative of field deployment where pilots will train in the United States desert areas and be sent to the Middle East desert as in Operation Desert Shield and Desert Storm. Large particles are more likely to cause erosion on airfoils (blades and stators) while small particles are more likely to block cooling holes in the turbine and cause corrosion. The SiO 2 in both particle size distributions are likely to melt in the combustor discharge gases and be deposited on first stage turbine nozzle vanes. A verification should be made to determine the effect of different sand compositions on the combustor and HPT at elevated temperatures. Notes on sand and dust concentration in guidance: The thrust or power loss above should be verified at constant turbine temperature for all engine classes. If verifying at constant commanded power setting, the engine thrust or power loss will be less. This is because engines commanding constant fan speed tend to increase thrust as components deteriorate rather than decrease. SFC loss should be verified at constant thrust or power output. The ratio of thrust or power loss to SFC loss varies with engine cycle. The ratio of shaft power loss to SFC loss is as high as 3:1 for typical helicopter engines. The relative losses can be verified by calculating performance with changes in compressor efficiency. The Air Force ATF sand and dust requirement is the MIL-E-87231 and MIL-E- 5007D and MIL-E-8593A sand and dust contaminant requirement at a “53 mg/m 3 (33× 10 −6 lb/ft 3 ) of air concentration” with a 2 hour test allowing a 10% loss of thrust and a 10% increase in SFC. Recent Army helicopter engine specifications since 1971 have all used “C-Spec” (MIL-E-5007C) sand with a concentration of 53 mg/m 3 33×10 −6 lb/ft 3 . The Engine D specification added an additional 54 hour sand test with 0–80 micron (0–312×10 −3 in) AC Fine air cleaner test dust. The Engine H specification uses a 0–200 micron (0–787×10 −3 in) sand (AZ Road Dust Coarse) at a “53 mg/m 3 33×10 −6 lb/ft 3 of air” concentration with a 20 hour test allowing a 10% loss of power and a 10% increase in SFC. The deterioration factors for power and SFC loss may be approximately the same for higher pressure ratio engines. Helicopter engines, with lower pressure ratios, experience power loss more rapidly than SFC loss (at constant T41). The factors may be more equal for higher pressure ratio engines. 612 Appendix H Coarse sand produces blade erosion and a near term loss of compressor efficiency. Fine sand causes plugging of cooling holes, deposits, and long-term engine distress. Fine sand tests should be conducted to determine the engine’s vulnerability to airhole plugging and nature of deposits. Deposits formed also depend upon the composition of the sand. Most of the arid regions of the world correspond to “C-Spec” sand in size except regions in the Middle East. Saudi sand, in general, is close to “C-Spec” in size, very sharp particle shape, with high alumina content. When mixed with water (i.e., waterwash), Saudi sand may form cement, which when heated may form porcelain. Saudi fine sand has been encountered at as high as 4572 m (15,000 feet). Very fine sand is found in parts of Israel and some other nearby regions. The performance loss due to coarse sand ingestion depends upon the total quantity of sand ingested, particle size, and engine power setting. Total quantity of sand depends upon engine airflow rate, sand concentration in the air, and duration of the test (53mg/m 3 33× 10 −6 lb/ft 3 for 50 hour may be equivalent to 530mg/m 3 33 ×10 −5 lb/ft 3 for 5 hour). Particle size ingested varies rapidly with hover height, for clouds created by rotor down- wash. Particle size is largest near the surface and decreases rapidly with height. Power setting affects the effectiveness of the IPS. The IPS is most effective at high power settings where high airflows result in high velocities. Filtration is needed to be effective at all power settings or with the finer particles. A helicopter on the ground at low engine power settings could conceivably experience more sand damage than when hovering a few feet above the ground at high power. A comprehensive sand test should, therefore, consider sand concentration level, composition, particle size in the bed, particle size in the cloud, hover height, time duration, and power setting. The present sand test (con- stant concentration, time, and power setting regime) forms a good benchmark for engine vulnerability to sand erosion, but may not be indicative of actual field experience. USAF bases and commercial airports are permitted to sand icy taxiways and runways with sand particles of up to 0.475 cm 0187 . If prolonged operation in these environ- ments is anticipated, the FOD potentials of this size sand need to be considered. AFI 32-1045 and FAA Circular 150/5200-30A should be consulted for specific information on allowable