LCA news and discussion

KrishG · Post by **KrishG** » 07 Aug 2009 00:34

The ADA website server has collapsed

. Wait! What difference will that make ?? All it used to do was to update the number of flights!

K Mehta · Post by **K Mehta** » 07 Aug 2009 00:41

Question to gurus, will the MRCA testing affect the LCA flight testing? Since ASTE is the agency involved in both these programs and it has been previously said that the lack of pilots for testing is affecting the schedule, would this not contribute too?

shiv · Post by **shiv** » 07 Aug 2009 16:03

KrishG wrote:The ADA website server has collapsed . Wait! What difference will that make ?? All it used to do was to update the number of flights!

Maybe someone tried to update it more frequently than once in 2 months and the server got overloaded?

AmitR · Post by **AmitR** » 07 Aug 2009 16:47

shiv wrote:
KrishG wrote:The ADA website server has collapsed . Wait! What difference will that make ?? All it used to do was to update the number of flights!
Maybe someone tried to update it more frequently than once in 2 months and the server got overloaded?

No wonder the LCA took so long to build.

rohiths · Post by **rohiths** » 07 Aug 2009 18:48

I keep hearing the loud noise of a plane with delta wing vrroming over south bangalore everyday

(To the best of my knowledge it is a LCA)
Yesterday morning the sound was so loud and sharp that it was like a sound of a lion

Proud of our SDREs

enqyoob · Post by **enqyoob** » 07 Aug 2009 22:42

Loud and sharp, hopefully flying supersonic. Over b'lore, kerala? All those flats don't have glass windows, do they?

BTW, I had an interesting chat with someone who is most definitely "in the know" - part of the signoff team on the LCA b4 the pilot strapped in. I merely asked him why in his view India is so sadly behind in jet engine technology. I did not mention GTRE at all.

His view, paraphrased:

Although they had no hands-on experience, GTRE did not bring in the HAL engine types to put the engine together. The engine factory people have decades of experience, but the GTRE types tried to build the components and assemble it themselves...

So it is not my imagination that there is something very very wrong with the way engine R&D is organized in India, and with GTRE.

These things really need a public enquiry, IMO. The sheer lack of seriousness in a project of this much national importance, is truly stunning.

suryag · Post by **suryag** » 07 Aug 2009 23:47

Was there news last month that the trainer was undergoing high speed taxi trials, any updates on that ?

SaiK · Post by **SaiK** » 08 Aug 2009 00:06

it could dont ask, dont tell .. question should be sent to the funding authority or program authority, for such administrative fups.

John Snow · Post by **John Snow** » 08 Aug 2009 00:36

N guru, thats exactly what I was lamenting when I said that India with its current infrastructure can launch hulls for 2 subs (at minimum) per year.

I have since the inception of BRF advocated the inclusion of proven, ITI ( not Indian Telephone Industries, but Industrial Training Institutes) graduates who have worked in PSU/Pvt sector to be part of the teams doing Research ( case point the B.Com graduate developing GIS software for ISRO).

Example. Tell me why we cant produce Indian Field Gun or UAVs with out "foreign know how"?
(rhetorical question)

There is an infatuation with Foreign collobration and the seduction is too tempting to go the desi way.

Please see artillery thread where I said a team of 12 3rd year students can easily design with manufacturing drawings under the leadership of factory foremen, univ prof and army chaps from capt to lt col ranks to give end user and field experience inputs.

Yes it can be done.

SaiK · Post by **SaiK** » 08 Aug 2009 00:47

..then why can't gtre?

John Snow · Post by **John Snow** » 08 Aug 2009 02:50

Long ago I had asked why not ceramic blades again obly Vina garu responded with the thermal gradient that a blade has to stand and the tensile and compressive forces that come into play.

http://www.unipass.com/predictionprobe/ ... ngines.htm

however it is being done ( ceramic, sintered ceramic blades for turbines,) one of my friends (PhD Physics from Mumbai IIT ws working on similar project but for ceramic coated with Pd for catalytic converters in Ford dearbon MI).

Why not a conventional blade be coated with ceramics to enhance temp withstanding just ike RCC (With steel reinforements). Either with spray ceramic material and then bake, sinter on to the conventional blade.

They key to all this is heat treatment furnaces

more here
http://www.freepatentsonline.com/6602548.html

A nickel base single crystal compliant layer on a ceramic blade has the capability to sustain high stresses and high operating temperature. Layers of nickel and platinum bonded on a single crystal superalloy over a sputtered gold-chromium layer support the high stress levels at elevated temperature without extrusion of the soft platinum or nickel layer and without destruction of an NiO compliant surface. The compliant layers have survived stress and temperature conditions without failure to the ceramic blade and the system can be stressed/heated and unloaded/cooled repeatedly without damage to the ceramic blades. A single crystal nickel base superalloy (i.e., SC180) has high strength properties at elevated temperature. Thin layers of chromium followed by gold are e-beam evaporated on one side of a polished surface of the alloy. Pure nickel is electroplated over this e-beam gold-chromium layer. Platinum is either electroplated or plated electrolessly over the nickel layer. The structure is annealed in vacuum or inert atmosphere to allow the diffusion of gold-chromium alloy into the superalloy and permit the nickel layer and diffusion of nickel into platinum to form a multilayer structure which is metallurgically bonded. The sheet is oxidized in air to allow diffusion of the nickel layer through the platinum to come to the surface and oxidize forming nickel oxide. This nickel oxide layer acts as the load distribution layer which does not extrude and the structural integrity of the compliant layer is maintained by the high-strength single crystal superalloy.

http://www.grc.nasa.gov/WWW/RT/RT1999/5000/5930min.html

Aditya_M · Post by **Aditya_M** » 08 Aug 2009 03:11

there is something very very wrong with the way engine R&D is organized in India, and with GTRE

since I have worked with private sector R&D in India, let me say that the wrongness is not limited to the public sector. Lack of interaction between design and production, management and design, management and customer, customer and end-user - we found out that last bit the very hard way! - I've seen it all. And I'm not even a grizzly old vet.

From what I have read, very few companies in India get it right. L&T *may* be one of them, Bajaj has done extremely well in the last few years by going from "Kawasaki Bajaj" to just "Bajaj".

pravula · Post by **pravula** » 08 Aug 2009 03:28

AmitR wrote:
shiv wrote:
Maybe someone tried to update it more frequently than once in 2 months and the server got overloaded?
No wonder the LCA took so long to build.

It up now.

LCA-Tejas has completed 1151 Test Flights successfully. (01-Aug-09).

* LCA has completed 1151 Test Flights successfully
(TD1-233, TD2-305,PV1-209,PV2-125,PV3-147,LSP1-52,LSP2-80).
* 209th flight of Tejas PV1 occurred on 30th July 09.

John Snow · Post by **John Snow** » 08 Aug 2009 04:00

We can do with out foreign collabration. Its like I doing the home work for my kid even if he is smart, he will always have the nagging doubt can I do it alone.

Yes we must

Bajaj is bogus, Only M&M Tata Godrej come close to doing it or reverse engineering or building on the expertise gained.

NRao · Post by **NRao** » 08 Aug 2009 04:14

John Snow wrote:We can do with out foreign collabration. Its like I doing the home work for my kid even if he is smart, he will always have the nagging doubt can I do it alone.

Yes we must

Bajaj is bogus, Only M&M Tata Godrej come close to doing it or reverse engineering or building on the expertise gained.

Ability is a given. It is management, structure and the will to get it done that is missing. Perhaps some combination of those items.

enqyoob · Post by **enqyoob** » 08 Aug 2009 06:05

Please see artillery thread where I said a team of 12 3rd year students can easily design with manufacturing drawings under the leadership of factory foremen, univ prof and army chaps from capt to lt col ranks to give end user and field experience inputs.

Yes it can be done

.

Snow garu, great and heartfelt post, exellent recommendations.

Long time ago in the days when the Soviet Cultural Society (whatever it was called) used to distribute free magazines, I remember seeing on the cover of one, a picture of something that looked like a diesel-electric locomotive, with a group of smartly dressed people around it.

Railway Locomotive Built By Young Pioneers

"YP" was Soviet answer to Boy Scouts and Hitler Jugend, I suppose, but these were supposed to be middle school / high school kids who had put together a locomotive. I have no idea if it went downhill any better than the "Gorshkov" but anyway, this has always stayed in my mind.

The GTRE Kaveri Engine seems to be something along these lines:

Combat Aircraft Engine Designed and Built By Young Pioneers

No real plans beyond PR value. BTW, the person I talked to, once I had triggered him, had several choice things to say about the disconnect between PR claims and reality on the Kaveri project.

If I was in Spin Mode I could use all the evidence above to prove conclusively that the Kaveri Project is a cover for the Ultra Secret XXXXXXXXX ( I cannot say what it is, obviously, paanwala and chaiwallah both emphasized this point) Project. Just a guess:

"Jatayu" Combined Cycle Sea2Space Multipurpose Gs Turbine Engine. Uses Air Liquefaction, Variable Pressure Ratio Integrated Radial/Axial Compressor, GOBAR gas converter, Plasma Jet Afterburner, Pependicular Fuel Injection Ferri Thermal Compression turboramjet stage, and several other Classified technologies.

This is how the much-maligned "AKILA" RPV (no relation to Akula Submarinov) was used as the public cover to siphon funds to the super-secret B-2 Stealth Bomber.

I mean, WHY ELSE would they make such a complete mess for 3 decades?

ramana · Post by **ramana** » 08 Aug 2009 06:41

You mean
Akula Subramanyam.

John Snow · Post by **John Snow** » 08 Aug 2009 07:34

Think of it I now recall that BARC has a Solid State Physics, Crystalography division with a whole bunch of Phds and wanna be Phds.

We had people who were in TIFR, BARC Trombay from 1960s and worked on Solid State physics division.

