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Manual TBD.23. The PT.1277 is a member of the Proteus family of engines of industrial and marine engines

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Manual TBD.19. The pack incorporates the Bristol Siddeley PT.1277 gas turbine. This is basically the well-tried Proteus engine, which powers the Britannia aircraft, and the Royal Navy's Brave Class fast patrol boats. The industrial engine, used for regular peak-lopping, is designed for an overhaul life of 10 years, and where long periods at continuous running are required, for example, as the sole source of supply in remote areas, the set can be provided at a reduced output rating

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Includes attached article from 'Flight': Bristol Proteus, A High-compression Turbopulp for Wing Mounting (18 August 1949)

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Includes card 'The Marine Proteus, A Gas Turbine Marine Engine'

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In German

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In Spanish

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Olympus aero engine. Bristol Aero-engines Limited. A subsidiary of the Bristol Aeroplane Company Limited. Section 1: Engines T.2358 issue 8; Section 2: Engines T2358 issue 10

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Olympus aero engine. TC.219. The Bristol Olympus B.O1.14R engine is a military version of the Olympus 551 engine, giving, under sea-level static conditions, 13,250 lbs of dry thrust or 23,050 lbs with reheat on. if required for take-off the dry thrust can be increaed to 15,450 lbs. The Bristol Olympus 551 is an engine which is subject of a licence agreement between this company and the Curtis-Wright Corporation, Woodridge, New Jersey - the engine being marketed in the USA as the Curtis-Wright Zephyr. The civil form from the engine has been offered to the world's airlines as an engine for medium range subsonic jet transports. The performance curves (as listed in Appendix A) 10 sheets and curve B.124012A. Appendix B (T.2280.A and curve B.68015.B)

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Olympus aero engine. APF.48. This performance folder is intended to be used in conjunction with the latest issue of MoS Technical Certificate BLT/5 and subsequent amendments, which should be consulted for operating limits on LP compressor rpm, HP compressor rpm, and jet pipe temperature. The accompaning curves are based on the measured flight performance of development engines; Appendix A details the correction to be applied to bring the folder into line with production engines. Appendix B explains the method of using the non-dimensional curves. In using the folder 100% LP rpm should be taken as 6,500 rpm

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Olympus aero engine. Restricted. TS.1324. Bristol Aero-Engines Limited. The Bristol Olympus Mk.104 is a two-spool turbojet with a maximum rating of 13,000 lb thrust. It has a high thrust/weight ratio and a low specific fuel consumption

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Olympus aero engine. PS.70/2. The Olympus BS.O1 22R is derived directly from the Olympus BS.10.15 by the addition of a zero-stage to the LP compressor. The engine is equipped with a fully variable reheat system which is being developed in conjunction with the Solar Aircraft Company of San Diego, USA. The performance in this brochure has been estimated by methods agreed with the National Gas Turbine Establishment and the Ministry of Aviation, and all data quoted are for an average engine

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Olympus aero engine. TRH.38. The BS Olympus Mk.104 engine is a two-spool turbojet with a maximum rating of 13,000 lb thrust. It has a high thrust/weight ratio and a low specific fuel consumption. Many thousands of hours of flying under service and development conditions have proved the soundness of the two-spool compressor principle which provides the engine with exceptional flexibility and ease of control in spite of the high pressure ratio

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Olympus aero engine. AD(C)4. The Olympus, the first twin-spool engine to be designed, ran for the first time in 1950. It has an excellent specific fuel consumption, good handling characteristics, and ease of accessibility of major components. The original engine, now the Olympus 100 series,was developed for the Avro Vulcan bomber Mk.1 aircraft; later derivations (200/300 series) now power the Vulcan B Mk.2 and further members of the family, 320 and 593 are being developed for more advanced supersonic aircraft. Some versions (200 series) also have industrial and marine application. Production of the Mk.101 started in 1954, followed by more powerful Mk.102 and Mk.104 engines. Production ceased in 1959

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Olympus aero engine. B Olympus 3-1. Bristol Aero-engines Limited. A subsidiary of the Bristol Aeroplane Company, Filton Bristol. After an official type test in November 1955, the Olympus 102 was approved for service at 12,000 lbs thrust without reheat; security restrictions do not permit the release of later information. In 1953 an earlier, less powerful version, installed in an English Electric Canberra MK.2, had raised the world altitude record to 63,668 ft in practical proof of the claim that the Olympus produced more power at higher altitude than any other engine. On August 29 1955 the same aircraft, with later type Olympus engines, improved its on record to 65,890 ft. In spite of this remarkable performance the Olympus is also one of the most economical turbojet in the world

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Olympus aero engine. Bristol Aero-Engines Ltd; Technical Publications Department, Filton, Bristol

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Ref: 434c.1/4. 10-page hand-written report (in pencil) of blade calculations by KER. Tabulated summary of blade stresses for roots of varying pitches

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Ref: 4343/5. Hand-written (in pencil) report and calculations by ENC. These plates should be as designed that when rotating, they exert a net axial thrust which balances the gas loading on the rotor blades. This axial thrust is brought about by the provision of a suitable cone angle and rim on one of the plates. In this case the gas loading falls back on to the front plate and so this plate must be designed to balance the load, the [ ] of which is 5500 lbs under ICHN conditions. Since the clamping load is solely dependent on speed it is evident that this load can only be correct for one set of atmospheric conditions. Under any other set the gas load will either be greater or less than the clamping load and this out-of -balance load will tend to deflect the plates. The plates must therefore have sufficient stiffness to ensure that this out-of-balance load does not produce a large deflection with corresponding shift of blades. Additional effects to be considered are gyroscopic pression and possible pressure drop acoss the plates. Graph 1: Rotational stresses in clamping plate (2nd design). Graph 2: Radial displacement of plate and labyrinth ring (2nd design)

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Ref: 434C.1/4. Hand-written (in pencil) by KER. Shows the root fixings in addition to the binding stress in the lobes. Page 2: tabulation of blade root stresses

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Ref: 4343/5. 16 page hand-written report (in pencil) with drawings of components and a slope curve for disc subjected to a rim load of 100lbs. by KER

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Ref: 434C.1/4. Hand-written (in pencil) 5-page report with diagrams of components and calculations by KER

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Ref: 4343/3 from Stress Office. Design memorandum GDM/BE/10/6. The rotor disc profiles should be in accordance with a previous recommandation (see GDM/BE/10/1 15 september 1948)

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Different content to PA1716/5/11/2/6 although references are the same. Stress office Ref: 434c.1/3. 434SR Design memorandum GDM/BE/10/6. The ending clamping plates conform with our earlier suggestions. 1 Drawing of clamping plate front Mat=S.II drawing of clamping plate front. Mat=S.II drawing of clamping plate rear

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Ref: 4343/3 Hand-written (in pencil) 10-page report with drawings of components and stress calculations by RD Fuller

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Ref: 434C.1/3. 2-page report

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434c.1/3. 10-page report of calculations by KER

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Ref:434c.1/3. Design memorandum GDM/BE.10/42

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Ref: 4343/4. design memo GDM/BE.10/30. Steel rotor blades for hollow steel blades. KWF/EN

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Ref: 434c1/4. Steel rotor blades for hollow steel blades. Drawings OLS.245 & OLS.246 KWF/EN

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Ref: 434c.1/3. Compressor driving shafts B.25541 & B.45334. Calculations 22 January 1951. Memo 26 February 1951

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Ref: 434c.1/4. Root fixing with hollow steel blades. sch/EMV/E to M/s 1123 and OLS.245 WM Owen. 7-page report, calculations and drawings of precision cast root to M/s1123

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Ref: 434c.1/3. 3-page report and calculations by KER

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Ref: 434c.1/3. report and drawing B.50214 attached

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Ref: 434c.1/4. 7-page (in pencil) report with calculations

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Ref: 434c.1/3. Calculations attached

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Ref:434c.1/3. Report with tabulated type of bolt and tightening extension details by KR Chamberlain

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Ref: 434c.1/3. Tightening extensions for new bolts of S.106 material compound with the existing bolts of Bace 140 material. Scheme drawing used OLS.313. Previous calculation 4343/3 10 December 1953; 4342SR/10. Page 6 graph: Tightening extension - .001 ins

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One page tabulated performance of Olympus Mk.104 rotor at sea level; M 0.84, 240Kts, 350Kts; Max rating etc.

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Three pages of calculation. Final page: Alternative breakdown of a low level sortie at ISA

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Ref: 4343/6. We will consider now the few outstanding points which have not been covered in previous calculations: 1 Corridor rim bearing length between blade platform and spacing ring, 2 Corridor binding between spacing rings. 3 Corridor shear in main driving shaft.KER

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Ref: 4343/5. Design memo GDM/BE.10/6. Alternative designs for low pressure compressor clamping plates. 3-page report with drawings Mat.SII, and Sixth stage disc replacing existing Dural disc and rear clamping plate

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Ref: 4343/6.TC Greaves. An appropriate investigation into the suitability of a modified spacing ring locating radically on the disc and axially on the back platforms

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Ref: 4343SR/17. RE Jenkins. Schemes or drawings used: OLS.412.37-page report with calculations and graphs: 37 Olympus Mk.2 LP compressor 1st & 2nd stage discs; 38Mk.2 LP compressor rotor 1st & 2nd stage discs; 39 Mk.2 LP compressor rotor 3rd & 4th stage discs. Ref: 4341c.1SR/6. Last page: Olympus Mk.1 spacing rings. Proposal to slot flange to receive firm radial location

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Ref: 4343.SR/8. KB Chamberlain. report and graph. Graph GR.2755: Stiffness test LP compressor rotor shaft

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Ref: 4343/10.Olympus Mk.3 LP compressor rotor shaft and rotor centre. Drawing used 035 116 and 144. Page 1,2: Calculations. Pages 3-5: Loading tabulations. Report by RE Jenkind

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Ref: 4343/5. Olympus Mk.1/2c LP compressor rotor assembly. Pages 1-8: Clamping load due to front flange. Pages 9-11: Clamping load due to rear flange. Pages 12-17: Deflection of rotor assembly. Page 18: Bearing pressure on spacing rings. Report by JA Taylor

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Ref: 4343SR/5. Olympus Mk.1/2c (101)ompressor (LP) front rotor flange. Drawing OLS 1317. This scheme incorporates the change to MS 1503. Pages 1-5: Calculations of clamping force. Pages 6-9:Stiffness of flange. Page 10: Stress distribution when deflected. Page 11: Overall rotor deflection. NB: Refer to MS 1503. Report byRE Jenkins

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Ref: 434C1/12. Olypmus Mk.1/2c LP compressor rotor. Data used TED 38 (6894/3). Previous calculations 4353/4 and 4363/11. Report by DJ Fortune

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Ref: 434C1/12. Memo: Experimental Department Defect report 145/54. Defect: Loss of 'nip' between the rotor flange front B.89455 and the last stage blades on engine rebuild. Action taken in shops: On most of the development a concession has been obtained to restore the 'nips' by removing metal up to a maximum of 0.10 from the first stage disc abutment face on the flange

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Ref: 434C1/5. AF Elliot 'Could you please furnish us the following centrifugal stresses based on front and rear rotor speeds of 6150 and 7750 rpm, or as close to these speeds as you have calculated. Memo attached 434c1/5 9 June 1954

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Ref: 434C1/1 and 434C2/1. Regarding the proposal to fit an adjusting washer under the rotor flanges (front) of the HP and LP compressor rotors we find that reducing the hub thickness of these parts by 0.50', required to accommodate this washer, leads to high stresses in the rotor and flanges

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Ref: 434c1/5. One page reply: Disc and rotor flange centrifugal stresses; set out in tabulated format

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Ref: 434c1SR/5. Olympus Mk.104 LP compressor end flange and casing. Scheme/drawing used OLS 3750. Previous calculations 4343SR/5 December 1952. Report by RE Jenkins

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Ref: 434c1/13. Internal memo: One page tabulated report by N Essex

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PFG/RCP/8549 from RC Plumb. I should like to review again the possibility of flying the Olympus 104 in low level role for training, particularly in respect of compressor blade vibration

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Ref: 434c1/13. Scheme/drawing used B.97852 hand-written report with a diagram of torque range and calculations

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Ref: 434c1/13. LP shaft and couplings when used in a low level training role

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Ref: 434c1/13. The pop-up manoeuver associated with the low level role involves several cycles at torque higher than normal TC. A review was therefore made of the stresses in both LP and HP couplings. Tables 1 and 2 give tabulated results of tests

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Ref: 434c1/5. Rear flange with labyrinth passage removed OLS.1021. Bearing pressure against disc face. Graph and tabulated stresses. Olympus 104 de-iced zerostage rotor blade scheme OLS.3750

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Ref: 4343/5. Rear clamping plate. General consideration of rear clamping plates. R Humpridge

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Ref: 434c1SR/9. Disc, bolts and spacing rings; schemes/drawings used 03S. Pages 1-2: Calculations of disc loadings. Pages 3-7: Disc profiles. Pages 8-12: Graphs. Pages 13-15: Possible design of 1st disc in steel. Pages 16-21: Reactive forces of cored discs. Pages 22-25 Main rotor bolts. Pages 26-29: Spacing rings. Page 30: Generator drive coupling

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Ref: 4343c/3. Drawing and calculations of LP compressor front rotor centre

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Ref: 623D/7. Bristol Gas Turbines, Bristol Aeroplane Company. 8 pages of tabulated calculations