sand dimensions and quantities. While this requirement paragraph is oriented around thrust or power and SFC perfor- mance, it is stated that the engine has to meet all requirements of the specification. This includes life and performance effectiveness of special technology materials and features on the engine inlet or front face which may be vulnerable to sand erosion and FOD. REQUIREMENT LESSONS LEARNED (A.3.3.2.4) The operation of aircraft in sand and dust environments has resulted in serious erosion damage to engine parts. Sand and dust particles are highly abrasive and tend to erode the thin metal tips and trailing edges of gas turbine engine compressor blades and vanes. Appendix H 613 Helicopter operations are most significantly impacted by the sand and dust problem, although fan blade leading edge damage has been experienced on Engine L in Aircraft B as a result of operation “near sandy, dry desert-like areas such as NAS Miramar.” One source pointed out that the “accelerated replacement of erosion damaged helicopter turbines in Southeast Asia was estimated to cost the U.S. Government about $150 million a year” (during the Vietnam war period). The finer particles of sand and dust have been known to clog turbine cooling passages and cause engine failures. Sand and dust has been encountered several hundred miles out at sea as well as over land masses. Sand has been determined to be more detrimental during foreign operation due to the differences found in the sand composition. A lower melting point has been observed in foreign sand test samples which allows the sand to melt and form a glaze coating on the turbine airfoils. Spalling of the coating and base metal then occurs. Due to the unexpected turbine blade failure and short life encountered in engines with gas temperatures in excess of 1093 C (2000 F), special consideration in blade design should be taken into account. The soil analysis of Saudi Arabia and other middle eastern countries reveals that very high amounts of sulfur, calcium, and magnesium exist. A study has shown that an HPT blade failure occurred because of excessive sulphidation. At high temperatures, a flux or coating of calcium sulfate and magnesium sulfate attaches to the blade, attacks the base metal, and corrodes it, leading to failure. USAF Aircraft A bases have raised questions about operating on taxiways and runways with sand particle sizes specified in FAA and Air Force Instructions. Subsequent analysis revealed that Aircraft A engines are vulnerable to stage 1 fan FOD when operating on sanded areas. Sand and dust ingestion and subsequent performance is a primary cause of engine removal on helicopters. References: Particle distribution: “Development of the Lycoming Inertial Particle Separator”, H.D. Conners, presented at the Gas Turbine Conference & Products Show, Washington DC, 17–21 March 1968. Particle densities: ASME 70-GT-96, J.C. Arribat, ASME Gas Turbine Conference, Brussels, Belgium, 24–28 May 1970. Particle distribution: SEA AIR 947, Issue 1, 2/71. Blade damage including blade and vane erosion, secondary airflow deposits resulting in power reduction and stall margin loss. ASME 68-GT-37, G.C. Rapp and S.H. Rosenthal, presented at Gas Turbine Conference & Products Show, Washington DC, 17–21 March 1968. Particle distribution: “Evaluation of the Dust Cloud Generated by Helicopter Blade Downwash,” Sheridan Rogers, MSA Research, Proceedings of the 7th Annual National Conference on Environmental Effects on Aircraft and Propulsion Systems, 25–27 September 67. Single helicopter dust concentration reached 16.2 mg/cu ft [572mg/m 3 614 Appendix H 357 ×10 −6 lb/ft 3 ] takeoff and approach reach 40 mg/cu ft 1413 mg/m 3 882 × 10 −6 lb/ft 3 . Two helicopter levels of 64mg/ft 3 22601mg/m 3 1411 ×10 −6 lb/ft 3 were reached. Particle distribution: Kaman Report No. R-169, “Amount of Dust Recirculated by a Hovering Helicopter,” dated 26 December 1969. Densities ranged from 0.29 gm/cu ft to 26 gm/cu ft. A.4.3.2.4 Sand and dust ingestion. The requirements of 3.3.2.4 shall be verified by: VERIFICATION RATIONALE (A.4.3.2.4) The effect of sand ingestion on the performance, bleed air quality, and internal air cooling system must be verified. VERIFICATION GUIDANCE (A.4.3.2.4) Past verification methods have included analysis, inspection, and test. ISO/FDIS 12103-1 can be used for guidance on Arizona test dust. The contractor should propose a sand ingestion test to verify the operation and perfor- mance of the engine at worst case power levels derived from the mission profile. This test could be combined with the AMT. If combined, the sand test should occur after post-AMT recalibration and endurance test completion, due to performance degradations of the engine. The following procedure may be used as a verification method. For fixed wing and VSTOL engines: “During the engine test, the coarse sand and dust shall be