How come these groups can't come up something for single crystal blades?

see this from 1998
http://www.ias.ac.in/currsci/apr25/articles7.htm

DAE–Solid State Physics Symposium 1998 – A report

The annual Solid State Physics Symposium (SSPS), sponsored by Board of Research for Nuclear Science, Department of Atomic Energy (DAE) was held this year in Kurukshetra University, Kurukshetra.

There were 304 papers scheduled to be presented. These included 22 invited talks. A seminar on nano-phase materials and two tutorial sessions, on Experimental Techniques, Data Processing and Scientific Visualization were scheduled. Out of the 257 contributed papers, 18 were chosen for oral presentation and the rest were posters. Fifteen theses were selected for presentation (oral + poster) out of which 10 were presented.

The scientific session started with a talk by Priya Vashishta (Louisiana State University, USA). He highlighted the current status of the molecular dynamics simulation (MDS) on parallel computers using multi-million atoms. Important device materials like silicon nitride and silicon carbide have been studied combining first-principle calculations and multi-million atom MDS to determine the stress distribution in a 54 nm pixel on a 0.1 m m silicon substrate.

The talk by Anil Kumar (Indian Institute of Science, Bangalore) dealt with 3-D lsing critical behaviour in aqueous solutions. He mentioned that the critical behaviour in complex fluids does not conform to the 3-D lsing type; instead a mean-field approach provides a better description. It was concluded that the breakdown of the anticipated 3-D lsing behaviour is mainly due to the structuring induced by the electrolytes.

R. Mukhopadhyay (Condensed Matter Physics Division, (CMPD), BARC) talked on classical and quantum dynamics of the methyl side-groups attached to the main polymer chain as studied by the neutron scattering technique. Neutron scattering being very sensitive to the protons is the best suited probe to study the dynamics in glassy polymers. He showed the first experimental evidence of quantum rotational tunneling in polymers. He also demonstrated how the methyl group dynamics can be consistently described by a distribution of barriers for random hopping in the classical regime and by a distribution of the tunneling frequencies in the quantum regime, both having the same physical origin: disorder inherent to the amorphous state of the polymer.

K. P. N. Murthy (Material Science Division (MSD), IGCAR) spoke on aging scaling, often found in slowly relaxing nonequilibrium systems. He attempted to describe the process using directed random walk models with site dependent transition probabilities. The aging was confirmed by calculating exactly the autocorrelation and its scaling with the ratio of the times. The talk by A. V. Rao (Physics Department, Shivaji University, Kolhapur) was on the technology of a novel material highly porous (99% air and 1% solid) silica aerogel as prepared by sol–gel technique. It has very low bulk density of <20 mg/cm3 but has visible transparency of 90% which makes this an important material that finds application in low mass liner for the generation of high intense soft X-rays, Cerenkov radiation detectors and as a host for many confinement requirements.

The tutorial session on Scientific Visualization dealt with two presentations: H. K. Kaura on Virtual Reality and P. S. Dhekne on Scientific Visualization, both from Computer Division, BARC.

In the evening talk, R. Chidambaram highlighted the importance of solid state physics for conducting nuclear tests of desired yields.

Vince McKoy (Caltech, US) discussed the progress in exploiting large parallel computers to generate electron collision cross-sections for gases of interest to the semiconductor industry. He gave an overview of the methods employed in these calculations and its computational demands and discussed the strategies used to parallelize the computer intensive steps. Parongama Sen (Surendranath College, Calcutta) reviewed the present status on coexisting spanning clusters in percolation. She discussed the different properties of such clusters and highlighted the existing problems in the spanning clusters in ordinary and directed percolations.

The seminar on nanophase materials was a topic of current interest for its numerous applications in the field of nonlinear optical devices, metallurgy and catalysis. A. K. Arora (MSD, IGCAR) who coordinated the seminar discussed various methods of synthesis and characterization of nanoparticles. S. Ramasamy (Physics Department,Madras University) described the synthesis of magnetic nanostructured materials using ultra high vacuum chamber which gives high purity grain boundary structure and results in various magnetic properties. S. Mahamuni (Physics Department, Pune University) discussed the optical properties of the quantum dots using solid state theories based on the delocalized electrons and holes within the confined volume. Her talk brought out the fact that enhanced nonlinear optical properties are the consequence of state filling effects as well as bleaching of the exciton absorption by the presence of surface trapped electron hole pair. She also discussed size and shape-dependent properties with reference to II-VI semiconductor dots.

S. C. Gadkari (Technical Physics & Prototypic Engineering Division, (TPPED) BARC) reviewed the solid state sensors used for gas sensing applications, in general and particularly emphasized the devices based on metal-oxide semiconductor thin films. He indicated future trends in the field of development of gas sensor array to realize the ‘electronic nose’ using neural network.

N. D. Sharma (Kurukshetra University) gave an overview on surface modifications on ion implanted materials with specific examples of 304 stainless steel, nimonic-90 alloy and aluminium implanted with 130 keV nitrogen, boron and argon ions at different doses. The possible increase of near-surface hardness was attributed to the formation of nitride and boride precipitates and dislocation pinning.

The talk by Prasenjit Sen (School of Physical Sciences, Jawaharlal Nehru University, New Delhi) on dissipative structure formation in heavy ion irradiation, provided evidence of realizing such phenomena in metals. He showed that these are formed as rearrangements in microstructures filled with stationary imperfections like dislocations and grain boundaries. He identified processes leading to such rearrangements.

Matti Lindroos (Tampere University, Finland) reviewed the electronic structure and fermiology of complex materials. He presented the results of first principles computations of the photo-intensity in high Tc material Bi2212. Substantial matrix element effects and remarkable anisotropy of the CuO2 plane band intensities were observed in the experimental spectra.

A presentation by R. P. Dahiya (IIT, Delhi) on surface nitriding of steel components in expanding plasma highlighted the industrial applications of such studies.

R. Ramakumar (Low Temperature Physics Division, Saha Institute of Nuclear Physics, Calcutta) discussed the nature of quasi one- and two-dimensional organic superconductors. He described the model developed to understand superconductivity in TTF[M(dmit)2]2 (where M is Ni, Pd or Pt). These materials have an electronic structure in which both highest and lowest unoccupied molecular orbitals derived bands cross the Fermi level and helped superconductivity.

P. S. Goyal (Inter University Consortium for DAE Facilities, Mumbai) gave a talk on anomalous thermal properties and tunnelling states in solids with a specific example of mixed salt system of ammonium and alkali iodides which show non-Debye specific heat behaviour. The results of neutron scattering and specific heat measurements on these mixed salts were presented.

S. Mazumder (CMPD, BARC) highlighted that a scaling law different from those observed in the case of binary alloys, is valid in describing the temperature-dependent phase separation behaviour of a multi-component alloy like maraging steel.

P. Ch. Sahu (MSD, IGCAR) described the laser heated diamond
anvil cell technique and discussed its importance in material synthesis with illustrations from his own work in
Japan.

The first tutorial session on Experimental Techniques: Characterization of materials was coordinated by S. C. Gupta (CMPD, BARC). He briefly described the various techniques like STM, AFM, and CCD-based X-ray imaging. A. Sinha (CMPD, BARC) discussed in detail his set-up of CCD-based X-ray imaging for both normal and high pressure applications. N. Venkatramani (Advanced Centre for Research in Electronics, IIT, Mumbai) showed with various illustrations the potentials of Scanning Probe Microscope as a tool to characterize materials. He mentioned that atomic scale resolution is possible in the AFM and a resolution better than SEM is possible in SPM. He stressed the fact that SPM imaging has reached a high degree of reliability and is an important technological tool, specially for storage media, profilometry and as a tribological device. G. Raghavan (MSD, IGCAR) discussed the characterization of interfacial evolution in film multilayers. He pointed out that the evolution of different phases in heat treated film multilayers involves issues related to interdiffusion and microstructure.

The tutorial session on Data Processing and Scientific Visualization (coordinated by B. K. Godwal and R. Mukhopadhyay, from CMPD, BARC) was an attempt to bring about an interaction between the physicists and the computer scientists. As the basic purpose of scientific computing is to gain an insight into the problem that is being probed, it was felt that a session based on such interaction will be useful. B. S. Jagadeesh and S. K. Bose (Computer Division, BARC) enlightened the participants through their detailed lecture-cum-demonstrations on data processing and scientific visualization.

J. Jayapandian (MSD, IGCAR) highlighted the novel design concepts in PC-based integrated data acquisition and control systems. He described various novel interfacing design techniques for automation in industries and laboratories. He stressed that such indigenous approach will save a good amount of foreign exchange.

vina · Post by **vina** » 08 Aug 2009 07:41

I think it is not just the physics part of getting the material. That is the easier part. The harder part is the engg problem. Will you be able to grow the crystals to a bigh enough size ?. There is no single text book that will teach you how to do it. That you will learn only by trial and error and actually doing it before you get it exactly right. Next comes the fabirication and machining part. Will you be able to machine it into the intricate shapes. Will you be able to drill the cooling air channels in it with super precision.

By no way it will be just a "lab" effort. It is much bigger than that

enqyoob · Post by **enqyoob** » 08 Aug 2009 07:54

Single crystal turbine blades

http://www.azom.com/details.asp?ArticleID=90

Since the 1950’s, the evolution from wrought to conventionally cast to directionally solidified to single crystal turbine blades has yielded a 250°C increase in allowable metal temperatures, and cooling developments have nearly doubled this in terms of turbine entry gas temperature. An important recent contribution has come from the alignment of the alloy grain in the single crystal blade, which has allowed the elastic properties of the material to be controlled more closely. These properties in turn control the natural vibration frequencies of the blade.

If metallurgical development can be exploited by reducing the cooling air quantity this is a potentially important performance enhancer, as for example, the Rolls-Royce Trent 800 engine uses 5% of compressor air to cool its row of high pressure turbine blades. The single crystal alloy, RR3000, is able to run about 35°C hotter than its predecessor. This may seem a small increase, but it has allowed the Trent intermediate pressure turbine blade to remain uncooled.