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Ref: 623D/7. Bristol Gas Turbines, Bristol Aeroplane Company. 7 pages of calculations

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Ref: 623D/7. Traced 29 October 1951. 1 B.59350 Olympus 1/2 (Bristol cannular conbustion chambers)

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Ref: 623D/7. Bristol Gas Turbines, Bristol Aeroplane Company. 11 pages of calculations with drawing

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Ref: 623D/7.We have today been asked to quote the critical speeds to avoid when operating Olympus and we give below the anticipated critical speeds based on calculations

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Ref: 623D/7. Confirming our conversation during the incremental running of Olympus 502 on the 1st November, it was very noticeable that the metal detector indicated the presence of metal between 7,100 and 7,300 hp rpm in appreciable quantities, but at other speeds it gave no appreciable indication of metal

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Ref: 623D/7. 3 pages of drawings and calculations

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Ref: 623D/7

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Ref: 623D/6. KCF 4 page report and calculations

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Ref: 623D/6. FH. 27 pages of drawings and calculations

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Ref: 623D/6. 9 pages of calculations

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Ref: 623D/6. 9 pages of calculations

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Ref: 623D/6. 11 pages of drawings and calculations

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Ref: 623D/6. 5 pages of calculations and drawings

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Ref: 623D/6. 13 pages of calculations

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Ref: 623D/6. Report No.T.521/53. Issued by Component Development Department Aero Engine Division Bristol Aeroplane Company Limited. In continuity of the tests already conducted on the B.0L.4 HP compressor rotor up to a maximum speed of 8,300 rpm ie 5% overspeed (see CDD Report No.513/53); it was necessary, in order to complete the specified schedule, to carry out further tests on the rotor up to a maximum speed of 8,690 rpm (10% overspeed), and hold the rotor at this speed for a period of five minutes

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Ref: 623D/6. 9 page report and calculations with Stress Graph 4346 SR/8 (GR.3416)

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Ref: 623D/7. 8 pages of calculations with drawing of Olympus Mk.101 Sunstrand constant speed unit

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Ref: 623D/7 A further vibration excitor test has been undertaken with the casing supported in a more realistic fashion and with the intention of discovering where the flexibilty lay

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Ref: 623D/7. Stress Department Report SV.73 dated 12 October 1955 showed the result of a test on 01.606 which suggested that the engine was whirling in the upper speed range

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Rf: 623D/7. Frequency checks have been carried out on the turbine drain valve to OLS.809 Issue 6 fitted to the exhaust annulus ex OL.525

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Ref: 623D/7

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Ref: 623D/7. 10 page report and calculations

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Ref: 623D/7. 6 pages of calculations and graphs

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Ref: 623D/7. 2 pages of calculations with drawing

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Ref: 623D/7. TP.2436

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Ref: 6234 SR/6

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Ref: 623D/6. Prog section ref: 12015/75. 9 pages of calculations and graphs

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Ref 623D/7. Graph

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Ref: 6234SR/7

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Ref: 623D/7. 3 pages of calculations

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Ref: 623D/7. An investigation has been carried out recently into the possible presence and source of dangerous vibrations in the Olympus auxiliary drive system

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Ref: 623D/7.Report 209/56. Results are shown in this report

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Ref: 623D/7. We have now completed a calculation of the torsional critical of the system incorporating modifications detailed in the following schemes OLS.3378 Issue 5; OLS.3582 Issue 1; OLS.3381 Issue 3

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Ref: 623D/7. 18 pages of calculations

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Ref: 623D/6. 12 pages of calculations and graphs

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Ref: 623D/6. 5 pages of calculations including graph

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Ref: 623D/6

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Performance and Power Plant Engineering Department: Graph TED23456 Olympus 104 LP compressor running lines

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Performance and Power Plant Engineering Department: Graph TED22738. Titanium O and 1st stage LP rotor blades. Avro HPEGV setting

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Issued by: Performance and Power Plant Engineering Department: Graph TED23417 Engine 104003 Build 11. Steel O stage LP rotors - 20deg LP EGV setting Engine 104008 Build 5. Aluminium O stage LP rotors - 5deg LP EGV setting Engine 104002 Build 12. Titanium O stage rotors - 20deg LP EGV setting

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Issued by Performance & Aerodynamics Department: Graph TED9158

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Issued by Performance & Aerodynamics Department: Graph TED9157. Data from curves 100% AF Factor =3.00ft/sec.

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Ref: CE/EW-D/HCH?9790. From Service Development - Turbine Engines and Power Plant Department. Includes related graph TT1736/1. 518/23

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Ref: CE/EW-D/GFH/9852. From Olympus ( Production Engines Project Engineer). The object of the meeting was to review the action being taken as a result of the failure of the 1st stage LP compressor blades of the 3 lobed root type in Engine 101045 installed in Vulcan XA898 in No.2 position, and the cracking of the 2 lobed root 1st stage blades (R-R.58) in flight engines at hours varying between 150 and 200. Examination of the flight programme of Vulcan XA.898 operating in the RAF showed that there had been a considerable amount of low flying whereas strain gauging in the air on the 2 lobed root had in the main been done at altitudes above 20,000 ft. As a result of this Vulcan XA.891 flew strain guage tests with Engine No.10117 (installed in in No.2 position) fitted with 3 lobed root blades at altitudes up to 10,000 ft at varying aircraft speeds, and throughout the speed range of the engine. Includes 3 related graphs

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Ref: CE/EW-D/GFH/10589. From Olympus (Production Engines - Project Engineer). The object of the meeting was to review further work being done on strain gauging, aerodynamic tests and mechanical tests to improve the overhaul life of the 1st stage LP compressor blades in the Olympus 101 engines, and to establish the flight life of 102 and 104 engines relative to zero and 1st stage LP compressor blades

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Blade Engineering Department: Notes on a meeting in Mr Blackwell's office 11 July 1957. 1.0. Strain gauge testing. 2.0 Analysis of strain gauge results. 2.1 Standard engine (Test 1a). 2.2 Venturi engine (Test 2). 2.3 Canberra intake and venturi nozzles (Test 3). 2.4 Test at -20deg EGV setting. 3.0 Discussion of results. 3.1 Effect of intake. 3.2 Effect of nozzle. 3.3 Effect of trimmers. 3.4 Explanation of failure of 519. 4.0 Future testing. 5.0 Flutter margin of Olympus 102 engines

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Project Department: Ref GN2066: From Project Department. Following the 1st stage compressor blade failure of an Olympus 101 in Vulcan XA898 at Waddington, it was realised that little was known about the aerodynamic flow conditions at the Olympus engine face of this aircraft. Since the aerodynamic reason for the blade failure was unknown, a series of tests were carried out by BAE and Avros to obtain more information about conditions prevailing in the intake duct. The test carried out by BAE were on a full scale Olympus 102 engine in the Vulcan wing rig and only simulated ground running. Whereas the Avro tests were on a 1/7.5 scale model of a wing fuelage combination carried out in a number of flight attitudes as well as the ground running case

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Ref: CE/PFG/123. From PF Green to Sir W Alec Corton. The recent failure of the port engine 104006 in the Canberra after 17 hours flying time has been established as originating from fatigue failure of two zero stage LP compressor blades just above the blade root platform, the crack starting from the leading edge. The marking on the blade root serrations on this stage was not abnormal, but the first stage blade root serrations were heavily fretted

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Ref: CE/EW-D/HCH/13350. From Service Department - Turbine Engines and Power Plant. The meeting was held to discuss the cause and effects of the zero stage blades on Olympus 104 engine No.104006 in the port position of the Canberra aircraft whilst on take-off

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Ref: CE/EW-D/GFH/13436: From Olympus (Production Engines) Project Engineer. The development programme agreed to investigate the resonant conditions of the zero and first stage blades in R-R58 and the programme to clear the same blades in steel for the Olympus 104 is a follows: 1 Engine No.101046. 2 Engine No.104003. 3 Engine No.104002. 4 Engine No.102007. 5. Page 2 of this report is missing

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Ref: CE/LBJ/EKA/7214. Mechanical Research department. The test is primarily to establish whether the amplitude of vibration in the 1F and 2F modes of the zero and first stage blades are significantly different from those of the Olympus 104 engines. It is also intended to explore the speed range between 800 rpm and 2,100 rpm to establish whether there is a severe vibration which would explain the 104 Canberra failure

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Ref: CE/LBJ/EKA/7217. Mechanical Research Department. In order to carry out an endurance test on the zero and first stage blades of an Olympus 104 engine which will be comparable to the most serious flight conditions, it is necessary to increase the amplitude of vibration which are normally obtained on the test bed.

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Performance Department No.225. Strain gauging tests carried out on the Olympus 102 & 104 in the Vulcan have shown that high forward speeds increase the amplitudes of vibration of the zero and first stage LP compressor. This has prompted an investigation into the effect of high forward speed on the LP compressor operating lines. A comparison has also been made of the test bed running lines for several standards of Olympus 102 and 04 engines. High forward speed at low altitude move the LP compressor nearer surge when the engine is running at low engine speeds. This only applies to the engine speed range where the engine final nozzle is unchoked on the test bed. The non-dimensional conditions obtained at high forward speed can be simulated to some extent by running the engine on the test bed with larger final nozzles. Pressure levels will, of course, be lower on the test bed

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Ref: CE/LBJ/ACN/7228. Mechanical Research Department. Engine operational history before failure (Ref: CE/AJT/ESG/3217. The tables showing a 'break-down' of the port engine (B.0L.104006) running time to failure were examined again. Full details of the last six performance flights were given by Mr Zanker and are summerised in this report

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Ref CE/LEJ/EKA. Mechanical Research Department. MRD Tech -Memo No.153/58

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Ref: CE/LBJ/ACN/7228. Mechanical Research Department. Technical Memorandum 157/58. The tests are to establish the amplitudes of vibration of the blading obtainable on the test bed for comparison with the flight results. The results will also enable the two blades to be used during the flight tests to be selected

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Ref: CE/LBJ/RNTR/7279. Mechanical Research Department. Engine No.104002 with the R-R58 'O' stage rotor blades will be available on the 12 March

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Ref: CE/LBJ/EKA/7296. Mechanical Research Department. No.2 Shadow

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Ref: CE/LBJ/EKA/7361. Mechanical Research Department

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Ref: CE/LBJ/EKA/7398. Mechanical Research Department. The aim of the test series is to establish that the steel blades in the zero and 1st stages will operate satisfactorily during flight testing in the Canberra. On a previous flight with an R-R58 bladed engine, large amplitudes of vibration in the fundamental flexural mode were observed at 45,000 and 50,000 ft. The test programme will broadly follow that used during the investigational flight into the recent Canberra failure, but will be extended to include as many operational conditions as possible

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Ref: CE/LBJ/EKA/7439. Mechanical Research Department. The aim of the flight test is to determine whether the large amplitudes of vibration, which were found on the zero and 1st stage rotor blades in the Canberra engine at 50,000 ft exist; also an aluminium zero stage and 1st stage when in a Vulcan. The maximum altitude covered in detail by the previous flight tests was 40,000 feet

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Ref: CE/LBJ/EKA/7475. Mechanical Research Department. During the second flight (No.40 & No.41) of the series to the modified programme CE/LBJ/EKA/7470 at 45,000 ft, the engine speed, while the aircraft was accelerating, was not held constant. It is thus proposed to repeat this part of the test during the third flight of the series (flight No.42). This may be achieved by coupling 2 'O' stage and 2 1st stage blades into the recorder. The detail requirements for the third flight as listed in this report

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Ref: CE/LBJ/EKA/7493. Mechanical Research Department

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Ref: CE/LBJ/RNTB/7512. Mechanical Research Department. A meeting was held on Wednesday 25 June to formulate a strain gauge programme and determine the motorised throttle running to be carried out on the steel zero and 1st stage rotor blades

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Ref: CE/LBJ/EKA/7519. Mechanical Research Department

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Mechanical Research Department. Tech- memo 205/58. A meeting was held in the Stress Office Conference Room on 18 July 1958. The meeting was held in the Stress Office on 18 July, in order to find some plausible explanation of why the 104 engine behaves differently in the Canberra and Vulcan. Past strain tests have shown that at altitudes of up to 20,000 ft large ampitudes of vibration are present over the zero stage blades of the Vulcan, which are not present over the same range in the Canberra. Tests on both aluminium and steel blades on the Canberra have shown large amplitudes at high altitudes which are absent on the Vulcan

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Ref: CE/LBJ/EKA/7570. Mechanical Research Department. In order to improve the impact resonance of the LP blading assembly of the Olympus 6 engines, a scheme has been drawn up in which the stator platforms are bolted to the casing. A strain gauge test is to be carried out to compare the amplitudes of vibration of the blades with hose for the previous standard loose dovetail fixing

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Memo: CE/LBJ/EKA/7566. Mechanical Research Department. During development running with the cooled turbine on engine No.101016, an LP compressor failure occurred. This was found to be due to a 1st stage blade failure. Due to the effects of the cooled turbine, the LP running line of the engine was moved much nearer surge than a standard engine. One aspect of the present testing is to reproduce this running line and see if there is significant increase in blade vibration amplitudes on the zero and 1st stage