ingested first, with the fine particle sand and dust ingested afterward. An engine disassembly and inspection shall be conducted between the coarse and fine sand tests as specified by the Using Service. The engine shall be tested at Intermediate thrust, with sand and dust ingested at the concentration levels and for the length of time specified in 3.3.2.4. During each hour of operation, at least one deceleration to idle and acceleration to maximum augmentation shall be made, with power lever movements of 0.5 seconds or less. If an anti-icing system is provided, ten periods of one minute operation of the anti-icing system shall be performed during the first test hour. During the entire test, maximum customer bleed air shall be extracted from the engine. The customer bleed air shall be continually filtered and the total deposits measured and recorded. Following the post-test performance check, Appendix H 615 the engine shall be disassembled to determine the extent of erosion, and the degree to which the contaminant may have entered critical areas in the engine. The test will be considered satisfactorily completed when the criteria of 3.3.2.4 have been met and the teardown inspection reveals no failure or evidence of impending failure.” For helicopter engines: “The engine shall be tested with a 0–80 micron 0-315×10 −3 in sand and dust ingested at the concentration levels specified in 3.3.2.4. The engine shall be tested for 9 hours at maximum, 27 hours at intermediate, and 18 hours at maximum continuous, for a total of 54 hours with not less than 27 starts. The test cycle shall be 10 minutes at maximum, 20 minutes at maximum continuous and 30 minutes at intermediate. All ratings shall be initially set to measured gas temperature associated with rated rotor inlet temperature. After test initiation, the engine shall be run at a constant gas generator speed unless the measured gas temperature associated with the T4.1 deterioration limit is reached in which case the measured gas temperature will be held constant. This test shall be terminated and the engine shall be removed and disassembled if power deterioration based on measured gas temperature is excessive or if engine failure or impending failure is evident. The engine shall be tested with the IPS if it is an integral part of the engine design. During each hour of operation, at least two deceleration to idle and acceleration to maximum continuous speed shall be made, with power lever movements of 0.5 seconds or less. A calibration test shall be performed at the beginning, at 25 hours and end of the test run. The engine shall be inspected at 25 hours by borescope and other visual techniques that do not require disassembly of the engine. Upon completion of the test run the engine shall be disassembled in such a way that the contamination displacement is minimized. The engine shall be disassembled to determine the extent of sand erosion, and the degree to which sand may have entered critical areas in the engine. Each major rotating and stationary component subject to the effects of sand ingestion shall be weighed at engine build, disassembly and after cleaning. The test will be considered satisfactorily completed when 54 hours of testing have been completed and the teardown inspection reveals no failure or evidence of impending failure.” “The engine shall be tested with the coarse sand at maximum continuous rated measured temperature with sand and dust ingested at the concentration levels and the length of time specified in 3.3.2.4. The engine shall be tested with the IPS if it is an integral part of the engine design. During each hour of operation, at least one deceleration to idle and acceleration to maximum continuous rated measured temperature shall be made, with power lever movements of 0.5 seconds or less. If an anti-icing system is provided, ten periods of one minute operation of the anti-icing system shall be performed during the first hour of each five hour cycle. The engine shall be shut down and cooled at least 12 hours following each five hours of sand ingestion. During the entire test, maximum customer bleed air shall be extracted from the engine. The customer bleed air shall be continually filtered, and the total deposits measured and recorded. If an engine internal washing system . coating of calcium sulfate and magnesium sulfate attaches to the blade, attacks the base metal, and corrodes it, leading to failure. USAF Aircraft A bases have raised questions about operating on taxiways. engine vulnerability to sand erosion, but may not be indicative of actual field experience. USAF bases and commercial airports are permitted to sand icy taxiways and runways with sand particles of up. LESSONS LEARNED (A. 3.3.2.4) The operation of aircraft in sand and dust environments has resulted in serious erosion damage to engine parts. Sand and dust particles are highly abrasive and tend