Continuing Developments
It is estimated that over the next twenty years a 200°C increase in turbine entry gas temperature will be required to meet the airlines' demand for improved performance. Some of this increase will be made possible by the further adoption of thermal barrier coatings. These coatings are produced from ceramic pre-cursors and have the potential to contribute about 100°C through the protection they provide.

Thermal Barrier Coatings
Thermal barrier coatings have been used for some years on static parts, initially using magnesium zirconate but more recently yttria-stabilised zirconia. On rotating parts, the possibility of ceramic spalling is particularly dangerous, and strain‑tolerant coatings are employed with an effective bond coat system to ensure mechanical reliability.

Ceramic Matrix Composites
Further increases in temperature are likely to require the development of ceramic matrix composites. A number of simply shaped static components for military and civil applications are in the engine development phase and guide vanes have been manufactured to demonstrate process capability, such techniques involve advanced textile handling and chemical vapour infiltration.

However, it is the composite. ceramic rotor blade that provides the ultimate challenge. It will eventually appear because the rewards are so high, but it will take much longer to bring it to a satisfactory standard than was anticipated in the 1980’s. Research work has concentrated for some years on fibre reinforced ceramics for this application, as opposed to monolithic materials which possess adequate strength at high temperatures but the handicap of poor impact resistance.

Today's commercially available ceramic composites employ silicon carbide fibres in a ceramic matrix such as silicon carbide or alumina. These materials are capable of uncooled operation at temperatures up to 1200°C, barely beyond the capability of the current best coated nickel alloy systems. Uncooled turbine applications will require an all oxide ceramic material system, to ensure the long term stability at the very highest temperatures in an oxidising atmosphere. An early example of such a system is alumina fibres in an alumina matrix. To realise the ultimate load carrying capabilities at high temperatures, single crystal oxide fibres may be used. Operating temperatures of 1400°C are thought possible with these systems.

Primary author: Stewart Miller
Abstracted from Materials World, vol. 4, 1996, “Advanced materials mean advanced engines”

For more information on Materials World please visit The Institute of Materials

Date Added: Feb 21, 2001

John Snow · Post by **John Snow** » 08 Aug 2009 07:55

N guru. I talked about Indian Field Gun, for which we are scouting collobration.

I did not talk about Jet Engine, just a gun to shoot projectiles, India made canons in the 17th century (cast Iron, made muskets) no?

We have been making field Gun, what is so new to make one? unless again the Field Gun requires Single Crystal tech for the barrel?

some answers some where in the Google land.

Barrels have about 6 different stresses working on them when a gun fires. Barrels are made from forged steel and a technique called autofrettage is used to strengthen them

Precisely: engineering problem,
actually technology problem, thats why we need ITI, foremen, and under grad guys, in colleges we are told how things are done, It is for us to develop how things can also be done ( from different perspective).

The amount of equipment and R&D facilities gathering cobwebs is to be seen to be beleived

steel hardening is an odd beast. some countries use different methods - arguably (on average) the germans have made the best. in a twist of irony, depending on method used, the harder a steel becomes, the more prone it is to shattering or becoming brittle. the tech to strengthen a barrel is different from strengthening a panel etc... some barrels are one piece, some are sleeved, some are jacketed etc.... there are heaps of issues to contend with such as heat build up, heat dissipitation etc.. - also, it's not really a design parameter to build in resistance to being struck by an incoming round etc... as an example, (generally) because the russians metallurgy was at a different level than western technology, they favoured larger calibre rounds to compensate for an engineering difference. thats an oversimplification - but it gives you an idea of how build differences could impact on eventual design solutions. it's a really quite complex topic, and probably some of the arty or tanker guys in here could give you a concise overview

I am conscious that the methods used in the production of the barrels strength it from the inside, and i would like to know what are the effects of them to pressure from the outside of the barrels. I know this is a complex subject and thus i do not need a complicated diffrectial anlysis of solid state metals, but more of a thumb estimate. Inm addition does anyone know the type of steel used in the production?

oh I dunno...The steel compositions of the various Russian and Chinese artillery and 100+mm guns should fairly well known by now... GERALD BULL KNEW, CL

heat treatment its all heat treatment

enqyoob · Post by **enqyoob** » 08 Aug 2009 08:04

"Crown Jewels"

The gas turbine manufacturer Rolls-Royce, for instance, outsources and offshores about 75 percent of its components to its global supply chain. But what of that remaining quarter?

Friedman quotes Sir John Rose, Roll-Royce's CEO, on the components the company still makes: "The 25 percent that we make are differentiating elements. These are the hot end of the engine, the turbines, the compressors and fans and the alloys, and the aerodynamics of how they are made. A turbine blade is grown from a single crystal in a vacuum furnace from a proprietary alloy, with a very complex cooling system. This very high-value-added manufacturing is one of our core competencies."

As Sir John Rose points out, single-crystal turbine blades are valuable to the companies that make them, as a high-technology—and profitable—product. Since they were invented and introduced, their unmatched resistance to high-temperature creep and fatigue have helped to advance the performance and durability of modern gas turbines. Were it not for the cost, they would replace thousands of conventionally manufactured turbine blades in higher-temperature applications.

A number of years ago, I worked in the engineering organization in which these "gas turbine crown jewels" were invented. It is surprising, given their importance, that even today much of the story behind their development is not well-known.

.... Generations of designers, engineers, and researchers have worked to increase gas turbine thermal efficiencies from an inaugural value of 18 percent to an unmatched 60 percent, in modern combined-cycle operation.

Vanes and Blades

...ASME's International Gas Turbine Institute over the last 50 years has had 14,000 reviewed technical papers presented at its gas turbine conferences worldwide.

..(see picture at the page)..
The beige single crystal blades on this GE 9H turbine—the world's largest—are approximately 18 inches long and weigh more than 30 pounds apiece.
Gas turbine thermal efficiency increases with greater temperature of the gas flow exiting the combustor and entering the work-producing component—the turbine. Turbine inlet temperatures in the gas path of modern high-performance jet engines can exceed 3,000°F, while nonaviation gas turbines operate at 2,700°F or lower. In high-temperature regions of the turbine, special high-melting-point nickel-base superalloy blades and vanes are used, which retain strength and resist hot corrosion at extreme temperatures. These superalloys, when conventionally vacuum cast, soften and melt at temperatures between 2,200 and 2,500°F. That means blades and vanes closest to the combustor may be operating in gas path temperatures far exceeding their melting point and must be cooled to acceptable service temperatures (typically eight- to nine-tenths of the melting temperature) to maintain integrity.

Thus, turbine airfoils subjected to the hottest gas flows take the form of elaborate superalloy investment castings to accommodate the intricate internal passages and surface hole patterns necessary to channel and direct cooling air (bled from the compressor) within and over exterior surfaces of the superalloy airfoil structure. To eliminate the deleterious effects of impurities, investment casting is carried out in vacuum chambers. After casting, the working surfaces of high-temperature cooled turbine airfoils are coated with ceramic thermal barrier coatings to increase life and act as a thermal insulator (allowing inlet temperatures 100 to 300 degrees higher).

Grain Boundary Phenomena

Conventionally cast turbine airfoils are polycrystalline, consisting of a three-dimensional mosaic of small metallic equiaxed crystals, or "grains," formed during solidification in the casting mold. Each equiaxed grain has a different orientation of its crystal lattice from its neighbors'. The resulting crystal lattice misalignments form interfaces called grain boundaries.

Untoward events happen at grain boundaries, such as intergranular cavitation, void formation, increased chemical activity, and slippage under stress loading. These conditions can lead to creep, shorten cyclic strain life, and decrease overall ductility. Corrosion and creaks also start at grain boundaries. In short, physical activities initiated at superalloy grain boundaries greatly shorten turbine vane and blade life, and lead to lowered turbine temperatures with a concurrent decrease in engine performance.

One can try to gain sufficient understanding of grain boundary phenomena so as to control them. But in the early 1960s, researchers at jet engine manufacturer Pratt & Whitney Aircraft (now Pratt & Whitney, owned by United Technologies Corp.) set out to deal with the problem by eliminating grain boundaries from turbine airfoils altogether, by inventing techniques to cast single-crystal turbine blades and vanes.

"Single-crystal airfoils have proved to have more relative life in terms of creep strength, thermal fatigue resistance, and corrosion resistance."

Single-crystal turbine airfoil development took place in the company's Advanced Materials Research and Development Laboratory, under the direction of Bud Shank. The first important development was the directionally solidified columnar-grained turbine blade, invented by Frank VerSnyder and patented in 1966.

(Patents expire in 17 years... so a starting point is to read that patent, which describes at least in general legalese gobbledygook how to go about it).

Direction solidification, carried out in a vacuum chamber, is accomplished by pouring molten superalloy metal into a vertically mounted, ceramic mold heated to metal melt temperatures, and filling the turbine airfoil mold cavity from root to tip. The bottom of the mold is formed by a water-cooled copper chill plate having a knurled surface exposed to the molten metal. At the knurled chill plate surface, crystals form from the liquid superalloy and the solid interface advances, from root to tip.

The mold is surrounded by a temperature-controlled enclosure, which maintains a temperature distribution on the lateral surfaces of the mold so that the latent heat of solidification is removed by one-dimensional transient heat conduction through the solidified superalloy to the chill plate. As the solidification front advances from root to tip, the mold is slowly lowered out of the temperature-controlled enclosure.

The final result is a turbine airfoil composed of columnar crystals or grains running in a spanwise direction. For the case of a rotating turbine blade, where spanwise centrifugal forces set up along the blade are on the order of 20,000 g, the columnar grains are now aligned along the major stress axis. Their alignment strengthens the blade and effectively eliminates destructive intergranular crack initiation in directions normal to blade span. In gas turbine operation, directionally solidified turbine blades have much improved ductility and thermal fatigue life. They also provide a greater tolerance to localized strains (such as at blade roots), and have allowed small increases in turbine temperature (and, hence, performance).