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Memo: CE/LBJ/EKA/7572. Mechanical Research Department. The basic purpose of the test series is to confirm the dangerous amplitudes of vibration obtained during flight strain gauge testing of Olympus 104004(3) have been reduced to a satisfactory level by use of steel blading in the zero and 1st stages. The programme is then intended to be extended to clear the blading over much of the operational range as practical

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Ref: CE/LBJ/EKA/7548. Mechanical Research Department. The aim of the tests is to examine the characteristics of vibration of Titanium blading. This stage has been chosen since comparative test results are available in both aluminium and steel

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Memo: CE/LBJ/EKA/7623. Mechanical Research Department. The aim of this series of tests is to observe if any dangerous amplitudes are caused by the attitude of the aircraft or by any other aircraft or engine manoeuvres of a special nature which are part of normal service procedure

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Mechanical Research Department. No.2 shadow

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Tech Memo: 225/58 Mechanical Research Department

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Ref: CE/LBJ/EKA/77433. Mechanical Research Department

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Ref: CE/LBJ/EKA/1563. Installation and Flight Development Department

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Ref: STR/J/E/50

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Ref: CE/LBJ/DAB/8054. Mechanical Research Department

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Olympus Jet pipe characteristics. From LE Leavy, Superintendent Wind Tunnels, Aircraft Research Division: I enclose some preliminary data on the swirl which we measured under flight conditions on the 1/.7.5 scale Vulcan model. This information should be treated as preliminary as we have as yet no accurate knowledge of the velocity in the engine duct, consequently the values Vo/VE quoted on the diagrams have a probable accuracy of no more than +10%. 4 graphs attached showing swirl in the engine ducts of the 1/7.5 scale Vulcan

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Jet pipe diameter 22.6'. Exit nozzle diameter 20.5'. Graph TED.2561. (Superceding TED.2122 & 2122A)

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Jet pipe diameter 22.6'. Exit nozzle diameter 20.5'. Graph TED.2697

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 24.5'. Exit nozzle diameter 21'. Graph TED.2756. Test bed obtained from Olympus 510 Build 5. Olympus 511 Build 6. Olympus 511 Build 7

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by JYB. Jet pipe diameter 24.5'. Exit nozzle diameter 21.5'. Graph TED.3943. Superseded by TED.4390. Test points for 1/2B Olympus engine 511 Build 12 shown on curves

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by JYB Jet pipe diameter 24.5'. Exit nozzle diameter 22'. Graph TED.3944. Test points for 1/2B Olympus engine 511 Build 12 shown on curves

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Jet pipe diameter 24.5'. Jet nozzle diameter 21.25'. Graph TED.4142. Note: In the absence of test results. These characteristics assume thrust and discharge coefficients taken from TED.2756 (characteristics with 21' nozzle). A new curve will be issued when test results are available

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Jet pipe diameter 24.5'. Jet nozzle diameter 21.25'. Actual nozzle diameter 21.203. Graph TED.4304. Test results: Olympus 101 No.523 Buiild 4 & 5

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Jet pipe diameter 24.5'. Jet nozzle diameter 21.25'. Based on test results obtained with Olympus 101 Engine No.523/5. Graph TED.4390

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC Jet pipe diameter 24.5'. Jet nozzle diameter 20.5'. Based on test results obtained with Olympus 101 Engine No.523/5. Graph TED.4391

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Jet pipe diameter 24.5'. Jet nozzle diameter 21'. Curve based on test results for Olympus 101 Engine No.523/5. Graph TED.4401

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 24.5'. Jet nozzle diameter 21.203 (Actual)'. Test result Olympus 101 Engine 521 Build 9. A= Test results (21.177' nozzle) Olympus 521/10. TED.4491

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 24.5'. Jet nozzle diameter 21.203 (Actual)'. Test result Olympus 101 Engine 521 Build 9. Also Engine 522 Build 9. A= Test results (21.177' nozzle) Olympus 521/10. TED.4491

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 24.5'. Jet nozzle diameter 21.4' (Actual 21.396'). Test result Olympus 101 Engine 521 Build 10. TED.4992B

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Jet pipe diameter 24.5'. Jet nozzle diameter 21.4' with 1 trimmer. Curve based on test for Olympus 101 Engine 523 Build 7. TED.5485

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Jet pipe diameter 24.5'. Jet nozzle diameter 22.6'(nominal). 522 22.591'(actual); 521 22.562'(actual). Curve based on test results for Olympus 98Z Engine No.522 Build 11 TED.6015 (superceded by TED6945)

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 24.5';Jet nozzle diameter 23' (nominal). 23.060(actual). Curve based on test for Olympus Engine 521 Build 14. Olympus Engine 522 Build 15. TED.6522

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Jet pipe diameter 22.5'; Jet nozzle diameter 22.6'. (Superceding TED.6015). Drawn by DWL

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. 27' diameter Avro pipe; 22.389' diameter inclined flight nozzle. TED.6987. This curve is consistent with TED.6091 Issue 2. Engine No.10132 Build 2 No,2.GP. 26.7555. TED.6987

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. 24.5' diameter Canberra jet pipe; 22.6' nominal diameter nozzle. Gross jet thrust. TED.20049

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. 24.5' diameter Canberra jet pipe (old type RCJ 4 CS 4120); 22.6' nominal diameter nozzle (old type 4 statio tappings). Mass flow. New curves drawn for Nimonic or st.steel. Jet pipe with 8 tappings. TED.20048

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. 24.5' diameter Nimonic or stainless steel jet pipe to latest standard (B.Static tappings) 22.6' nominal diameter nozzle. TED.20790

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. 24.5' diameter Nimonic or stainless steel jet pipe to latest standard (B.Static tappings) 22.6' nominal diameter nozzle. Gross jet thrust. TED.20789

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. 27' diameter Avro jet pipe. 22.4' diameter flight nozzle. TED.8208

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. 27' Avro diameter jet pipe. 22.4' diameter flight nozzle. TED.8207

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 27'; Jet nozzle diameter 21.75'.TED.2895

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 27'; Jet nozzle diameter 21.75'.TED.2895

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 27'; Jet nozzle diameter 21.75'.TED.3445. Calibrated on Engine No.2603 Build 14

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 27'; Jet nozzle diameter 20.75'(nominal) with trimmers.TED.3813

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Jet pipe diameter 27'; Jet nozzle diameter 21.2'(nominal) with trimmers.TED.4654B

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Covering ground run range only. TED4827

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Supercedes TED.4827)

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. Production jet pipe: 21.3'nozzle with one 3/16'th by 10'. TED.5275

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DC. Production jet pipe: Final nozzle diameter 21.3'; Cone angle 24deg (included). TED.5464

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. 22.55' diameter nozzle with trimmers. Note: In the absence of results, this curve has been taken to be the same as that for the Acr pipe with 22.413' diameter flight nozzle plotted on TED.6091. TED.7817

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. 22.55' diameter nozzle with trimmers. TED.7816

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. 22.423' diameter nozzle with two trimmers (7.5' by 1/4'). TED6090

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. TED.6824

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. 27' diameter Avro jet pipe. 22.389' diameter inclined fight nozzle. TED.6091. Issue 2

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. 27' diameter pipe. 22.389' dia nozzle. NB: Both jet pipe and static pressure were measured at front of pipe. TED.7330

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. Note: In the absence of test data these characteristics have been based on the thrust and discharge coefficients used in preparing the 21.2' nozzle curves given on TED.5241. TED.7342

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. NB: This curve is consistent with TED.5241. Includes related graph - 27' dia Avro jet pipe; 22.4' dia (nominal) flight nozzle. TED.7352 (Supercedes TED.7330)

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. TED.7482

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. TED.7483 and 27' dia jet pipe; 22.4' dia nozzle with 24 square' trimmer area. TED.20382

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by DWL. This curve is consistent with TED.8207 & 8. TED.8955

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by FGY. 27'dia Avro jet pipe; 22.4' dia (nominal) flight nozzle with 4 bar trimmers. TED.9100

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by FGY. 27' dia Avro jet pipe; 22.4' dia nozzle with 6 aerodynamic trimmers (each 4'square). Data from Engine No.526 Build 24. TED.9406

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by FGY. 27' dia jet pipe; 22.4' dia nozzle with 6 aerodynamic trimmers (each 4 square'). Data from Engine No.526. Build 24. TED.9407

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. 27' dia Avro jet pipe; 22.4' dia nozzle with 6 square' trimmers aera. TED.9976

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. 27' dia Avro jet pipe; 22.4' dia nozzle with 6 square' trimmers aera. TED.9977

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. 27' dia Avro jet pipe; 22.4' dia nozzle with 6 square' trimmers aera. TED.20381 (Supercedes TED.9406)

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. 27' dia jet pipe; 21.75' dia slave nozzle. TED.5152

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. Based on ground run of port engine No.527 Build 10 on 5 November 1958 with venture intake. 22.5 nom dia nozzle with 30 square' trimmer area. Air mass flow. TED.21638

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Issued by the Performance & Aerodynamic Department. Engine Division Filton Bristol. Drawn by EAW. Based on ground run of starboard engine No.522 Build 31 on 6 November 1958 with venture intake. 22.5 nom dia nozzle with 15 square' trimmer area. Air mass flow. TED.21637

Ref: 434c.1/3. 4-page report and calculations

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Ref: 434c.1/4. Estimation of weight of blade above base (2nd group). 8-page report and calculations. Drawing of 4th stage; Root sections

Ref:434c.1/3. Design memo GDM/BE.10/71. Olympus front section compressor driving shaft B.51025; calculations attached KWF/VAK

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5214SR/17. Drawing of parts tested: B.47197. The starter drive shear neck has been subjected to a static load/deflection test in order to determine the yield and ultimate torques for comparison with for comparison with the design failing torque of 110lb.ft. An applied torque of 108lb.ft caused marked yielding, and the neck shared at 125lb.ft.

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Ref: 434c.SR/3. report and calculations

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5214SR/18 As promised by Mr Stockdon during today's telephone conversation we give the part numbers for the above bearings: MRJ.15 - FB.196110 (inner, bage and rollers). XRJ.11/2 - FB .196111 (cage and rollers). XIJ.11/2 - FBS.225/12/AZB.

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5214SR/18. We are enclosing herewith a copy of our scheme MS.1453 which gives preliminary details of the bearings and their running conditions in the above drive

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5214SR/18. We are enclosing prints illustrating the inner, cage and rollers of our LRJ.17 and the cage and rollers of XLRJ.1.1/2' and we trust the information will enable you to proceed with your design. Includes related drawing

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5214SR/18. Further to your letter of the 1st inst and subsequent telephone conversation with Mr Stockton, we understand that you now wish to use the inner, cage and rollers of our MRJ.12 instead of LRJ.17and we are therefore enclosing a print giving the necessary information for the 12mm bearing. Includes related drawing and letters

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5214SR18. Revision of stresses given in 5214SR187 September 1953 for shear neck failure condition. Schemes and drawings used OLS.1793 (B.87250)

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5214SR/18. Stressing of gear shafts and determining of bearing loads

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5214SR/18. Stressing of gears, bearings and shafts in generator drive increaser gearbox (in nose bullet)

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5214SR/18. Stressing of shear coupling and drive shaft for P3 generator (6kw) in nose bullet. Drawing used OLS.1626

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5214SR/18. Stress clearance. (a) Normal running 33.5hp at 9870 generator rpm. (b) Shear neck failure 420lb.ft. The bearing arrangement has been examined by the manufacturers and was considered to be satisfactory

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5214SR/18. Schemes OLS.1795 (B.75242)

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5214SR/19

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5214SR/18. Memo refers to working centre distance for both pairs of gears is 2.773 (see casing drawing B.77846). It then goes on to refer to five drawings for the future

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5214SR/18. From running experience it is apparent that the tip and root relief should be increased. No drawing available

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5214SR/20. The report goes on to examine the strengths and weaknesses of the above parts. It goes on to give details of the stress cage and torsion factor summaries

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5214SR/20. Contents: calculation of stresses; Conclusion and summary

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5214SR/20. We have examined scheme OLS.3582 showing the strengthened starter drive quill shaft and find it satisfactory from the stress viewpoint

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5214SR/20. Stresses in bearing housings whenfitted with PH.B liners. Schemes or drawings. Used:OLS.3770; OLS.3667; & OLS.3769. Calculations 1 to 6; Tables of stresses and factors 7

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5214SR/20. Strengthened starter drive with 200lbs.ft shear neck. OLS.3528. Contents: Explanatory notes (Pages 1-3) on use of 200lb.ft shearneck. Routine stressing sheets (pp 4- 13). Summary sheets as follows: Shafts - factor; Serrations - B.pressures; Gears - stresses (pp14 - 17). Gear teeth and serrations added

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5214SR/20. Olympus 103 fuel pump, hydraulic pump, starter; main oil pumps; auxiliary scavenge pump; and tachometre generator drive. Scheme OLS.3381; OLS.3378 (subsid). Arrangement drawings: B.75018 (inner end); B.102407 (Outer end)

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5214SR/20. Torsional stresses, serration bearing pressure, and flexibility calculations. Scheme drawings used: OLS.4078 starter clutch drive with improved flexibility

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5214SR/20. Comparison of strengths; sleeve/shear neck. Scheme/drawings used: B.80375 unit; B.54219-20 & B.80376 details

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5214SR/20. Fuel pumps, hydraulic pump, starter, main oil pumps, auxiliary scavenger pump, & tachometre generator drive. Routine stressing, bearings, gears, serrations, shafts. Routine stressing sheets, tables Pages 1-47; I to XIII inclusive p48; Correction factors p49. Stress summary filed separately under 5214SR/20. Clearance note 5214SR/20. See also 5214SR/20 July 1958