Building upon direction solidification, Pratt & Whitney reached the goal of liquidating turbine airfoil grain boundaries in the late 1960s.

Barry Piearcey patented Pratt & Whitney's first single-crystal turbine blade, which was followed by a patent on an improvement by Bernard Kear. Maury Gell patented single-crystal blade alloy composition improvements that raised the incipient melting temperature by 150 to 200 degrees, which provided for a direct potential increase in engine performance.

One Airfoil, One Crystal

The making of a single-crystal turbine airfoil starts the same as a direction solidification airfoil, with carefully controlled mold temperature distributions to ensure transient heat transfer in one dimension only, to a water-cooled chill plate. Columnar crystals form at the knurled chill plate surface in a mold chamber called the "starter." The upper surface of the starter narrows to the opening of a vertically mounted helical channel called the "pigtail," which ends at the blade root. The pigtail admits only a few columnar crystals from the starter. As solidification proceeds up the helix, crystal elimination takes place so that only one crystal emerges from the pigtail into the blade root, to start the single crystal structure of the airfoil itself.

Again, one-dimensional transient heat conduction must be maintained as the mold is withdrawn from the temperature-controlled enclosure. Any heat conducted to mold lateral surfaces can cause localized crystallization, which disrupts the single-crystal structure, with secondary grains.

In the 1970s, Pratt & Whitney developed techniques to manufacture single-crystal turbine airfoils, and to overcome casting defects (such as secondary grains, recrystallized regions, and freckle chains). This early pioneering work has been taken over by other manufacturers and improved upon over the past 30 years. Yields greater than 95 percent are now commonly achieved in the casting of single-crystal turbine airfoils for aviation gas turbines, which minimizes the higher cost of SX casting compared to equiaxed casting.

Early on, Pratt & Whitney investigated single-crystal turbine airfoil use in various jet engines. (One of the first was the J58, which powered the Lockheed SR-71 Blackbird. The very first actual engine use was in Pratt & Whitney's JT9D-7R4 which received jet engine flight certification in 1982. This first single-crystal bladed engine powers the Boeing 767 and the Airbus A310.)

Strength and Resistance

In jet engine use, single-crystal turbine airfoils have proven to have as much as nine times more relative life in terms of creep strength and thermal fatigue resistance and over three times more relative life for corrosion resistance, when compared to equiaxed crystal counterparts. Modern high turbine inlet temperature jet engines with long life (that is, 25,000 hours of operation between overhauls) would not be possible without the use of single-crystal turbine airfoils. By eliminating grain boundaries, single-crystal airfoils have longer thermal and fatigue life, are more corrosion resistant, can be cast with thinner walls—meaning less material and less weight—and have a higher melting point temperature. These improvements all contribute to higher efficiencies.

The newest chapter of the story is their recent introduction in new, large land-based gas turbines, where they hold promise of minimum life cycle cost for increased turbine temperatures. Gas turbines used to produce electric power in the 200 to 400 MW range have turbine airfoils that can be 10 times larger than jet engine turbine airfoils. These large SX castings have had production problems in the industry, causing casting yields to go down, driving costs up.

As an example, one 1999 study done for the U.S. Department of Energy found that for a nominal $6,000, 13.6 kg single-crystal blade, a 90 percent yield would raise the cost to $7,000, while a 20 percent yield would shoot unit costs up to $30,000 each. Much work has been going on in the casting industry to increase yields for these large turbine blades and vanes, using liquid aluminum or tin cooling, or inert gas jet cooling, to increase the efficiency of the critical one-dimensional transient heat transfer process that controls single crystal solidification.

....

Lee S. Langston, professor emeritus of mechanical engineering at the University of Connecticut, is the editor of ASME's Journal of
Engineering for Gas Turbines and Power.

There u go, now that's the Google search is done, all that remains is to actually produce it.

John Snow · Post by **John Snow** » 08 Aug 2009 08:15

I am assuming the SCB is made of ferrous alloy (or magnetic alloy), so can not magnetic field application make the crystal grow the desired way? Is this possible?

Does the current Mig 21 engine incorporate Single Crystal Blades?

If MTBF is the only issue why cant Kaveri go ahead run without SCB?

How many a/c in IAF actually use SCB in their engines?

I am guessing Mirage 2000, Mig-29, SU-30 MKI, Jaguar? MIg-27? Mig-21 last three I doubt but I could be wrong

enqyoob · Post by **enqyoob** » 08 Aug 2009 08:19

c Technical paper.They probably have put in a few fundamental errors, just to make sure ppl do the phirsht prinjipal derivashun. When I was a PIGS in my first Quarter, I was given the job of coating 25-micrometer thermocouples with a "Eutectic" coating of Beryllium Oxide (which is deadly poison) and Ytriium Oxide (which is used to treat insanity). Very simple, procedure was given in a technical paper from Dera Kangaroo Khan.

A small step was omitted : exactly how to get BeO and Y2O3 (or whatever its formula was) into the required form, which was not what came in the bottle. Took me a year+ to figure it out. The poison obviously did not kill me, no comments on the other aspect. Point is, it had to be figured out to do all the rest of the research on flames that was to follow, though it took months of wading through chemistry handbooks (I HATE chemistry) and very very careful experimentation with good stuff like conc. H2SO4, HCL, etc. etc. The only time my dear Advisor really was about to go through the roof was when I presented my solution for increasing the flame temperature from the 1500 C that the Bunsen burner was achieving, to the 1800C + range to see if the eutectic coating was actually doing its job (when it worked right, it was transparent, and extremely thin, barely visible under the sort of microscope I had - like the Emperor's Clothes. The only way to figure out that it was working was by seeing that the thermocouple was actually surviving and reading temperatures close to the melting point of Platinum). I suggested connecting the 2500PSIA hydrogen cylinder into the bunsen lab's gas line.

U R NOT GOING TO PUT HYDROGEN INTO THE BUILDING'S PLUMBING SYSTEM!

None of these things are easy even when you have the technical paper - but this is all anyone is going to give u. If I had a coated thermocouple to use, like desi injineers have from the hajaar injins that are being taken apart every day, that use single-crystal blades, I would have succeeded much sooner, not wasting all that time trying to prove I had got to the right answer. But it cannot be done by ppl who can't even figure out the benefits of being able to do reverse engineering.

John Snow · Post by **John Snow** » 08 Aug 2009 08:26

N guru you were ahead of yourself, in those days they did not think much of Green Energy and the contribution of Hydrogen to green economy

But I am glad you did not but most of neigborhood in Hyderbad used H + O2 do welding for window grills, some times rig up a device to generate H by creating a small generater with water and CaCo3...

Glad we both live to see real green economy no

bala · Post by **bala** » 08 Aug 2009 08:27

John Snow wrote: Yes it can be done.

I am going with this single quote of Snow/Spinster Garu. I am sure everyone in GTRE and DMRL has read every book on single crystal blades. The issue is trying to create one. From now, all Metallurgy Depts in Engineering Institutes can be given a single mission: create a Single Crystal Blade which can meet or beat the specs of a RR, Pratt & Whitney, GE, Saturn. The reward is Rs 100 crores or more.

negi · Post by **negi** » 08 Aug 2009 08:33

Are you sure about Hydrogen Snowji ?

I believe it must be standard issue Oxy Acetylene torch ; H2 is expensive to extract and requires a more expensive means to compress in a equally rugged cylinder . And how does one produce H2 from CaCO3 and H2O ?

vasu_ray · Post by **vasu_ray** » 08 Aug 2009 08:39

bala wrote:create a Single Crystal Blade which can meet or beat the specs of a RR, Pratt & Whitney, GE, Saturn. The reward is Rs 100 crores or more.

DARPA style challenge is great, however does the right infrastructure exist? and accessible to the teams?

John Snow · Post by **John Snow** » 08 Aug 2009 08:48

Naveen Negi garu>> I was trying to recall from my memory of Bal & Tuli Inorganic chemistry I read in 1970/71, so I started google search. And found this

Pelletizing or granulating process - Patent 5468446compacting the material coated with calcium carbonate to produce said pellets ... over the reaction of the calcium hydride in water to produce hydrogen
The pelletisation or granulation of a material or mixture of materials the or at least one of which is reactive in a liquid to produce a gas is improved by treating the reactive material prior to final compaction to form a coating thereon of a substance which is less soluble in the liquid than the reactive material. The preferred reactive material is calcium hydride and the preferred coating is calcium carbonate with or without calcium hydroxide.

I stand corrected not calcium Carbonate but Calcium Hydride. It could be ALminum Hydride also.

the balanced equation to this is
CaH2+2H2O----Ca(OH)2+2H2

********
I did this in 7 th grade for fun.
Take a empty homeopathy glass bottle its about 6 inches with a cork lid (like wine bottle)

Add Pan chuna (CaCo3) add Aluminim foil from cigerette pack, in those days packs used to have aluminium foil with a wax paper like tissue paper attached to it. add them as small balls
Then add water and shake the bottle and put it a safe distance with in couple of minutes you see bublles in the glass container and the cork would shoot almost 10 feet. If the cork was too tight the glas jar would break hence the safe distance. I was caught and my ears were really really boxed, then I branched out into eletronics by 8th grade to make crystall radio and then on to class B pus pull amplifiers (with Hitachi and Toshiba 2sb77 Mullard equivalent of OC71 and OC72... thats a story for different time)

***
There is nothing that is not in India

enqyoob · Post by **enqyoob** » 08 Aug 2009 08:50

Infrastructure for crystallization

Laser-heated pedestal growth

The main advantages of this technique are the high pulling rates (60 times greater than the conventional Czochralski technique) and the possibility of growing materials with very high melting points.[2][3][4] In addition, LHPG is a crucible-free technique, which allows single crystals to be grown with high purity and low stress.

Wikipedia single crystal

I gather that this is the technology needed, to make powerful pulsed lasers? Had no idea..
I cubic Zirconia ("AmirKhani Diamond") what comes when the single crystal turbine blade doesn't work?

Anyway, point is there is so much info available, and so many Chemistry MSc grads and Materials Engineers in yindoostan, that this should not be such a big deal. "Infrastructure" just has to be built by determined people.

John Snow · Post by **John Snow** » 08 Aug 2009 08:55

That means the alloy is zirconium alloy.
NFC has great facility to do Zirconium stuff. there is huge shop with furnace etc to make fuel rods, there is zirconium tubes mill. there is ball bearing plant with everything from MAN & Seimens Germany. Midhani should also be able to picth in for raw materisl.

***
from uncle wiki

As its name would imply, cubic zirconia is crystallographically isometric and, as diamond is also isometric, this is an important attribute of a would-be diamond simulant. During synthesis zirconium oxide would otherwise form monoclinic crystals, its stable form under normal atmospheric conditions. The stabilizer is required for cubic crystal formation; it may be typically either yttrium or calcium oxide, the amount and stabilizer used depending on the many recipes of individual manufacturers. Therefore the physical and optical properties of synthesized CZ vary, all values being ranges.

It is a dense substance, with a specific gravity between 5.6–6.0 — at least 1.6 times as dense as diamond. Cubic zirconia is relatively hard, at about 8 on the Mohs scale— much harder than most natural gems.[1] Its refractive index is high at 2.15–2.18 (B-G interval, compared to 2.42 for diamonds) and its luster is subadamantine. Its dispersion is very high at 0.058–0.066, exceeding that of diamond (0.044). Cubic zirconia has no cleavage and exhibits a conchoidal fracture. Because of its high hardness, it is generally considered brittle.

Under shortwave UV cubic zirconia typically luminesces a yellow, greenish yellow or "beige". Under longwave UV the effect is greatly diminished, with a whitish glow sometimes being seen. Colored stones may show a strong, complex rare earth absorption spectrum.

No wonder women dont like Cubic Zircona that much unless blessed naturally

**************
ceramic abrasive tip to SCB.

Note this Patent expires in two years 2012, GTRE should wait round the clock and claim it

http://www.patentstorm.us/patents/5264011/fulltext.html

An abrasive system and a processing procedure is provided which permits the direct installation of a thick abrasive blade tip cap onto a cast single crystal turbine rotor blade during a single heating schedule requiring only one furnace operation. The composition of the abrasive blade tip cap advantageously utilizes the high temperature performance capabilities of the single crystal alloy without significantly affecting its mechanical properties as a consequence of the processing necessary to permanently bond the abrasive blade tip cap to the rotor blade. A semi-rigid blade tip cap preform is first formed and positioned on the tip of the rotor blade. The preform and rotor blade are then heated in a vacuum furnace according to a temperature schedule entailing heating rates, holding temperatures and durations which are sufficient to bond and consolidate the preform. The rotor blade is then rapidly cooled in the vacuum furnace to retain the single crystal structure of the rotor blade.
Claims

The embodiments of the invention in which an exclusive property or privilege is claimed are defined as follows:

1. An abrasive blade tip cap preform for bonding to a rotor blade tip to form an abrasive rotor blade tip cap, said abrasive blade tip cap preform comprising:

a metal powder matrix comprising a cobalt alloy and boron; and

abrasive ceramic particles interspersed in said metal powder matrix, said abrasive ceramic particles being coated with a reactive metal;

wherein said boron is present in sufficient amounts to aid in wetting and bonding together said metal powder matrix and said abrasive ceramic particles into a fully densified matrix upon sufficient heating of said abrasive blade tip cap preform.

2. An abrasive blade tip cap preform as recited in claim 1 wherein said metal powder matrix comprises:

a cobalt-base braze alloy and a boron-containing cobalt alloy, said boron being present in said boron-containing cobalt alloy in sufficient amounts to aid in wetting and bonding together said metal powder matrix and said abrasive ceramic particles into a fully densified matrix upon sufficient heating of said abrasive blade tip cap preform;

wherein said abrasive blade tip cap preform comprises about 20 to about 30 weight percent of said cobalt-base braze alloy and about 42 to about 52 weight percent of said boron-containing cobalt alloy.

3. An abrasive blade tip cap preform as recited in claim 1 comprising from about 10 to about 50 weight percent of said abrasive ceramic particles.

4. An abrasive blade tip cap preform as recited in claim 1 further comprising a fluorocarbon powder distributed throughout said metal powder matrix in sufficient amounts to promote a bond between said rotor blade tip cap preform and said rotor blade tip when sufficiently heated.

5. An abrasive blade tip cap preform as recited in claim 4 comprising up to about 5 weight percent of said fluorocarbon powder.

6. An abrasive blade tip cap preform as recited in claim 1 further comprising an organic binder distributed throughout said metal powder matrix in sufficient amounts to improve the green strength of said abrasive blade tip cap preform.

7. An abrasive blade tip cap preform as recited in claim 6 wherein said abrasive ceramic particles are aluminum oxide particles.

8. An abrasive blade tip cap preform as recited in claim 6 comprising from about 1 to about 7 weight percent of said organic binder.

9. An abrasive blade tip cap preform as recited in claim 1 wherein said reactive material is titanium, said abrasive ceramic particles comprising about 2 to about 4 weight percent of said titanium.

10. An abrasive blade tip cap preform for bonding to a rotor blade tip to form an abrasive rotor blade tip cap, said abrasive blade tip cap preform comprising:

a first layer, said first layer comprising:

a metal powder matrix including cobalt and boron; and

abrasive ceramic particles interspersed in said metal powder matrix, said abrasive ceramic particles being coated with a reactive metal;

a second layer comprising a cobalt-base braze alloy; and

an organic binder between said first and second layers;

wherein said boron is present in sufficient amounts to aid in wetting and bonding together said metal powder matrix and said abrasive ceramic particles into a fully densified matrix upon sufficient heating of said abrasive blade tip cap preform.

11. An abrasive blade tip cap preform as recited in claim 10 wherein said metal powder matrix comprises:

a cobalt-base braze alloy and a boron-containing cobalt alloy, said boron being present in said boron-containing cobalt alloy in sufficient amounts to aid in wetting and bonding together said metal powder matrix and said abrasive ceramic particles into a fully densified matrix upon sufficient heating of said abrasive blade tip cap preform;

wherein said abrasive blade tip cap preform comprises about 20 to about 30 weight percent of said cobalt-base braze alloy and about 42 to about 52 weight percent of said boron-containing cobalt alloy.

12. An abrasive blade tip cap preform as recited in claim 10 further comprising a fluorocarbon powder distributed throughout said first layer, said abrasive blade tip cap preform comprising up to about 5 weight percent of said fluorocarbon powder.

13. An abrasive blade tip cap preform as recited in claim 10 further comprising said organic binder throughout said first layer, said abrasive blade tip cap preform comprising from about 1 to about 7 weight percent of said organic binder.

14. An abrasive blade tip cap preform as recited in claim 10 wherein said reactive material is titanium, said abrasive ceramic particles comprising about 2 to about 4 weight percent of said titanium.

15. A rotor blade comprising:

a distal tip on said rotor blade;

an abrasive composition adhered to said distal tip, said abrasive composition comprising;

a metal powder matrix including cobalt and boron; and

abrasive ceramic particles interspersed in said metal powder matrix, said abrasive ceramic particles being coated with a reactive metal;

wherein said boron is present in sufficient amounts to aid in wetting and bonding together said metal powder matrix and said abrasive ceramic particles into a fully densified matrix upon sufficient heating of said abrasive composition.

16. A rotor blade as recited in claim 15 wherein said metal powder matrix comprises:

a cobalt-base braze alloy and a boron-containing cobalt alloy, said boron being present in said boron-containing cobalt alloy in sufficient amounts to aid in wetting and bonding together said metal powder matrix and said abrasive ceramic particles into a fully densified matrix upon sufficient heating of said abrasive composition;

wherein said abrasive composition comprises about 20 to about 30 weight percent of said cobalt-base braze alloy and about 42 to about 52 weight percent of said boron-containing cobalt alloy.

17. A rotor blade as recited in claim 15 wherein said abrasive composition comprises about 10 to about 50 weight percent of said abrasive ceramic particles.

18. A rotor blade as recited in claim 15 wherein said metal powder matrix and said abrasive ceramic particles comprise a first layer of said abrasive composition, said rotor blade further comprising a second layer comprising a cobalt-base braze alloy adhered to said first layer.

19. A rotor blade as recited in claim 18 wherein said metal powder matrix comprises:

a cobalt-base braze alloy and a boron-containing cobalt alloy, said boron being present in said boron-containing cobalt alloy in sufficient amounts to aid in wetting and bonding together said metal powder matrix and said abrasive ceramic particles into a fully densified matrix upon sufficient heating of said abrasive composition;

wherein said abrasive composition comprises about 20 to about 30 weight percent of said cobalt-base braze alloy and about 42 to about 52 weight percent of said boron-containing cobalt alloy.

20. A rotor blade as recited in claim 18 wherein said abrasive composition comprises about 10 to about 50 weight percent of said abrasive ceramic particles.
Description

The present invention generally relates to abrasive materials and methods for adhering abrasive materials to a substrate. More particularly, this invention relates to an improved abrasive material and method for adhering the abrasive material to a turbine blade wherein the method entails a single furnace operation such that the process is particularly suitable for single crystal turbine blades.
BACKGROUND OF THE INVENTION

In the turbine section of a turbine engine, the turbine rotor is circumscribed by a shroud such that the shroud is adjacent the tips of the rotor blades extending from the hub of the rotor. The shroud serves to channel the combustion gases through the turbine section of the turbine engine and prevents the bulk of the turbine engine's combustion gases from bypassing the turbine rotor blades. However, a portion of the gases are able to bypass the rotor blades through a gap present between the rotor blade tips and the shroud. Because the energy of the gases directed through the rotor blades is used to rotate the turbine rotor assembly and any compressor upstream of the turbine section, turbine engine efficiency can be increased by limiting the gases which are able to bypass the rotor blades through this gap.