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5214SR/20. Gears, bearings, shafts etc; Rouine stressing to SM.63. Scheme drawings used: OLS.3378, OLS.3381, OLS.3582, OLS.3834, OLS.3948, OLS.3953. Tables and diagrams to SM.63 Pages 1-66

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5214SR/20. Torsional stresses, serration bearing pressures, and flexibility calculations. Previous calculations 5214SR/20 OLS.54078 dated 25 May 1957. Flexibility of vertical shaft and spindle. Pages 1-2; Routine stressing sheets, tables I, II, VII, VIII, IX, XII; Summary sheets XIV, XVI

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5214SR/20. Scheme drawings B.75018, B.133483. Previous calculations 5214SR/20 dated 2 May 1956 and July 1958. Routine stressing sheets SM.63 1 - 14. Summary sheets 15 - 19

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5214SR/20. Stressing of bevel gears (comparison with existing design) OLS.3953. Table I Case tabulation Page 1; Table XIII clear tooth stresses page 2; Existing starter gears Table XIII clear tooth stresses Page 3

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5214SR/20. Routine stressing, gears, bearings, shafts, serrations. Scheme B.75242. Previous calculations: 5214SR/18 dated 13 January 1953, and 17 September 1953. Routine stressing sheets pages 1-14; Flexibility of drive Page 15; Strength of planet gears layshaft Page 16

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5214SR/19. TP.3130. Prior to being used for the engine test programme TP.3095 series, the intermediate casing and associate parts ex OL.102007 (6th build) are to be made available for the test

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Engine 200835 at Drax 1005 hours 2nd root seration. Prim Y or SecY 5. Blocked NGV & one CC failed to light. Engine 200627 Drax 609 hours. Top root serration. Prim Y or SecY 5. Displaced NGV. Engine 200602 Ferrybridge 211 hours. Top root serration. Prim Y or SecY 5. Displaced NGV. Engine 200614 Ratcliffe 499 hours Top root serration. Prim Y or SecY 5. Displaced NGV. Engine 200638 Ironbridge 1048 hours. Top root serration. Prim Y or SecY 5. Suspect displaced NGV. Heavy damage. Engine 200814 Drax. 1048 hours. Top root serration. Prim Y or SecY P? No blockage. Being investigated. Engine 200807 Rye House 904 hours. 2nd root serration. Prim Y or SecY 5. 3rd flare break-up. Graph under report showing overall failure rate 7/46.5 or 0.15/1000 hours.

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Standardisation sheet 6178. Proposed types affected: Olympus 106, 108, 110, 113, 115, 116, 119, 120, 133, 134. To restrict blade untwist and fretting by replacing existing blade, which has a shroud gap of .000/.002 clearance, with interference fitted shroud of .000/.006, and puddle welded shroud abuttment faces of Satellite 12. An overhaul of approx 50% of blades require salvaging. It is considered the introduction of ineterference fitting shrouds will restrict blade vibration, and thus the rate of shroud edge wear, fiving at leat 500 hours in gain life expectancy. Modification D sheet 1 Approval: All the above plus 112, 112/RN.114. To be introduced with future contrscts. Will be available in 8 months. D sheet 2: Additional rework information introduced. Ref IPN 12742/5/C to facilitate early embodiment of the important features of this modification, blades procured to OLP2701 (S.00) can be mixed with blades to this mod standard. The only difference being OLP2701 instructs single pass S12. ten sets of blades to OLP2701 standard are currently available in IMD 75 stores. Mod OLP3411 (Eng). D sheet 3 GAPL section drawing.OLP 3411 (Eng): Schedule 4711007113 turbine Part 1 DIS 106, 113, 116. Schedule 4711006393 turbine Part 1DIS 108, 115. Schedule 4711021345turbine Part 1: DIS 112,112/RN. Schedule 4711023701 turbine Part 1: DIS 114. Schedule 4711008321 turbine Part 1: DIS 110. Schedule 4711040390 turbine Part 1: DIS119, 133. Schedule 4711040633 turbine Part 1: DIS 120, 134. Turbine rotor blade BDA9535 may be reworked to BDD5210 to make BDD5202. Turbine rotor blade BDA9542 may be reworked to BDD5202 standard using Drawing BDD5210.

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Ref: WKA/DEJ: This disc was fully machined at Derby and has no connection with any other failure report to date, in that it was produced approximately three years earlier than discs which have been subject of recent investigations. The report covers the three slots identified by Defects Investigation, including Slot 92 where the failure was believed to have occurred. To conclude, it is my opinion that the disc was not responsible for this particular failure.

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Ref: JLE/DTL/5/41/72-50-06. Tabulated list of stators etc.

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Drawing Nos.03-4545/-/4. The present standard of Olympus Mk.200 series and Mk.301 HP turbine rotor blades is to modification OLY.1488 standard. To date, there has been a total of 14 blade root failures in Service engines and a further five engines have either suffered blade failures or have been found with cracked roots during acceptance testing or engine repair. Three of the failures were experienced on blades to modification OLY.1488 standard. The failures occurred in the top or second root serrations, usually moleating at the acutely angled trailing edge concave side, or less frequently, at the leading edge convex side of the blade root. This then propogated in fatigue, resulting in fracture across the serrated base section. Tests by IMD indicate a loss of 1/2% thrust. Performance will be checked on first 12 engines

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Hand-written (in pencil) report from Aero Development Department Parkside. During the period, April 1963 to April 1968, there has been a total of 17 fatigue failures of 1st stage turbine rotor blades in service flight engines. These blades fail at one of three positions: 1 In the aerofoil from approx 1' to 1.25' above the platform. 2 In the top fir tree root serration. 3 In the 2nd fir tree root serration. The instance of these failures in both Mk.200 and Mk.300 series engines is as follows: a) Form failures Mk.200 Mk.300 Form failures 3 1 Top root failures 1 6 2nd root failures 2 4 Total 6 11

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This report gives a complete breakdown of all disc and blade failures up to 1970

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An earlier memo dated 13 June 1968, gavean estimated variation of LP/HP rpm, based on production acceptance figures, altitude effects being determined from early flight testing in the Vulcan aircraft (1962). In order to verify these figures for the latest standard of engine, a flight trial has been carries out on Vulcan XM606 instrumented to record HP rpm over the flight range. The results of these tests (Made on Mk.301 engine ), shows that the altitude effect is not quite as much as originally estimated, and it has been assumed that a similar effect would be applicable to the Mk.201 engine

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Hand-written report. 1 Thick disc but without differential pitch ie1570 standard. Engine 650224. 17 hours. Top root serration. Engine 650447. 962 hours.Top root serration. Engine 650393. 689 hours. Top rot serration. The failure rate based on the above three failures and 19,620 hours flying before withdrawal of the disc was = 0.152/1000 hours. Thick disc with differential pitch ie 1643 standard. Engine 650039. 813 hours. Top root serration. Engine 620266. 4 hours. Top root serration. On test bed. Engine 650306 1014 hours. Top root serration. Engine 9000003 1660 hours. 2nd root serration (+480 hours in other engine). The failure rate in Service based on the two failures and an estimated 600,000 hours flying Service embodiment of the Mod is = 0.0033/1000 hours. If the test bed failure is included then the rate increases to = 0.005/1000 hours.

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Contract K/47b/65: K/42b/93.Development Engineering Report No.253: Comparative fatigue testing on HP turbine rotor blades machined by BED and Hepworth & Grandage, using HP spool Rig. Engine type HP Spool Rig No.MRE.2: Builds 25, 26 & 27. This report covers a three-part test programme that was arranged following a series of short term HP turbine rotor blade fatigue failures on the latest standard IMD engines. A total of 12 root failures occurred at running times between 156 and 666 hours

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The continued incidence of Olympus LP turbine blade failures has led to the need for restricted operation in the 95+-1.12% rpm band where resonance and consequent fatigue of the LP blades can occur. In order to minimise the effects of this restriction on the flexibility available to pilots in setting rpm, the existing restricted bands are lifted on engines to post Olympus modification 1570 or 1643 standard. This relaxation has been made possible by the majority of engines in service, now have one or other of the modifications which introduce a stronger HP turbine disc

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Ref: E/FGPG/JDW/TN. Laboratory Report LE.78/158. This blade had fractured transversely above the top root serration as a result of fatigue cracking which initiated at the concave trailing edge corner of the root. The blade material loading appeared to have been satisfactory. The blade material was found to be metallurgically satisfactory

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Ref: E/FGPG/JDW/TN. Laboratory Report LE.78/191. During acceptance testing of engine 620266, fracture of an HP turbine blade root occurred after 4 hours. Subsequent NDT confirmed that two other blades were cracked at the root position. The three blades were submitted to the Laboratory for investigation into the failure. The three failed blades, reference 16, 13 and 8 were identified by the code HWTO. Ten further blades were used in the manufacture of the turbine and these were also obtained for examination

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Ref: RH/16/08. Mr Armstrong (the Chairman) reviewed the disc measurement as follows: The disc had ben turned at Dunlop and broached at Derby in 1976. The investigation proved that the correct broach had been used and was not a repeat of the situation in 1973 when 80 discs had been scrapped off. The measurements had shown the throat dimension to be undersize and secondly the across wires dimension was incorrect with shortage of metal on the pressure flanks on one side of the fir tree slot. The amount of error and the position in which it occurred was consistent, with the discrepancy that had been found previously, when the broach has been taken through the same slot twice. The stagger would also be affected

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Hand-written un-signed report. 1 Critical rpm band indicated by MES review is 83 - 91% Nh. This is coincides with the IF-6th engine order where the max AR was 55%. 2 The original SF1, which was introduced to avoid the 4th engine order combined disc and blade mode of vibration involving the 'thin' disc, avoided 81-87% LP pm in the case of he Mk.202 engine and 82.5 - 90% LP pm from the Mk.301 engine (ie rpm at 15C deg). 3 The equivalent LP rpm from the HP rpm band given at 0 is as follows:- Mk.202 70-83% LP rpm; Mk.301 69.5-83.5% LP rpm. 4 The failures to date have all been on Mk.301 engines (ie 3 cases in Service) and the bias has always been towards the Mk.30. May be due to more excitation ie higher operating pressure (HOP). 5 Because of 4 it would indicate that the fit of the blades at the root and the outer shroud is not the critical feature leading to the failures, but must be related to the rpm at which the blade is likely to resonate plus the (HOP). 6 Since the recent failures have occurred after the SFI was withdrawn, this also would tend to support comments at 5. 7 Since the likely speed band to avoid is now wider than before (ref 1 and 3) this may lead to compliance problems. Also, most of the band would have been flown in prior to the withdrawal of the SFI and apparently did not produce failures. This would also indicate that the area to worry about is the overlap and this is indicated below: Drawing of Mk.301 showing old SFI and Mk.202 showing new band. The overlap of the Mk.202 is 81-83% LP rpm. The overlap of the Mk.301 is 82.5-83.5% LP rpm. Using the above speed bands the HP rpm. Avoided is 90/91% for the Mk.202 and; is 90.5/91% fo the MK.301. To allow or scatter, the band for both marks could be made 80-85% LP rpm, which would avoid HP rpms between 89/92.5%. It could also be argued that since with the original SFI, there were no failures, that this should be re-introduced. This would encompass the top end of the 6h EO resonance condition covered by the ovelap condition. Olympus engine operation: In circuit 35 67/70% rpm. Approach 75 75-78% rpm. Roller 93% rpm. At below 5,000 not operated between 80-90% LP rpm.

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Report PS276/75. To Jo Pemberton. In view of the failure rate of blades in the Mod.1570 thick discs compared with the original thin discs, an investigation based on dimensional checks has been made. This entailed comparing the slot dimensions of a disc which suffered a blade failure with one without failure. The results including drawings and graphs re. contained in this report

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For the attention of Jo Pemberton: Investigation disc No Code 379. Serial No CJ635. 4 broached slots equi-spaced checked. Number 1 slot taken at random, and numbered in clockwise diection. Details see attached sheets

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Internal memo Superintendent of Metrology, Producer Centre 6 Ref JM/QK

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Ref RH/16/08. Jo Pemberton Project Engineer AQD, Ansty. Subject to the meeting held 8 February 1979 Ref RH/16/08, report has now been issued by headquarters Quality Engineering

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Laboratory report No. LE/79/120. This turbine blade had failed through the top root serration in the blade form. Photo Fig 1 View of blade fragments. Fig 2 Close-up of root fracture. Photo P6269 close-up of Fig 2. Photo P6270 righr hand side close-up of Fig 2.

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Internal memo Superintendent of Metrology, Producer Centre 6 Ref JM/QK. Blade root drawing B409197, cracked blade. Two Photos: 1 P6379 close-up of blade No 29. Photo 2 P6380 close-up of other side

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Ref: 443C SR/37. The engine, which retained the shroud rings from the previous test, had one of its 2nd stage turbine blades machined in accordance with OLS.4404 Issue 2 (serrations partly removed - the Armstrong Siddeley Motors method), so that it would pull out of the disc

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Ref 4434/37. OL TR 2068. Contract No. 13001/KG/M/011/CB12(a) Engine No 101030. Olympus Mk102. Build 11. Turbine containment test

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Ref 4434/37. OL TR 2068. Contract No. 13001/KG/M/011/CB12(a) Engine No 101030. Olympus Mk102. Build 12. Turbine containment test

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Ref: 4434SR/36. Calculation of radial changes in N labyrinth members for Olympus Mk.101, Mk.102 and Mk.104. Calculations attached

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Ref: 4434/35. Scheme drawings D/OLS3772. OLS4366. Drawing MS2543 of Front diaphragm (cooled turbine). Pages 1- 4 Calculations considering coned portion of cover plate. Page 5: Application of hoop & radial loads on cone to the hub section on Donath Chart. Page 6-9: Deflection due to rotation. Page 10: Rorational stresses on hub from Donath Chart. Investigation of cone under pressure conditions D/OLS/3772

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Ref: 4434SR/34. Scheme drawing OLS3560. Page 1 Graph of stresses. Page 3-4 Results from 'deuce'. Page 5 Section details. Page 6-8 Effects of offfset

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Ref: 4434SR/34. Streses in diaphgram with bayonet fixing. Scheme drawings OLS4563. Previous calculations

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4434SR/31. Scheme drawings OLS3382. Previous calculations 4444SR/15. Page 1: Individual stresses. Page 2 Total stresses. page tabulated stresses. page 4 Cross sectional view. Page 5 Donath Chart tabulations. Pages 6-7 Themal stresses. Pags 8-10 Spigot loadings. Page 11 Growth at labyrinth diaphgram. Page 12 Pressure differential

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4434SR/29. Scheme or drawings OLS2803. Previous calculaions 4444SR/9 & 4432SR/2. Page 1 Tabulation of stresses. Page 2 drawing GR3703 of interstage diaphgram. Page 3 Graph of interstage diaphragm (thickened web). Page 4 calculations of rotational stresses (no rim load; no side load) Donath Charts. Page 5 Calculation of thermal stresses. Page 6 Estimation of thermal stresses, and stresses due to external loading. Page 7 Side loading 1000lbs/in.

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Ref: 4434SR/29. Fretting at bayonets, growth of labyrinth face diameter and inner spiot, and difficulty of removal from 1st stage turbine wheel

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Ref: 4434SR/25. Photoelestic report PE.142

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Ref: 4434/11.Referring to your note Ref CE/EW-D?CH/2151 dealing with the inspection and detailing of blade roots and slots, it appears that up to a variation in inspection figures obtained by Experimental and Production Department when viewing the same turbine disc, that some looseness is suspected in the dimensioning of the drawing

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Ref: 4434SR/11. Notes of a meeting held in Mr Hill's office on Wednesday 19 January 1955 following the failure of Olympus mk.101A engine No.520. The substance of his report indicates the proposed design action to improve the cooling of the 1st and 2nd stage turbine wheels

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Ref: 4434SR/11. The first set of Olympus Mk.101 HP turbine blades on an engine which has run satisfactorily has now been checked in the Bartlett gauge for fir tree root errors. Mr Oliver has confirmed that future blades will have a stagger error between the two sides of the fir tree as recorded by the Bartlett gauge of +0.00025. This figure is considered less than the maximum measured on the set of blades ex Olympus 520

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Ref: 4434SR/10. After three builds on this engine rubs were found to have occurred between the inner platforms of the first stage turbine stators and the wheel blades. After the first build only, the air baffle attached to the seal ring support cone also rubbed the turbine wheel severely enough to wear through part of the baffle. The total running times on the three builds were 6 hours 49 minutes (18 November 1950 to 23 November 1950); 4 hours (20 December to 21 December 1950) and 2 hours on 7 January 1951

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Ref: 4434SR/8. Fatigue cracking in the 2nd from bottom root serration of the 1st stage turbine wheel has ocurred on Olympus Mk.100 engine 507, after a running time of 702 hours 46 minutes. This is the first failure of his nature on Olympus Mk.100 engines. The engine failed on 7 January 1956. Failed parts B.62396 1st stage turbine wheel, serial No.396 in H.46 mateial to OLS.745. One blade root slot was cracked in the 2nd from bottom root serration, at the front face acute corne. The crack as due to fatigue which had initiated from a sharp corner formed where the serration corner radius coinsided with the undercut on the front face of the whel rim

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Ref:4434SR/8. Stress clearance. The above scheme has ben examined and is considered acceptable from the stress point of view. It is estimated that the modification to the blade root decreases its strength by about 7%

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Ref:4434SR/5. On the 1st stage turbine fitting bolt, B66567, was found to be fractured in the undercut between the threaded section and parallel section of the bolt

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Ref: 4434SR/5. Notes on a meting in Mr Chamberlain's office to discuss the material of the turbine disc bolts, following the failure of a 1st stage disc bolt on Olympus 100 engine 2602. The material specification of the 1st and 2nd stage turbine disc bolts for Olympus 100 and 101 series engines is BACE.236 Issue 1

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Ref: 4434SR/5. It has been bought to our notice that the four air holes drilled in the above shaft are not provided with a blend radius at the intersection of the holes and the wall of the shaft. We would point out that in our assessment of the strength reduction due to holes are suitably radiused. An absence of such radiusing gives rise to a source of weakness in the shaft which becomes considerable in the event of stress fluctuation due to torsial vibration and lateral bending movements

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Ref: 4434SR/5. Investigation into the desirability of radiusing air bleed holes. Contents: 1 Determination of mean clear stress across critical section of hole. Four pages of calculations and drawings

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Ref: 4434/5. Gyro effects and tightening torques. Five pages of calculations and drawings. Includes related addendum

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Ref: 4434SR/4. Turbine wheels in H.46 and Rex.448 material can be regarded as stable in so far as overspeed testing is concerned, and therefore not subject to changes if the size of the slots alter broaching. Olympus wheels in G.18B can be looked upon in the same light since by the time they are broached they will have been already overspeed tested and therefore, in part finished condition. In the above cases therefore, ie, all wheels in M.46 and Rex.448 and G.18B Olympus only, the grading can be done at the first view after completion of broaching. The aim in this is to enable blade production to proceed without waiting for final completion of the wheels

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Ref: 45SR/6. CE/SWM/47425.Bearings numbered 1 to 7 from front to rear

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Ref: 45SR/6. CE/SWM/47425. Issue 2/2. Bearings numbered 1 to 7 from front to rear

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Ref: 45SR/6. CE/ SWM/E/51393. Drawing ES.999 Issue C, mentioned are not in our possession, nor any other drawing in this report

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Ref: 45SR/6. CE/HCM/S/51928/S

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Ref: 45SR/6. P/U.10/G/2. The necessesity of pressurising the compartment forward of the LP compressor, in order to reduce the thrust loading on the bearing, has been investigated for maximum RPM, ICAN, static, and 600 mph, Arctic conditions

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Ref: 45SR/6. CE/SWM/E/53924. We do not have the drawing mentioned

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Ref: 45SR/6. P/U.10/G/3. A means of maintaining the above bearing load within permissible limits, without resorting to pressure balancing on the HP compressor, has been devised as follows: Three labyrinths to be provided at the rear of the HP compressor, having diametres of 22', 17' and 7' approximately. Each of the two labyrinth chambers so formed to be vented to atmosphere via passages of not less than 3.0 sq ins, with blow-off vents opening at 70lb sq in. also located in the vent(s) from the outer chamber (ie between the 22' and 17' dia labyrinths). The table below summarises the required dimensions for the larger diametere labyrinths. This report is in very poor condition and must be handled extremely carefully

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Ref: 45SR/6. P/U.10/G/3. A means of maintaining the above bearing load within permissible limits, without resorting to pressure balancing on the HP compressor, has been devised as follows: Three labyrinths to be provided at the rear of the HP compressor, having diametres of 22', 17' and 7' approximately. Each of the two labyrinth chambrs so formed to be vented to atmosphere via passages of not less than 3.0 sq ins, with blow-off vents opening at 70lb sq in. also located in the vent(s) from the outer chamber (ie between the 22' and 17' dia labyrinths). The table below summarises the required dimensions for the larger diameter labyrinths. This report is ia properly typed report of the previous record

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Ref: 45SR/6. CE/JC/CAG/54322. It is agreed that a programme of bearing development rig testing is to be initiated and continue for Olympus bearings in the Mechanical Rig Shop in the same manner as for Proteus and Theseus bearings

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Ref: 45SR/6. TED.113. Report

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Your ref: 45SR/6. CE/SWM/E/53924. We have now finalised our consideration of the bearings for the above application as shown on your drawing MS.999 Issue D and give below our comments on the various bearings

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Ref: 45SR/6. CE/SWM/E CEB. The pincipals of the design described in PAN-107 are adhered to, but various dimensional changes have been found necessary since the date of the previous note, in order to ensure that the labyrinth blow-off valve closes before 20,000ft, altitude is reached, to restrict oil leakage from No.4 bearing, and to meet the bearing load limitations

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Ref: 45SR/6. EW/TPD/DD/MB/25. A test assembly in connection with scheme SCG.645 (No.4 bearing) which will cater for radial loads in addition to the thrust load previously mentioned. The normal radial loading device is capable of applying a maximum of 5,000lbs per stacked test pair and this will, no doubt, be adequate for your requirements

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Ref: 45SR/6. EW/TPD/DD/MB/25. The recipiant of this report has underlined the following: gyroscopic action in flight will apply radial loads considerably in excess of those previously envisaged

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Ref: 45SR/6. CE/SWM/WC/60828. Par 3: It is to be noted, that the first tests to be carried out will be the thrust balancing device tests as outlined on SCG.645, and furthermore the parts cleared by the department necessary for such test have been completed since 15 February 1949. Consquently as soon as the parts put in hand by your department become available, testing will commence

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Ref: 45SR/6. CE/F/G/61457

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Ref: 45SR/6

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Ref: 45SR/6

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Ref: 45SR/6. Forthcoming visit to Bristol to discuss the outstanding bearing problem on the Olympus

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Ref: 45SR/6. Proposed visit of Messers Storer and Rutherford to Bristol aeroplane Company to discuss Olympus bearings

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Ref: 45SR/6. CE/SWM/47425. Issue.3. Bearings numbered 1 to 8 from front to rear

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Ref: 45SR/6. Telephone conversation re. proposed visit of Messers Messers Storer & Rutherford to Bristol Aeroplane Company re. Outstanding order for Olympus bearings

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Ref: 45SR/6. RW/TPD/DD/M/79. Since receipt of our memo of 10 January 1949, a complete redesign of Olympus bearings has taken place and the design of test assemblies has consequently been held up. I have now obtained a print of MS.1054, which supercedes MS.999, and gives particulars of the redesigned bearings. Design of the test assemblies can now proceed. It has been agreed that bearings Nos.2 4, and 8 positions will be tested and first priority will be the No.2 stacked pair with thrust balance device. Existing assembly RM.999 will cover the No.8 position. The shaft for the No.2 position will probably suit the No.4 position when fitted with adaptor sleeves etc.

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Ref: 45SR/6. EW/TPD/DD/MF. Para 4: Radial and thrust loading devices will be applicable in all three cases and it should be noted that the maximum capacity of the existing rig is:- Total thrust load 15,000lbs. Total radial load 10,000lbs.

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Ref: 45SR/6. CF/F/G/65081. Attached are two pages of all Olympus ball and roller bearings that have been ordered up to and including 29 June 1949

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Ref: 45SR/6. This is a hand-written report which goes into details of the affair, however, it is unsigned

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Ref: 45SR/6. CE/SWM/WC/6577. The position regarding Olympus bearings has now been clarified, and as a result, does not line up with your list issued on 28 June 1949 ref: CE/F/g?65081, and for your information we have prepared lists showing the bearings now cleared under the various schemes, from which you will be able to amend the respective assemblies as follows:- Oil sump delivery assembly No.302. Compressor rotor assembly No.304. Starter drive assembly No.314. Fuel pump drive assembly No.321 etc.