Manufacturing tolerances, differing rates of thermal expansion and dynamic effects limit the extent to which this gap can be reduced. Any rubbing contact between the rotor blade tips and the shroud will spall the tips of the rotors. Spalling will tend to further increase the gap described above, thereby reducing engine efficiency. In addition, spalling tends to promotes structural fatigue in the rotor blades, causing the useful life of the rotor to be shortened.

As an alternative, it is well known in the art to form a dynamic seal between the rotor blades and the shroud by forming an abrasive tip cap on the end of one or more rotor blades, and more preferably, on each rotor blade. During operation of the turbine, the abrasive tip caps abrade a groove in the shroud as a result of numerous "rub encounters" between the abrasive tip caps and the shroud. The groove, in cooperation with the rotor blade tips as they partially extend into the groove, forms a virtual seal between the rotor blade tips and the shroud. The seal reduces the amount of gases which can bypass the rotor blades, and thereby improves the efficiency of the turbine engine.

Various materials and processes have been suggested to provide a suitable abrasive tip cap on turbine rotor blades. Typical abrasive materials used include silicon carbide, aluminum oxide, tantalum carbide and cubic boron nitride. Aluminum oxide, or alumina, is generally preferred because of its high temperature capabilities and oxidation resistance. In that such abrasive particles do not provide a structurally sound material, they are incorporated with a metal matrix, including for example, nickel or cobalt-base alloys, to provide a sufficiently strong structure which can be bonded to the blade tip. However, the thickness of such a metal matrix is often limited because of the structural weakness of the abrasive composition.
In some applications, it is conventional to apply the abrasive composition to the rotor blade tip using a thermal spray technique, such as plasma spraying or detonation gun spraying. While suitable for many purposes, thermal spray techniques are inefficient in that only part of the abrasive composition contacts and adheres to the rotor blade tip, while much of the thermal spray completely misses the target. More importantly, thermal spraying damages or destroys the morphology of the abrasive particles, making them unsuitable for the intended purpose. In addition, subsequent processes are typically necessary to provide the adhesion and structural integrity necessary for the abrasive composition to survive the hostile environment of a turbine engine. Such steps often include adhering the abrasive composition to the blade tip during a first heating and cooling cycle, and later depositing an additional quantity of the metal matrix over the abrasive composition through a second heating and cooling cycle, such as during hot isostatic pressing. As an alternative, it has also been suggested to melt the tip of the blade, such as with lasers, introduce the abrasive to the blade tip, and then resolidify the blade tip.

While the above processes may be suitable for some turbine blade structures, turbine blades used in modern gas turbine engines are often fabricated from cast high temperature nickel-base superalloys having a single crystal microstructure. Single crystal blades are characterized by extremely high oxidation resistance and mechanical strength at elevated temperatures, which are necessary for the performance requirements of modern turbine engines. However, the single crystal microstructure must not be affected by the process by which the rotor blade abrasive tip caps are secured to the rotor blades. In particular, the process must not recrystallize the single crystal microstructure of the rotor blade, such that the high temperature properties of the rotor blade are lost or diminished. As a result, processes which entail melting the rotor blade tip to the single crystal rotor blade are entirely unacceptable. In addition, repeated thermal cycling of the rotor blades runs the risk of degrading the single crystal microstructure of the rotor blade.

Thus, it would be desirable to provide an abrasive composition which can be readily formed into an abrasive blade tip cap and which can be attached to a turbine rotor blade in a single heating and cooling cycle so as to minimize any degradation of the microstructure of a single crystal turbine rotor blade.

SUMMARY OF THE INVENTION

It is an object of this invention to provide a method for attaching an abrasive blade tip cap to a rotor blade in a single heating and cooling cycle to preserve the microstructure of a single crystal rotor blade.

It is a further object of this invention that such a method include the formulation of an abrasive blade tip can preform which can be adhered as a unit to a rotor blade during a single heating and cooling cycle.

Lastly, it is another object of this invention that such an abrasive blade tip cap preform be of sufficient thickness so as to provide sufficient stock for machining the abrasive blade tip cap to tolerance while retaining adequate thickness to perform repeated rub encounters with a turbine engine shroud over the life of the turbine engine.

In accordance with a preferred embodiment of this invention, these and other objects and advantages are accomplished as follows.

According to the present invention, there is provided an abrasive composition and a process for attaching the abrasive composition to a rotor blade of a turbine engine, wherein the process entails a single heating cycle to adhere the abrasive composition to the tip of the rotor blade. As a result, the process is particularly suitable for forming abrasive tip caps for single crystal rotor blades, such as those formed from nickel-base superalloys.

The abrasive composition is preferably formed as a mat from which a rotor blade tip cap preform can be readily sized and shaped to fit the shape of the rotor blade tip. The abrasive composition generally includes a metal powder matrix containing a cobalt-base braze alloy and a cobalt alloy containing boron in sufficient amounts to aid in wetting and bonding together all of the preform constituents into a fully densified matrix. Ceramic abrasive particles, and preferably aluminum oxide particles, are interspersed in the metal powder matrix. The ceramic abrasive particles are coated with a thin layer of a reactive metal, such as titanium, which serves as a wetting agent to promote a metallurgical bond between the abrasive particles and the metal powder matrix.

More specifically, the rotor blade abrasive tip cap preform includes a first layer which contains the ceramic abrasive particles interspersed in the metal powder matrix, and a second layer formed from the cobalt-base braze alloy alone. A binder is provided between the first and second layers to adhere the second layer to the first layer. In addition, it is preferable that a binder and a fluorocarbon powder be distributed throughout the first layer. The binder serves to impart green strength to the preform while the fluorocarbon powder serves to create a cleansing action between the matrix and the abrasive particles and between the preform and the bonding surfaces during the heating process to ensure a sufficient bond between the preform and the blade tip. Finally, it is also preferable to temporarily attach a layer of cobalt-based braze tape to the blade tip using a suitable transfer tape prior to metallurgically bonding the preform to the blade tip with a furnace operation.

The bonding process of this invention consists of a single heating and cooling cycle so as to minimize degradation of the single crystal microstructure of the rotor blade alloy. The abrasive composition of the present invention is specifically formulated to take advantage of the high temperature capabilities of the single crystal rotor blade such that a single heating cycle will suitably bond the preform to the rotor blade. The bonding process and abrasive composition of the present invention permit a thick abrasive rotor blade tip cap to be permanently secured to the rotor blade tip. As a result, even after grinding to the final dimensions, the abrasive rotor blade tip cap possesses the capability for long service life, and provides the requisite rubbing action with the shroud during the operation of a turbine engine to form a seal between the rotor blades and the shroud.

The preferred heating schedule includes heating the preform, which is mounted on the rotor blade tip, in a vacuum to about 1100° F. for a duration sufficient to volatilize the binders and fluorocarbon powder but at a rate sufficiently low to prevent porosity in the abrasive composition due to outgassing. The preform and rotor blade are then further heated to about 2000° F. at a rate sufficient to maintain vacuum pressure no greater than about 1×10-4 torr, and for a duration sufficient to thermally stabilize the rotor blade prior to reaching bonding temperature. The temperature of the preform and rotor blade is then further raised at a rate sufficient to prevent liquation of the braze tape, and held at about 2225° F. for a duration sufficient to braze and consolidate the abrasive tip preform to the rotor blade tip. Finally, the preform and rotor blade are furnace cooled to a temperature of about 2000° F., after which the preform and rotor blade are further cooled to below about 1400° F. at a rate sufficient to maintain the desired microstructure and strength of the single crystal rotor blade.

Using this preferred heating schedule, the abrasive particles are tightly bonded within the metal powder matrix, and the rotor blade tip cap is tightly bonded to the rotor blade such that the rotor blade tip cap forms a structurally integral portion of the rotor blade in terms of strength and durability. The process of the present invention permits relatively thick preforms to be bonded to the rotor blades such that subsequent machining of the rotor blade tips in the assembled rotor can be performed to bring the rotor assembly into tolerance, while also ensuring that sufficient abrasive material will remain to provide repeated rub encounters over the life of the turbine engine.

In addition, the heating schedule ensures that alteration of the single crystal microstructure is minimized, such that the high temperature capabilities of the superalloy will remain. The single furnace operation involved in the heating schedule of the present invention also provides a significant economic advantage from the standpoint of time, labor and energy requirements. As a result, the present invention is highly suitable for the mass production of turbine rotor blades.

Other objects and advantages of this invention will be better appreciated from the following detailed description.

BRIEF DESCRIPTION OF THE DRAWINGS

The above and other advantages of this invention will become more apparent from the following description taken in conjunction with the accompanying drawing wherein:

FIG. 1 shows an exploded view of the component details of a turbine rotor blade and abrasive rotor blade tip cap in accordance with this invention;

FIG. 2 shows a side view of a turbine rotor blade on which there has been attached a rotor blade tip cap in accordance with this invention; and

FIG. 3 shows a representative illustration of a cross-sectional microphotograph at 50× magnification of the turbine rotor blade of FIG. 2.

DETAILED DESCRIPTION OF THE INVENTION

An abrasive system and a processing procedure is provided which permits the direct installation of a thick abrasive blade tip cap onto a single crystal cast turbine rotor blade during a single heating schedule that accordingly requires only one furnace operation. The composition of the abrasive blade tip cap advantageously utilizes the high temperature performance capabilities of the single crystal rotor blade 12 without significantly affecting the mechanical properties of the single crystal rotor blade 12 which would arise as a consequence of the processing necessary to permanently bond the abrasive blade tip cap to the rotor blade.