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Ref: 45SR/6. CE/SWM/E/66938. We enclose herewith a copy of our Design Scheme OLS.216, showing our latest proposals for the above bearing, which we are specifying in tungsten-chrome steel, fitted with a granodised steel cage and made to precision limits of accuracy. The loads and running conditions will be as shown on our Scheme MS.1054, already in your possession. We shall be pleased to receive our comments on the above and a copy of your preliminary data-sheet. A modified copy of our Scheme MS.1054 will be forwarded to you in due course

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Ref: 45SR/6. CE/SWM/E/69201. It will be noted that we propose to supply the bearing with a convex outer track and a diametrical clearance of .0045' to .005.' This we think should be adequate if the temperatures given in your earlier scheme are likely to be retained

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Ref: 45SR/6. CE/SWM/E/69201. You will notice that some of the bearing race temperatures have been modified; this has been done to bring them into line with your latest temperature estimations. Attached, hand-written note of questions: What are changes introduced by his (as distinct from bringing records into line with revisions already notified to R & Ms). Temperature changes and insignificant rounding-off of load figures

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Ref: 45SR/6. CE/SWM/E/71021. Scheme MS.1054 Issue B, which gives the latest particulars and running condition to the above bearings. This scheme superedes MS.1054 Issue A, any copies of which should be destroyed

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Your Ref: 45SR/6. CE/SWM/E/69201. Thank you for supplying a copy of Scheme MS.1054 Issue B, giving up to date particulars of the runnng conditions for the bearing as above. In view of the revised temperature figures, we would suggest that the diametrical clearance should be increased on No.5 bearing your part No. B.29579 to .0055' to .006', this will bring it into line with the fit of the outer bearings and take into account the increased temperature

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Ref: 5SR/6. CE/SWM/E/71641. We are accepting your recommendation to increase the diameter-clearance of No.5 bearing

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Ref: 45SR/6. List of tables

Ref:45SR/6. list of velocities

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Ref: 45SR/6. DCE/RCF/272. We should b glad to know your delivery promises for the main shaft bearings for which you have not already quoted

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Ref: 45SR/6.DCE/SWM/KC/10. In connection with the first test run of the Olympus, Messers Ransome & Marles have brought to our notice that the thust bearings at positions No.2, 4 and 8 will fail if the thrust load is less than the radial load. As it is poposed to 'run-in' the first engine at a low rpm, we would like to know the thust loads which will be imposed on the aformentioned bearings, for a range of speeds up to 4,000 pm (LP compressor), in order that we may advise the test engineers regarding suitable speeds

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Ref: 5SR/6. DCE/FCM/EB. Recent changes in concessions have been made to the main shaft bearings for Olympus engines, and in order that the position may be claified, we give the following list showing part numbers, with any alterations which may be used on the first engines

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Ref: 45SR/6. TPR.8002. Lighting up tests were carried out successfully and a good start with little flame was obtained by motoring up to 600 rear com rpm; lighting up and then motoring up to full speed on the starter rotor. The maximum jet pipe temperature recorded during this start was 400degC

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Ref: 45SR/6. DCE/SWM/KC/135. We thank you for your letter of the 7th inst and closing a copy of Scheme MS.1154 showing a proposed bearing arrangement for the above application. We do not approve of the proposal to make the inner race integral with the gear. This makes a very complicated component which would be unnecessarily difficult to manufacture. We note that you suggest a material similar to DTD.551 Nitrided but would point out that we understand that this material has failed quickly in the test bearing which you have recently tried and we therefore doubt whether it is suitable material for producing bearing races. We would strongly advise you to use a bearing of standard shape and we feel that by a slight re-design around this postion of the job, you could accommodate our XLRJ.4'E, the overall dimensions of which are 4'by 5.5/8' by 7/8' and we would most strongly recommend you adopt this course

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Ref: 45SR/6. Mr Rutherford (Ransome & Marles) confirmed that the above bearing as shown on the peliminary scheme handed to him on 28 June, would be satisfactory

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Ref: 45SR/6. DCE/SWM/KC/CEB. Following the attention made to N labyrinth on the HP turbine, the diameter of G labyrinth can be reduced

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Ref: 45SR/6.DCE/SWM/KC/CEB. Due to changes in design it has now been found possible to reduce the diameter of the C labyrinth

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Ref: 45SR/6. I promised to send you particulars of our Olympus bearings in order that you could put up proposals for Hoffman alternatives. These bearings fall into two general classes 1 Main shaft bearings; 2 Accessory drive bearings, all have steel cages. I am, therefore, attaching copies of drawings which have been prepared by Mr Chamberlaine, giving 1 general arrangements of each bearing, together with the operating conditions, and 2 Detailed drawings of existing bearings, where they would be required by you for information on overall dimensions

Ref: 4366SR/3. Contents. Pages 1-3 calculations. Page 4 Graph axial gas loading. page 5 Tangental gas loading. Page 6 to end calculations

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Ref: 4366/2. Pages 1-3 calculations. Page 4 Enlarged annulus area; tangental gas loading. Page 5 Enlarged annulus area, axial gas loading. Pages 6-25 calculations

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Ref: 4366/2. Stresses in blade root fixings including gas bending effects and stresses in 5th stage disc lobe including asymmetric loading effect. Contents: 4 pages of calculations and drawings

Ref: 4366/2.Pages 1-13 calculations and tables

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Ref: 4366/3. Proposed new root fixing broach. Previous calculations 4366/3. Contents: Pages 1-9 calculations and drawings

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Ref: 4366/2.Proposed new root fixing broach. Previous calculations 4366/2. Contents: Pages 1-7 Calculations and drawings

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Ref: 4366SR/2. Recent fatigue cracking of Proteus III fir stress disc lobes, resulting in one case in complete fracture of the lobes and consequent destruction of the compressor, indicate that the fillet stresses in the disc lobes are too high to give extended disc life. Calculations of these fillet stresses is complicated by the presence of interminate vibration loads, the value of which can only be properly obtained by running experience

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Ref: 4366/3. Effect on root stresses increasing root width by 0.50'(using oversize broach). Contents: One page of tables

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Ref: 4366/2. Performance calculations rederred to: 4366/2 & 4356/2. Contents: Page 1 Summary. Pages 2-3 calculations

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Ref: 4366SR/2. Recent fatigue cracking of Proteus III fir tree lobes, resulting in one case in complete destruction of the compressor

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Ref: 4366/3. Calculationa DJ Fortune. Performance office data used PDM.41. Previous calculations referred to 4366/3 & 4356/3

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Ref: 4366SR/2. Calculation RE Jenkins. Performance office data used GP/7 6896SR/2. Previous calculations referred to 4363SR/7. Olympus IV factors to be applied to blade stresses for different engine conditions. Reference GP/7 BOL.4 internal engine conditions and memo from P and A department giving Mach 2 internal engine conditions. dated 22 October 1953. Table listed below gives a breakdown of engine conditions at various speeds

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Ref: 4366/3. This note gives general stress clearance to the HP rotor blades with section data as shown on 03S 131-4, Ref PDM.41, for the following flight conditions

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Ref:4366/2. This note gives general stress clearance to the LP rotor blades with section data as shown on 03S 241-4 Ref GN.593 PDM.42 & 114, for the following flight conditions: 1 ICAN Sea level static 2 ICAN Sea level 600 knots 3 -25deg.C Sea level 557 knots 4+25deg.C Sea level 630 knots 5 ICAN Sea level Tropopaue 500 knots Material should be BACE 245

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Ref: 4366/2. Scheme or drawing used PDM.220. Stresses in BOR.6 & all rotor blades. Blades physically as BOL.4 but with increased t/c tapering from 12% at the root to 5% at the tip

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Ref: 4366SR/2.BY KR Chamberlain. The enclosed tabulation gives stresses in the rotor blades for the design speed 6320rmp and for the 5% & 16% up-speed cases, also the adjustments to the blades considered necessary to give stress clearance for these latter cases

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Ref: 4366SR/2. GN 896

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Ref: 4366SR/2. This meeting was held primarily to review the proportions of the blade roots on the Olympus 4 - 6 LP compressor. These have been criticised in view of the apparent excessive depth of serration in comparison with the thickness at the neck of the blade root

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Ref: 4366SR/2. An investigation has been made into the use of the blade root form standardised for BE.25 and BE.26 engines, with the aluminium alloy compressor blades for the Olympus 6 engine

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Ref: 4366/2. Oversize broaches for LP discs

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Ref: 4366/2.By RSL Pritchard. Stresses in Olympus 6 LP compressor rotor blades under various aerobatic conditions. Previous calculations referred to 4316/2, 4366SR/2, 8 August 1952. Contents Page 1 summary; Evaluation of factors for the engine conditions pages 2 & 3; data given Page 3; Calculation of correction factors pages 4-8

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Ref: 4366/12. Cancellation of LP 1st stage rotor blades ordered on 03N.316. The order for the two sets of HP 1st stage rotor blades in BASE308 are to stand

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Ref: 4366/2. DH hughes & R Pritchard. Stresses of root fixings with 101 type roots. Note: The rim shapes are those of B 01.6D. Scheme or drawings used: 03S 273/14-13

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Ref: 4366/2. A report of the primary total failures on B 0L.601 was first given, this was the same as was reported in report ref CE/RCP/PJEH/18094 with no further additions. The results of fatigue tests on 4th stage blades show them to be weak in 2nd flap vibration which leads to the conclusion that the most probable cause of failure is 2nd flap vibration excited by 8 intermediate vanes at 5730 rpm, for the 4th stage blades. Because of exceptionally high resonance on the 4th stage blades it is quite reasonable to believe that the 5th stage blades failed off resonance, thus the blade is very critical

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Ref: 4366SR/5. The failure was analysed as follows: 1 Fatigue failure of four 1st stage blades; two blades had fractured through the top root serration; one blade was cracked in the top serration and one cracked in the bottom serration. One fractured blade had fatigued over about one-third of the area of the top serration. 2 Fatigue failures of three 2nd stage blades, one blade had fractured through the top root serration and was also cracked in the bottom serration

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Ref: 44366SR/5. Failure of the 5th blade roots and cracking of the serrated disc rim due to fatigue, derived from running at resonant conditions under a powerful excitation, is probably aggravated by local stresses arising from certain design features

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Ref: 4096SR/4. It should be noted that the Allen screw locking device applies only to the 2nd stage dummy weight, the positive location of which is necessary when overspeeding the 1st and 2nd stage rotor assembly alone

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Ref: 4096SR/4. The following overspeed assemblies are to be manufacured for test on the above rig: 1 B.01.4 1st & 2nd stage LP rotor assembly (minus blades). 2 B.01.4 1st & 2nd stage HP rotor assembly (minus blades). 3 Suitable dummy blade weights for these units are to be designed by the Stress Office.

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Ref: 4096SR/4. It should only be necessary to overspeed the first of these rotors. The balanced compressor assembly, complete with dummy blade weights to DT.291567-71 is to be overspeeded in the following manner. 1 Measure the outside diameter at four equally spaced stations around the predetermined circumference on the 1st, 3rd & 5th stage discs. Note the room temperatures. 2 Set probes over the following to a clearance of 0.040. i) 1st stage dummy blade weight. ii) Spacing ring between 3rd & 4th stage blade weights. iii) 3rd stage dummy blade weight. iv) 5th stage dummy blade weight.

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Ref: 4096SR/4. In view of the high loads which will be imposed on the bolts connecting the 1st & 2nd stage rotor discs to the rotor centre, it is necessary to carry out initial tests as soon as possible to ascertain the reserve factor of the bolts under static and dynamic conditions and the behaviour of the coned rotor assembly.

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Ref: 4096SR/4. We have examined your scheme RL.19824 showing the arrangement for overspeeding the HP & LP Olympus Mk.4 1st & 2nd stage compressor rotor assemblies. Incorporating our suggestions to make the respective rotor centres of the same material and as dimensionally similar to engine components we find this arrangement satisfactory. This is a memo from the Stress department to Mr Stevenson Engineering

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Ref: 4096SR/4. Since the behaviour, under running conditions of dummy blade weights retained by Allen type screws in a tapering root slot as present in the 2nd stage disc, in unknown, it is thought desirable to safeguard against axial movement of the weights, should the nip caused by the screw due to be lost due to rim expansion

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Ref: 4096SR/4. In view of the high loads which will be imposed on the bolts connecting the 1st & 2nd stage rotor discs to the rotor centre, it is necessary to carry out initial tests as soon as possible to ascertain the reserve factor of the bolts under static and dynamic conditions and the behaviour of the coned rotor assembly

Ref: 4096SR/4. In order to prevent the 2nd stage dummy blade weights from moving axially, a locking ring B.88789 is to be fitted behind the 2nd stage disk. With the fitting of this locking ring, the original 1/4 UNF bolts UB.83877 24-off will be replaced by 24-off UNF studs UB.80548

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Ref: 4096SR/4. Drawing of retaining ring for 2nd stage dummy weight

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Ref: 4096SR/4. List of tabulated checks and tests. Torque recommended 80ft.lb. + 5 using engine oil

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Ref: 4096SR/4. The B.OL.4 LP compressor rotor contains two coned assemblies, the first incorporating the 1st & 2nd stage discs, and the second, the 4th & 5th stage discs

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Ref: 4096R/4. 1 The overspeed tests had disclosed permanent distortion of the rim of the second stage disc which had heeled in a forward direction. The spacer ring flange had also run on the rear face. At design rpm, dilation of the second stage disc was more than twice that of the first stage

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Ref: 4096SR/4. Correction of existing weights in order that the CG corresponds to that of the blades. Mods using BACE 76 plugs. Mods using BACE bolts

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Ref: 4096SR/4. The combined 1st & 2nd stages were assembled to Schemes SCS1483and RL19824, and subjected to an overspeed test in accordance with the Stress Office instruction 4096SR/4 Issue 2

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4096SR/4. Memo from Stress Office: In view of increased abutment between the 1st stage dummy blade weight root fixing and the redesigned spacer ring of the 1st and 2nd stage LP compressor coned assembly, it is thought that the nine rejected dummy weights mentioned in your report of the 7th January 1954 ref ED/INS/KJ/IM/98. The dummy blade weights in question, should have their retaining tangs on the forward face made true to drawing, and would Mr Newport please make arrangements for this work to be carried out. The full 1st stage set of weights will be modified to our instructions detailed in our note to Mr Edmunds dated 12 January 1954 Ref CE/SWM/KC/3889

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Ref: 4096/4. Memo from Chief Engineer No.4 shop to Stress office: With reference to your note of the 12th instant, and confirming our conversation yesterday, concerning the modifications which you have requested shall be made to dummy blade weights to Scheme MS.1699, we would confirm that there are some 400 weights involved and that the cost of the modifications to these weights will be approximately £2,000 to £2,500. You may consider this to be somewhat excessive but we would explain that the heavy metal plugs which are required to be fitted to the blades will have to be split up from hexagon bar of fixed size, since this particular metal is only supplied in that form and, in addition, each of the weights requires jig boring or jig drilling for at least two holes. Drawing DT291567-71 attached

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Ref: 4096SR/4. The following procedure is to be adopted when overspeeding the Olympus 4 1st & 2nd stage LP coned assembly

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Ref: 4096SR/4. Further to our note of the 17th inst ref CE/SWM/KC/4031 regarding the rebuild and testing of the 1st and 2nd stage coned assembly in reverse, the following additional checks would be of invaluable consequence in the redesigning of the coned pair

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Ref: 4096R/4. It has been decided to carry out overspeeding investigations on the Olympus 4 LP compressor 4th & 5th staged coned assembly

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Ref: 4096R/4. It was decided at a meeting on the 16th February that spinning tests should be carried out on the 4th and 5th stage LP compressor coned assembly in view of the recent findings on the first two stages of this rotor

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Ref: 4096SR/4. The following procedure is to be adopted when overspeeding the Olympus Mk.4 4th & 5th stage LP coned assembly: Drawing CE/SWM/KC/4032 attached

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Ref: 4096/4. Following the overspeed test of the above assembly after the initial build, ref EDR.745/54, the unit was rebuilt in accordance with Research Programme Note 2129/B. The following new parts to scheme were incorporated...

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Ref: 4096SR/5. Section 1: Olympus 4 coned assembly - 1. 2nd stage dummy weights unsatisfactory. 2 Torque tests carried out on 20 - 7/16 bolts in test blocks resulting in 70lb.ft. 3 Tested to 7000 rpm when a 2nd stage weight moved axially about 1/4'. Test discontinued. 4 Although large elastic displacements were apparent, no permanent distortion was found in the discs. Section 2: Compressor rotor

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4096SR/4. Issued by Component Development Department. Aero Engine Division Bristol Aeroplane CO. Ltd. In view of previous tests on the B.OL.4 LP compressor 1st & 2nd stage coned assembly, which indicated that the rim of the 2nd stage disc was tipping excessively under centrifugal load, a re-designed spacer ring intended to give more support to the disc rim was incorporated in the 2nd build of the assembly. A series of spinning tests have now been acrried out to obtain dilation and axial displacement characteristics of the assembly under centrifugal loading

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Ref: 4096/4. Stress Office. In order to test a coned assembly of the 4th & 5th stage rotor discs schedule assembly B.91681 has been prepared listing the existing parts to be obtained from the rotor assembly, also a new part B.91862 Blade weight retainer (4th stage), which is to be manufactured. Test Plant Design Office will make an adaptor to suit these discs

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Ref: 4096SR/4. Dr Hooker has ruled at a recent meeting that the Olympus 4 LP compressor rotor shall be designed and tested to try light alloy blades (BACE 301) to replace the existing steel rotor blades in BAC 245

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Ref: 4096SR/4. Report T.561/54. OO.703102. Previous tests on the B.OL.4 LP compressor rotor 1st & 2nd stage coned assembly, have shown the unit to be incapable of withstanding the centrifugal loads imposed on it at proposed engine running speeds without undue distortion. Before revising the design of the assembly, it was considered expedient to obtain as much information as possible from the existing assembly, to supplement the theoretical basis on which design is based

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Ref: 4096SR/4. The Stress office reviewed the design position in the light of the latest decision to introduce aluminium blades in the LP compressor assembly. It has been found that whilst the material change offered some relief, the problem of elastic canting of the rims of the current compressor discs still persisted. The maximum axial movement at the blade tip occurred on the 2nd stage and amounted to approximately 3170 in. After reviewing the blade clearances it was agreed that this movement be tolerated for initial running

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Ref: 4096SR/4. 1. Includes related Engine Stress Investigation

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Ref: 4096SR/4. Attached full report with drawing

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Ref: 4096SR/4. 1st & 2nd stage dummy blade weights to refinement 10%T/C 7% T/C tip A1 alloy blades with no correction applied for rim weights. Previous calculations referred to: 4096/4. Drawing on last page

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Ref: 4096SR/4. Modified dummy blade weights representing 12% T/C root 5% T/C tip (3rd, 4th, 5th stages) 10% T/C root, 7% T/C tip (1st & 2nd stages) A1 alloy blades without correction to rim couples. Scheme of drawing used: MS.1571

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Ref: 4096SR/4. Enclosed is the overspeed schedule covering the investigations to be carried out by the Olympus 4 1st & 2nd stage coned LP assembly, incorporating the modified 1st stage spacer ring to 03S.273/7. 3-page report including drawing on final page

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Ref: 4096SR/4. Fourth interim report on spinning tests carried out on the B.OL.4 1st & 2nd stage LP coned assembly. Report T.573/54. OO.703102. Previous tests on the B.OL.4 1st & 2nd stage LP compressor coned assembly have shown the assembly to be incapable of withstanding the loads imposed upon it by the employment of steel blades at proposed engine running speeds; the centrifugal loads due to to the blade weights being out of line with the resisting forces in the disc sections, set up a couple at the rims of the discs causing them to tilt to such an extent that permanent distortion occurred at speeds below proposed engine take-off rpm

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Ref: 4096/4. The above assembly is now to be built with dummy blade weights to simulate aluminium blades. Details of the assembly are shown on Scheme 03S273/7

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Ref: 4096SR/4. Enclosed herewith is the overspeed schedule covering investigations to be carried out on the Olympus 4th 7 5th stage coned LP compressor assembly

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Ref: 4096SR/4. Following the overspeed test of the above assembly, after the 3rd build, ref EDR.534/4, the unit was stripped and rebuilt inacordance with Research programme 212/D

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Ref: 4096SR/4. The above is now to be built with dummy blade weights DT.527773 and DT.327774 to simulate aluminium blades. Scheme SCS.1641 shows details of the assembly with the exception of the dummy blade weight. Includes related Research Build Note RB.6/2. B.OL.4 LP 1st & 2nd stage coned assembly for overspeed testing: Ref: 4096SR/4. The above wheel is being returned to No.4 Shop following completion of the testing and is to be check balanced and dimensionally checked as laid down in the Stress office Schedule

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Ref: 4096SR/4. 1 First two stages LP compressor. 2 HP compressor - complete assembly.

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Ref: 4096SR/4. 1st & 2nd stage dummy blade weights to represent 10% root, 2% T/C tip titanium blades for curved discs (see MS1783). Scheme or drawings used MS1783, BT111101 & 3. Previous calculations referred to: 4096SR.4 30 April 1954; 4356/2 4 May 1954

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Ref: 4096SR/4. 1st & 2nd stage dummy blade weights to represent 10% T/C root, 2% T/C tip AL.Alloy (BACE 301) blades for redesigned discs to MS1756. Scheme or drawings used: MS1756. Previous calculations referred to: 4096/4 13 May 1954

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Ref: 4096SR/4. Further to note RB.6/1 difficulty in tightening the nuts on the hub bolts has been experienced due to the small diameter spot face on the adaptor; to overcome this, special stepped nuts are now to be fitted, therefore the UB.80554 special nuts 10-off, are replaced by BS.26023 Special Nuts 10-off supplied by TPDO. Scheme SCS.1641/2 has been prepared to show this change

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Ref: 4096SR/4. Enclosed herewith is the overspeed schedule covering investigations to be carried out on the complete LP compressor rotor assembly comprising Olympus 4 components. Includes related Fifth Interim report on spinning test carried out on the B.OL.4 1st & 2nd stage LP coned assembly: Ref: 4096SR/4. Although previous rig test have shown the B.OL.4 LP compressor rotor 1st & 2nd stage coned assembly to be incapable, at the proposed engine speeds, of withstanding the loads imposed by the steel blades for which it was originally designed, it was visualised that the substitution of aluminium alloy rotor blades might permit the use of the existing rotor design

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Ref: 4096SR/4. Dimensional check before assembly: A complete check should be taken of all major dimensions, together with the degree of coning on the forward faces, of the 1st & 4th and aft faces of the 2nd & 5th stage discs. Report T.603/54. OO.703130

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Ref: 4096/4. The above assembly has been returned to No.4 Shop after completing its tests and is to be balanced and dimensionally checked, as laid down in the Stress Office Schedule, after which it is to be completely stripped and inspected

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Ref: 4096SR/4. Total running time 22.5 hours. Issued by Component development Department, Aero Engine Division, Bristol Aeroplane Co. Ltd. Report T.630/54. OO.703130. Fifth Build Al blades

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Ref: 4096/4. Running time 5.5 hours. Report T.635/54

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Ref: 4096/4. The above new discs were subjected to overspeed dilation and checks to Schedule 4096SR/4 Issue. Upon satisfactory completion of tests, subsequent examination and dimensional checks revealed no appreciable radial or axial changes and all components were in a satsifactory condition. Summary of tests is attached in tabulated format

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Ref: 4096SR/4. Following individual testing of stages 1 & 2, and stages 4 & 5, the complete rotor unit is now to be built with dummy blade weights to simulate aluminium blades, reference scheme SCS.1641/2, 03S.273/3 and 03S.273/7

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Ref: 4096SR/4. It should only be necessary to overspeed the first three of these rotors

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Ref; 4096SR/4. Further to RB.6/5 the following modification is to be incorporated to enable the assembly to be built

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Ref: 4096/4. Further to TP.1801/10, the following modifications are to be incorporated into the initial build of this engine

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Ref:4096/4. Further to RB.6/5, it has now been decided to bring the above rotor in line with main engine by fitting the latest type dowels and studs at the spacer ring and rotor discs, stages 3-5, stages 1 & 2 rotor discs are already in line

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Ref: 4096/4. Attached: LP compressor 6,600 rpm 5th stage. Blade drawing B.118255

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Ref: 4096/4. Further to RB.6/7, please note B.79986 rotor disc stage 4, which was modified to drawing B89222, is now to be repart numbered B.101474ND, due to the 10-off holes being 37/64 dia. instead of .5625 dia.

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Ref: 4096SR/4. Dummy blade weights representaing AL. Alloy rotor blades with Mk.101 blade root fixings for 1st, 2nd, 3rd, 4th & 5th staged at 6,600 rpm. Schemes or drawings used: B.102381-5 blades; B.102386-90 discs. Previous calculations referred to: 4364/2 20 August 1954

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Ref: 4096/4

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Ref: 4096/4. It has been agreed by the Stress Department that the overspeed test on Olympus No.603 HP compressor rotor can be waived since an overspeed will be required on Olympus no.604's rotor with Titanium discs

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Ref: 4096/4. With reference to TP.1955 which instructed the build of the HP compressor rotor for overspeed testing, it has now been agreed with Stress Department that overspeed test on this rotor need not be carried out and the Instruction on TP.1955 is therefore cancelled. MoS technical agreement CE/RCP/PJEH/18335

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Ref:4096/4. Followin inspection, the above units is to be rebuilt to the following standards, which will change the title from B.OL.4 to B.OL.6

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Ref: 4096/4. Total running time 10 hours. MoS contract SP/6/Eng/7274

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Ref: 4096/4.Further to TP.1956/4, the following additional modifications are required during the initial build of this engine

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Ref: 4096/4. Three Olympus 6D oil separators have been successfully overspeeded to cover maximum engine take-off conditions at 6320 rpm. This process could now be deleted from the manufacturing process for these de-rsted engines

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Ref: 4096/4. On completion of crack testing of the LP compressor rotor discs, the LP compressor is to be built with dummy blade weights and despatched to the Research Department overspeed rig, Gypsy Patch, where overspeeding 10% over 6600 rpm is to be carried out to Stress Department instructions. On completion, the LP rotor is to be built for fitting to the engine as already instructed

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Ref;4096SR/4. The balanced Olympus 6D compressor assembly, complete with dummy blade weights to DT.327709-10 and DT.327772-4 in the first to 5th stage respectively, in the following manner...

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Ref: 4096SR/4. In view of the satisfactory overspeed experienced on the first Olympus 6D LP compressor rotor assembly, carrying representative A1 Alloy rotor blades, it has been ruled by Dr Hocker, that a similar rotor be further overspeed tested to 7260 rpm, to cover Olympus 6 take-off conditions of 6600 rpm. It has been agreed that on the next strip of Olympus 6 engine, the LP rotor will be built for overspeed and the test carried out in accordance with Stress Department Schedule Issue

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Ref: 4096SR/4. Enclosed, herewith, is the Stress Department overspeed schedule covering acceptances test on the Olympus 6D lp compressor rotor, carrying representative A1 Alloy dummy blade weights with Olympus 101 type root fixings in stages 3, 4 and 5. Dilation checks should be carried out over the last three stages on the first rotor only

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Section D. Issued by Performance & Aerodynamic department PIM.42. The compressor design is based on the following design point of assumption at ICAM. SLS conditions: Page A1: Air mass flow 212 lbs/sec. Compressor rpm 6,200. Compressor total pressure ratio 2.38. Page A2: Olympus III data LP compressor 6,200 rpm 212 lbs/sec ICAN sea level static (tables). Page A3: Olympus III LP compressor rotor blade data. Page A4: Olympus III LP compressor stator blade data. Page A5: Graph of Olympus III LP compressor rotor blades. Page A6: Graph of Olympus LP compressor stator blades. Includes related addendum

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Section D. PDM.41. The compressor design is based upon the following assumptions at ICAN. SLS conditions. Air mass flow 212 lbs/sec. Compressor rpm 7,800. Compressor total pressure ratio 4.0. Report also contains HP compressor tables: 7,800 rpm data at sea level. Stator blade data. Graph for HP inlet guide vanes; Graph HP compressor blades

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Section D.GN.399. DCE/NRO/847B/P. Drawing MS.1394: Temperature and pressure in turbine at ICAN sea level static maximum rpm entry. Total temp 1020CK. Drawing GR 2770: Inertia bending moment diagram. (REJ 2 January 1952). Drawing GR 2769: Inertia shear force diagram. (RJ 2 January 1952). Drawing GR.2773: Pressure thrust diagram. (REJ 4 January 1952). Drawing GR.2774: Torque reaction diagram. (RJ 8 January 1952). Drawing GR.2772: Gyroscopic bending moment diagram.(RJ 4 January 1952). Drawing GR.2771: Gyroscopic shear force. [RJ]. Graph GR.2803/A: Dead load shear force moment for Olympus III LP system. (KLB 21 January 1952). Sheet LP.1. Graph GR.2803/C: Dead load force & bending moment for Olympus III HP system. (KLB 21 January 1952). Sheet HP 1. Graph GR.2803/B: Gyroscopic shear force & bending moment for Olympus III LP system.(KLB 21 January 1952). Sheet LP 2. Graph GR.2803/D: Gyroscopic shear force & bending moment for Olympus III HP system.(KLB 21 January 1952). Sheet HP 2.