The composition of the preferred abrasive blade tip cap and the preferred heating schedule are particularly adapted for cast single crystal nickel-base superalloys, such as used in the rotor blade 12. The preferred nickel-base superalloy consists of, by weight, about 10 percent tungsten, about 10 percent cobalt, about 9 percent chromium, about 5.5 percent aluminum, about 1.5 percent tantalum, about 1.5 percent titanium, about 1.0 percent hafnium, about 0.02 percent boron, about 2.5 percent molybdenum, about 0.15 percent carbon, and about 0.05 percent zirconium, with the balance being nickel. Such an alloy is commercially available from Cannon-Muskegon under the trade designation CMSX-3. However, it is foreseeable that other suitable nickel-base alloys, as well as cobalt or iron-base alloys, could be substituted with similar results.

In accordance with the preferred embodiment of this invention, an abrasive blade tip cap preform 10 is brazed to a single crystal turbine rotor blade 12 (FIG. 1) to form an abrasive blade tip cap 30 of an abrasive rotor blade 14 (FIG. 2). As seen in FIG. 1, the rotor blade 12 has a tip portion 16 which is remote from the rotor blade's base 18 by which the rotor blade 12 is mounted to a rotor hub (not shown) to form a turbine rotor assembly (not shown). The tip 16 of the rotor blade 12 is substantially flat, though having a compound curvature referred to as the arc drop. The arc drop of the blade tip 16 results from the rotor blade 12 being ground to conform to the cylindrical internal surface of the turbine shroud (not shown).

The abrasive blade tip cap preform 10 can be attached to one or more rotor blades 12 according to the method of the present invention, though in the preferred embodiment, each rotor blade 12 would have an abrasive blade tip cap preform 10 bonded thereto. With the abrasive rotor blades 14 mounted to the hub within a turbine engine (not shown), the abrasive-blade tip caps 30 will be proximate to the shroud which circumscribes the turbine rotor assembly. The abrasive blade tip caps 30 serve to wear-form a seal track in the shroud, resulting in a virtual seal between the abrasive rotor blades 14 and the shroud which substantially prevents combustion gasses from bypassing the rotor assembly. A particular aspect of the preferred composition of the abrasive blade tip caps 30 is the ability to withstand repeated and severe rub encounters with the shroud, with only minimal loss of material from the abrasive blade tip caps 30 and preferential wear of the shroud material.

The preferred abrasive composition from which the abrasive blade tip cap preforms 10 are formed includes a metal powder matrix combined with ceramic abrasive particles. The abrasive particles are preferably about 80 to 120 mesh grit aluminum oxide particles which are coated with a reactive metal. The coated aluminum oxide particles preferably make up about 24 to about 28 weight percent, and more preferably about 26.3 weight percent, of the abrasive blade tip cap preform 10, though it is foreseeable that the coated aluminum oxide could be present in quantities of as little as 10 or as great as 50 weight percent.

Most preferably, the reactive metal coating on the aluminum oxide particles is a titanium coating which constitutes about 2 to about 4 weight percent of the coated aluminum oxide particles. As a reactive metal, the titanium serves to wet the surface of the aluminum oxide particles to promote a metallurgical bond between the particles and the metal powder matrix. Although titanium is preferred because it is known to react to both aluminum oxide and the matrix to form a metallurgical bond, other reactive metals could also be used. It is preferable that the titanium coating be applied using known fluidized bed chemical vapor deposition techniques so as to ensure uniformity of the coating on the particles, though other suitable processes known in the art are acceptable.

The metal powder matrix is a mixture of a cobalt-base braze alloy combined with a cobalt alloy that includes boron. The cobalt-base braze alloy is preferably Aerospace Material Specification 4783 (AMS4783) having a nominal composition by weight of about 8 percent silicon, about 19 percent chromium, about 17 percent nickel, about 4 percent tungsten, about 0.8 percent boron, with the balance being cobalt. The cobalt-base braze alloy is preferably provided in particle form and has a particle size no greater than about 325 mesh. The cobalt-base braze alloy makes up about 22 to about 26 weight percent, and more preferably about 24.1 weight percent, of the abrasive blade tip cap preform 10, though it is foreseeable that the cobalt-base braze alloy could be present in quantities of as little as about 20 or as great as about 30 weight percent.

The boron-containing cobalt alloy is a proprietary composition manufactured by Union Carbide Specialty Powders of Indianapolis, Ind., and designated as Alloy No. CO-274. Boron is present in sufficient amounts to aid in wetting and bonding together all the preform constituents into a fully-densified matrix. Similar to the cobalt-base braze alloy, the boron-containing cobalt alloy is provided in particle form and has a particle size no greater than about 325 mesh. In the preferred embodiment, the boron-containing cobalt alloy makes up about 46 to about 50 weight percent, and more preferably about 48.2 weight percent, of the abrasive blade tip cap preform 10, though it is foreseeable that it could be present in quantities of as little as 42 or as great as 52 weight percent.

In addition, the abrasive blade tip cap preform 10 preferably includes a fluorocarbon powder distributed throughout the metal powder matrix. The fluorocarbon powder is preferably a polytetrafluoroethylene (PTFE) micropowder, marketed as Product No. MP 1100 and available from E.I. DuPont de Nemours and Company, Inc., Polymer Products Department, of Wilmington, Del. The fluorocarbon powder preferably makes up about 1.0 to about 2.0 weight percent, and more preferably 1.4 weight percent of the abrasive blade tip cap preform 10. However, the fluorocarbon powder may be omitted completely or provided in amounts as great as about 5 weight percent, in that the desired amount is dictated only by the need to provide a final cleansing of the matrix and abrasive particle surfaces and the bonding surfaces of the preform 10 and the rotor blade tip 16 during the high temperature consolidation of the preform structure, which will be described in detail below, and without producing excessive porosity in the abrasive blade tip caps 30.

As a final preferred additive, the preform 10 may include a binder which is distributed throughout the metal powder matrix in sufficient quantities to impart green strength to the preform 10. Such binders are well known in the art. However, the preferred embodiment utilizes a proprietary organic binder, Allison Type GAB/Production, available from Vitta Corporation of Bethel, Connecticut. The binder preferably makes up between about 2 and about 5 weight percent of the abrasive blade tip cap preform 10, though it is foreseeable that the binder can be present in quantities of as little as about 1.0 and as great as about 7.0 weight percent, in that the desired amount is dictated only by the amount of green strength desired in the abrasive blade tip cap preform 10 and the level of porosity resulting from binder volatilization which can be allowed in the consolidated abrasive blade tip cap 30.

The preform 10 is preferably cut or punched from a mat of uniform thickness and according to the shape of the blade tip 16, with an allowance being made for shrinkage during consolidation of the preform 10 as a result of the bonding process. The preform 10 has two distinct layers, as best seen in FIG. 1. The first layer 20 consists of the metal powder matrix and the titanium-coated abrasive particles combined with the fluorocarbon powder and the binder, while the second layer 22 is comprised only of the AMS 4783 cobalt-base braze alloy which serves as a reservoir to replace the organic constituents volatilized during the consolidation heating cycle, and thereby minimize the resulting porosity which would otherwise be created. The second layer 22 is nominally about 0.003 to about 0.004 inches thick, and the preform 10, combining the first and second layers 20 and 22, has a thickness of about 0.059 to about 0.065 inch. However, the thickness of the preform mat may vary substantially, depending on the blade design. Generally, mat thickness is dictated by desired cutting life and machining stock required. Due to the thickness of the preform 10 and the green strength contributed by the binder, the preform 10 is sufficiently rigid to permit handling under most manufacturing conditions. However, the binder is not otherwise an essential ingredient in the preform 10, in that the binder is volatilized during the bonding process to be described below.

The environment in which the abrasive blade tip cap preform 10 is applied to the rotor blade 12 must be clean to prevent contamination of the bonding surfaces of either the rotor blade 12 or the preform 10. The procedure for applying the preform 10 to the rotor blade 12 includes forming a bonding surface on the rotor blade tip 16 which is ground smooth with no edge breaks. The rotor blade 12 is further prepared by being degreased with a suitable solvent or detergent of a type well known in the art. The rotor blade 12 is then masked to expose only the tip 16 of the rotor blade 12, which serves as the bonding surface. The blade tip 16 is then blasted using a blasting medium, such as a nickel-base blasting medium sold under the name NICROBLAST MEDIA by Wall Colmonoy Corporation of Madison Heights, Michigan. Such a nickel-base blasting medium is preferred because it leaves a nominal nickel layer (less than about 0.0001 inch) on the rotor blade tip 16 which serves to wet its bonding surface and thereby promote bonding of the preform 10 during consolidation. However, it is foreseeable that other blasting mediums known to those skilled in the art can be used with acceptable results. Thereafter, the entire rotor blade 12 is flushed with dry, filtered air to remove any excess blasting medium.

Preferably, a one eighth inch band of stop-off, such as NICROBRAZ-GREEN STOPOFF, a product of Wall Colmonoy Corporation of Madison Heights, Michigan, is then applied to the rotor blade 12 surfaces surrounding the bonding surface to prevent brazing at these regions. A braze tape 24, and more preferably a cobalt-base braze tape comprised of the aforementioned AMS 4783 cobalt-base braze alloy having a thickness of about 0.004 inches, is then applied to the blade tip 16 using a suitable transfer tape 26, such as type 9710 Transfer Tape, a product of 3M Company of St. Paul, Minn.

The previously cut preform 10 is then temporarily attached to the tip 16 of the rotor blade 12 using a suitable transfer tape 28. The preform 10 and rotor blade 12 are then further readied for bonding and consolidation by orienting the rotor blade 12 vertically such that the preform 10 rests on top of the tip 16 of the rotor blade 12. The bonding and consolidation process must be performed in a clean, out-gassed vacuum furnace, with the furnace preferably being evacuated to a pressure of no more than about 1×10-4 torr.