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4263/7. Section D. Graph GR2798: Olympus III LP compressor 1st stage disc and stresses. Graph GR2799: Olympus III LP compressor 2nd stage disc and stresses. Graph GR2800: Olympus III LP compressor 3rd stage disc and stresses. Graph GR2801: Olympus III LP compressor 4th stage disc and stresses. Graph GR2802: Olympus III LP compressor 5th stage disc and stresses. Page B7 to B12: Olympus III LP compressor rotor spacing rings. tables and drawings. Page B13: Compressor and turbine rotating parts. Polar moment of inertia. Page C2. Graph GR2812: Olympus III HP compressor 1st stage disc and stresses. Page C3. Graph GR2813: Olympus III HP compressor 2nd satge disc and stresses. Page C4. Graph GR2814: Olympus III HP compressor 3rd stage disc and stresses. Page C5. Graph 2815: Olympus III HP compressor 4th stage disc and stresses. Page C6. Graph 2816: Olympus III HP compressor 5th stage disc and streses. Page C7. Graph 2817: Olympus III HP compressor 6th stage disc and stresses. Page C8. Graph 2818: Olympus III HP compressor 7th stage disc and stresses.

Section C. Now part of 1716/5/11/2/427

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Section C. Basic scheme 03S 127/8. This note gives general stress clearance to the HP compressor rotor assembly as shown in the above scheme, with the exception of the blading which is the subject of a further note

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Section C. Two pages of tables

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Section C. Scheme 3S.127 Material S.106. ICAN SL Static

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SectionC. Component radial extensions at ICAN SL Static condition (relative to 20 degC)

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Section C. Scheme 03S.127. Material S.106

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Scheme 03S.127. Material S.80

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OIMI 48/2

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Section D. Scheme 03S.114

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Section D. Scheme O3S.136

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Section D. Scheme O3S.137

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Scetion D. Scheme O3S.115

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Section D. Scheme O3S.200

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Section D

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Section D. Scheme O3S.138-141. With drawings

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Section D. Scheme O3S.103-108. With tables and drawings

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Section D. Scheme O3S.127. 1 stand 2nd stage discs

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Section E. Graph 2422 Olympus Mk.3 turbine HP & LP rotor blade loadings: 1st stage 101 blades; 2nd stage 79 blades

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Section E. Blade to PDM.38. Tables. Graphs (2) GR2860 Olympus III HP turbine rotor blade stresses (4454SR/10); Nimonic 80A. PDM38. Graph GR2716 Olympus III turbine HP rotor blade stresses.(4454SR/10)

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Section E. Blade to PDM.38. Page E.7 Blades at maximum cruise. Page E.8 Summary of blade data. Sheet 1: Graph GR2896 Olympus III turbine. Redesigned LP rotor blades stresses (4454SR/10). Sheet 2: Graph GR2896 Olympus III turbine. Redesigned LP rotor blade stresses

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Section E

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Section E

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Section F. 2 page report with drawings and tables

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Section F

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Section G

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Section H: Gear drives and couplings

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Section H: Gear drives and couplings

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Section H: Gear drives and couplings

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Section H: Gear drives and couplings. Conditions (take-off) Fuel flow 1100 gall/hr. Fuel pump 23 hp (Lucas); Pressure rise 1500 lbs/sqr in. tachometer 1 hp. Pressure & scavenge pump 1 hp

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Section H: Gear drives and couplings

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Section H: Gear drives and couplings

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Section H: Gear drives and couplings

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Section J: Balancing. Olympus III: The labyrinth sleeve has been extended from the final dynamic balance (Fig 5 MS.1320 sheet 1). Olympus Mk.1/2C The thrust housing and locating bolts have been excluded from the final dynamic balance (Fig 4 MS.1338, not in our possession) for the same reason as was the labyrinth sleeve in the Olympus III

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Section K: Overspeed testing

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Section K: Overspeed testing. Attached drawing of the above

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Section K: Overspeed testing. Attached drawing of the above

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Section K: Overspeed testing

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Section K: Overspeed testing. Drawing 4096SR/10 attached

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Section K: Overspeed testing. Drawing 4096SR/10 attached

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Section L: Pressure testing

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Section L: Pressure testing

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Section L: Pressure testing

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Section L: Pressure testing

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Section L: Pressure testing

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Section L: Pressure testing

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Section L: Pressure testing

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OLE.201

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Ref: 5216SR/18 attached: F.1021 Schedule of bearing loads for auxiliary drives (Olympus 6) dated 15 September 1954

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Ref: 5216SR/16. We enclose herewith a copy of our scheme MS.2736 which shows the majority of the bearings in the above drives and gives their operating conditions. The remainder of the drives have yet to be finalised and will be forwarded at a later date

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Ref: 5216SR/16. We enclose herwith a copy of scheme MS.2701 which will illustrate the bearings in the above drives and give their operating conditions

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Ref: 5216SR/16. We acknowledge receipt of your letter of the 27th ultimo also a copy of drawing MS.2701 illustrating the bearings you propose in the above drives. We are enclosing herewith a copy of our chart F.1068 outlining all particulars of the bearings in question

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Ref: 5216SR/16. We wish to thank you for your letter and enclose data sheets of the 20th instant M.35 originally shown

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Ref: 5216SR/16. Further to our letter of the 18th instant enclosing a copy of our Scheme MS.2736, will you please note that the bearings at position K and Y have now been changed to your LEJ.10 type ie. the lips will be on the inner track

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Ref: 5216SR/16. Further to your letter of the 18th instant (ref: CE/SWM/KC/9347) we are enclosing a copy of our scheme RS.2736 sheet 2, which illustrates the remainder of the Sundstrand drive and gives the bearings operating conditions. The only portion of these accessory drives not completed is the LP tachometer drives, and this will be forwarded in due course when the design is complete

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Ref: 5216SR/16. Thankyou for your letter and attached data-sheets of the 12th instant, from which it is noted that you consider the bearings satisfactory. With regard to remarks on bearings at position N and R, we shall be retaining them as shown on our Scheme as position N and then be identical to position D in the inner end drive, and position R cannot be reduced on outside diameter due to the adjacent seal

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Ref: 5216SR/16. We thank you for your letter and attached data-sheet 139 together with with sheet F.1068.B, from which it is noted you consider the bearing satisfactory. The information has been passed to the Drawing Office and you should receive copies of our drawings in due course

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Ref: 5216SR/16. CE/SWM/KC/9635. Further to our letters of the 26th september and 22nd November 1957, these drives have been modified to incorporate an American fuel system instead of our standard Lucas system. This has involved some redesign as follows: MS.2701. On your copy of this Scheme, will you please delete the loads and speeds and add this note; 'Loads and speeds are given on MS.2736 sheet 4'. The bearings remain unchanged. MS.2736 sheet 1. This sheet has now been cancelled and the new drives are shown on sheets 3 and 4, which are enclosed with this letter for your attention. MS.2736 Sheet 2. On your copy od this sheet, will you please delete the loads and speeds and add this note: 'Loads and speeds are given on MS.2736 sheet 4'. It is unfortunate that these changes have been required after we have both completed all the technical work on the drives and we regret the inconvenience you have been caused. However, if you could help us once more by clearing these new schemes at an early date, it would of great assistance, as our drawing office is most anxious to clear the details with America

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Our ref: EW-D/KBC/11958. In order that components for the Olympus 22R engine may be tested in advance, we are proposing to use the existing Olympus 200 engine suitably modified. One of the first items to be tested is the above Dowty pump, which it is anticipated will be ready at the end of the month. A copy of our Scheme MS.3195 is enclosed showing the installation of this pump on one of the accessory drive faces of a 200 engine, and you will see that the only new bearing required is an XXJ.40. This bearing has been specified by the designer as FBS.245/40/AD, but we are wondering if you will be satisfied with this in view of the high speed and virtually 'no-load' running condition

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Our Ref: Tech/LEH/KH/508. Your ref: EW-D/KBC/11958. SR521D/19. We thank you for your letter of November 12th with scheme MS.3195, and note you would like to have our comments regarding the bearings selected. The remainder is technical details of the project

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Your Ref: Tech/LEH/KH/508. Our ref: EW-D/KBC/11958. SR521D/19. The bearing XXJ.40 will be used in the test engine only and not in the Olympus 22R itself. Therefore it will not be necessary to have any of these bearings in BACE.173. We do, however, agree that as the tendency is for loads, speeds and temperature of accessory bearings to increase, consideration should be given to specifying future bearings in BACE.173

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Our ref: EW-D/KBC/14096. SR521D/22. Your ref EW-D/KBC/2730. We have examined the above scheme and find that the starter drive to basic schemes 03S.460/1 and 03S.924 is satisfactory for the following conditions when schemes 03S.3004 and 3297 are incorporated. Technical details in rest of memo

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Approved LTC July 19th 1968 - Item 1886

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Ref: 521/SR17. Scheme drawings used 03S2025, 03S1955, 03S1872, 03S1961, 03S1951. 43-page report with stress calculations

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Ref: 5216SR/18. Scheme drawings used: 03S1872, 1951, 1955, 1961, 2025. Previous calculations: 5216SR/14, 5216SR/15, 5216SR/17. Page 1: Tabled information (isued to Technical Records). Pages 2-4: Bearing load sheets for cases other than as 5216SR/15

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Our ref EW-D/HFR/2641. Internal memo from HF Rodgers to Mr K Chamberlaine Stress Office: Strengthened hydraulic pump drive & quill shafts with shear neck failing torques of 4784 lb/in and 3500 lb/in. 4129 lb/in and 2963 lb/in. 03S3292 Issue.1

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Scheme drawings used: 03S3292. Full treatment SM74 and modified SM71 calculations of bearing stress to balance tooth thickness. Page 1: Calculation basis and summary of bend stress. Pages 2-3: Initial calculations (basic proportion) of 8.75 DP bevels. Pages 4-21: SM74 & SM71 tables for 8.67 DP bevels. Page 22: Determination of entries for 8.75 bevels in GDS400. Page 23: Copy of detail DOs; GD5400 sheet 10 of calculations.

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Scheme/drawings used 03S3292. Data referred to GDS400

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Schme/drawings used: Dowty Drawing A.8032.Y.6

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Our ref EW-HFR/2677. Letter from HF Rodgers Assistant Chief Designer, production Engines Bristol Siddeley Engines Limited: We are at present investigating an increased strength engine drive to prevent fatigue failures which have occurred on the shear neck in the engine to the Super Verdal hydraulic pump. In order that we may accurately assess the optimum diameter for the stronger shear neck, we will be pleased if you would send us a detail drawing of the pump shaft for stress evaluation since we understand this shaft to be the weakest part of the pump itself

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Olympus No.600275 was returned from RAF Conings by for suspected metal fragments in the pump. The engine was rejected following an investigation which showed that after a quill failure, part of the quill was missing. This quill is the breaking or shear quill for the hydraulic pump drive

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Scheme/drawings used: 03S460/1&4, 03S3292. Pages 1-13 Diagrams and tables to SM60. Calculations by RSL Pritchard

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Scheme Nos: 03S460/1, 03S460/4, 03S3292. Front sheet plus 10 sheets to SM63. Note! Bearing criteria has been obtained by proportioning existing summary (SR521D/9 & 11) values. This does not take into account the slight differences in mid-face width position of the new 15/31/15 & 13/43 bevels, but this effect is small. Inner end spur gear criteria has also been predetermined from existing summary. Gear & SC's do not allow for MT52 (i.e.as exists Sim for 03S5495 etc.)

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Standard drive with 1880 lbs/in shear neck stiffened drive with 3000 lbs/in shear neck and changes to standard drive to use 3000 neck. Scheme/drawings used: Standard drive 03S495, 03S2555; 03S3292 stiffened drive. Previous calculations 521D/20 of 21/4/6. Pages 1 & 2: basis of calculations. Page 3 Shaft factors & gear SFs. Page 4 Shaft and gear factors. Page 5 Changes required to standard drive to take 3000 neck