The preferred heating schedule between room temperature and about 1100° F. is determined by the rate at which the binder and fluorocarbon powder will volatilize. In particular, the binder and fluorocarbon powder must not volatilize at a rate which will produce porosity or distortion in the preform 10, and thus the final abrasive blade tip cap 30. Therefore, the heating rate is purposely chosen to be relatively slow so as to allow complete diffusion of the volatilized gasses through the preform 10 material without creating porosity in the preform 10.

Once the fluorocarbon powder and binder have been volatilized (i.e., above about 1100° F.), the heating schedule is determined by the need to bond and consolidate the preform 10 while also preserving the single crystal structure of the supporting nickel-base superalloy turbine blade 12. While specifically adapted to the property limitations of the preferred nickel-base superalloy, the composition of the preform 10 and the preferred heating schedule described below may also be applicable to other single crystal alloys where mechanical property requirements are met in conjunction with the constraints imposed by the required bonding cycle.

The heating rates, durations and limits for the preferred heating schedule are detailed in the table below.

TABLE I ______________________________________ PREFERRED FURNACE SCHEDULE RATE TEMPERATURE/TIME ______________________________________ Heat 10° F./minute Room temperature to 500 +/- 25° F. (max) Hold at: 500 +/- 25° F. for about 5 minutes Heat 3° F./minute 500 +/- 25° F. to 750 +/- 25° F. (max) Hold at: 750 +/- 25° F. for about 5 minutes 6° F./minute 750 +/- 25° F. to 900 +/- 25° F. (max) Hold at: 900 +/- 25° F. for about 5 minutes Heat 3° F./minute 900 +/- 25° F. to 1100 +/- 25° F. (max) Hold at: 1100 +/- 25° F. for about 10 minutes Heat 15° F./minute 1100 +/- 25° F. to 2000 +/- 25° F. (min) Hold at: 2000 +/- 25° F. for about 10 minutes Heat 30° F./minute 2000 +/- 25° F. to 2225 +/- 15° F. (max) Hold at: 2225 +/- 15° F. for about 110 to about 130 minutes Vacuum or gas cool 2225 +/- 15° F. to 2000 +/- 25° F. Gas fan cool 2000 +/- 25° F. to 1400 +/- 25° F. ______________________________________ Note: Temperatures given are set points; +/- tolerances are the required contro ranges.

The preferred bonding and consolidation process includes heating the preform 10 and rotor blade 12 in a vacuum of no more than about 1×10-4 torr. The temperatures and durations indicated above are selected to perform the following. First, the preform 10 and rotor blade 12 are heated to a temperature of about 1100° F.+/-25° F. at a rate and for a duration which will be sufficient to ensure complete diffusion of the volatilized gasses through the preform 10. As shown in Table I, intermediate holding temperatures of 500° F., 750° F. and 900° F. are preferred to prevent porosity formation, but these intermediate holding temperatures are not absolutely necessary. The preform 10 and rotor blade 12 are preferably held at about 1100° F.+/-25° F. for about 10 minutes, which is sufficient to prevent porosity within the preform 10.

The temperature of the preform 10 and rotor blade 12 is then further raised at a rate of about 15° F. per minute minimum, which is sufficient to minimize the exposure of the rotor blade 12 to high temperature, and held at about 2000° +/-25° F. for about 10 minutes, a duration which is sufficient to thermally stabilize the rotor blade 12. Thereafter, the temperature of the preform 10 and rotor blade 12 is further raised at a rate of about 30° F. per minute maximum, which is sufficient to prevent liquation of the AMS 4783 braze tape 24, and held at about 2225° +/-15° F. for about 110 to about 130 minutes, a duration which is sufficient to melt the metal powder matrix. The molten metal powder matrix forms a liquid phase which surrounds the abrasive particles and wets the rotor blade tip 16. In addition, the molten metal powder matrix wets and reacts with the titanium coating on the abrasive particles in a manner that produces a strong metallurgical bond upon cooling.

In particular, it is believed that the titanium in immediate contact with the aluminum oxide surface bonds to oxygen in the aluminum oxide, essentially becoming a part of the oxide structure. Titanium which is located in the coating further from the aluminum oxide particle remains metallic. Because of its metallic nature, the titanium coming in contact with the molten metal powder matrix improves wetting and probably alloys itself with the braze alloy. Thus, both the abrasive particle-titanium interface and the titanium-matrix interface are strengthened by chemical bonding so that the overall bond between the abrasive blade tip cap 30 and the rotor blade 12 is stronger than mere mechanical joining.

FIG. 3 is a 50× magnification of the single crystal abrasive rotor blade 14 and abrasive blade tip cap 30 after bonding. In the abrasive blade tip cap 30, there can be seen aluminum oxide abrasive particles 32 and the molten metal powder matrix 34, which consists essentially of the cobalt-base braze alloy (AMS 4783) and the boron-containing cobalt alloy. The bond joint interface using the AMS 4783 braze tape 36 is also shown.

The above heating rates, temperatures and durations are recommended for the preferred nickel-base superalloy, such as is used in the rotor blade 12. In addition, it is believed that the brazing temperature could vary between about 2210° F. and about 2240° F., while still achieving adequate results--i.e., minimum porosity in the abrasive blade tip cap 30 without degradation of the rotor blade's single crystal microstructure. However, the above temperatures are preferred for the particular combination and proportions of materials used. In addition, it is foreseeable that suitable results could also be obtained with holding durations which are outside of the preferred range, such as between about 110 minutes up to about 130 minutes, although the preferred range is favored since it provides the desired results within a practical production schedule.

Following the above heating steps, the preform 10 and rotor blade 12, now as the unitary abrasive rotor blade 14, are gas cooled, such as by flowing an inert gas within the furnace chamber, to a temperature of about 2000° +/-25° F. Gas cooling to this temperature is preferred because it ensures solidification of the abrasive blade tip cap 30 prior to gas fan cooling. Thereafter, the abrasive rotor blade 14 is gas fan cooled to below about 1400° F. at a rate of at least 50° F. per minute, which is sufficient to maintain the desired single crystal structure of the rotor blade 12 and resultant strength level of the rotor blade 12. The abrasive rotor blade 14 is then cooled below this temperature to room temperature by gas fan cooling or by furnace cooling. The rate of cooling below about 1400° F. does not appear to be critical to the success of this invention.

The abrasive rotor blade 14 is then assembled, along with other abrasive rotor blades 14 and possibly uncapped rotor blades 12, to a turbine wheel or other appropriate fixture and ground to the final dimensions using silicon carbide or diamond grinding wheels, following machining parameters which are generally well known in the art. In addition, the surface of the abrasive blade tip cap 30 may be chemically or electrochemically etched to better expose the abrasive particles 32 to improve initial abrasiveness.

The relatively thick (equivalent to multiple abrasive particle diameters) abrasive blade tip cap 30 provides sufficient stock for machining while retaining adequate thickness to accommodate repeated rub encounters over the life of the turbine engine. This feature is contrary to the teachings of the prior art, wherein an abrasive cap has a thickness equivalent to only one grit particle which is applied to a finish-machined rotor blade. As a result, significant assembly and disassembly operations are typically necessary because the application environment may be detrimental to some components of the rotor assembly, and the abrasive blade tip cap has a significantly shorter service life due to its limited thickness.

It has been determined that the heating schedule of this invention is capable of sufficiently consolidating and bonding the preform 10 to the rotor blade 12 with minimal loss to the integrity of the single crystal structure of the single crystal nickel-base superalloy turbine rotor blade 12. As a result, the high temperature properties of the superalloy are essentially retained, a critical factor in the environment of a modern turbine engine.

It should also be noted that the preform 10, once consolidated to form the abrasive blade tip cap 30 during the above heating schedule, is characterized as having sufficient structural and bond strength to survive the high rotational speeds and temperatures of a turbine engine and numerous rub encounters with the engine shroud. Specifically, the particular composition of the preform 10 is able to fully utilize the various stages of the heating schedule to complete the consolidation and bonding processes. In addition, the abrasive blade tip cap 30 is inherently corrosion resistant due to the presence of cobalt as the primary constituent of the metal powder matrix 34.

It is a particular feature of the present invention that the abrasive particles 32 are tightly bonded within the metal powder matrix 34, and that the active metal coating on the abrasive particles 32 serves to wet the surface of the abrasive particles 32 so as to promote a metallurgical bond between the abrasive particles 32 and the metal powder matrix 34. As a result, the retention and durability of the abrasive particles is enhanced. The metal powder matrix 34 is formulated to provide bond strength between individual abrasive particles 32 and between the abrasive blade tip cap 30 and the rotor blade 12, while also providing corrosion resistance. The bond strength is characterized as being sufficient to meet the tensile strength necessitated by the numerous rub encounters demanded of the abrasive blade tip cap 30. In addition, the bond strength is further enhanced by the fluorocarbon powder which provides a final cleansing action to the bonding surfaces of both the preform 10 and the rotor blade 12 during the consolidation and bonding cycles.

Moreover, prior to heating, the preform 10 is sufficiently rigid and has predictable shrinkage characteristics to permit a simple and economical punching or cutting operation to form the preform 10. The process of the present invention also permits relatively thick preforms 10 to be formed and bonded to the rotor blades 12 such that subsequent machining of the abrasive blade tip cap 30 can be performed to bring the rotor blade 12 into tolerance while ensuring that sufficient abrasive material will remain to provide repeated rub encounters over the life of the turbine engine.

Finally, as a primary advantage of the present invention, the heating schedule employed ensures that alteration of the single crystal microstructure is minimized, such that the high temperature capabilities of the superalloy will remain intact. The single furnace operation involved in the heating schedule of the present invention also provides a significant economic advantage from the standpoint of time, labor and energy requirements. As a result, the present invention is highly suitable for the mass production of turbine rotor blades.

While our invention has been described in terms of a preferred embodiment, it is apparent that other forms could be adopted by one skilled in the art; for example by substituting other matrix or braze compositions, such as nickel-base alloys, for the preferred cobalt-base compositions, other abrasive materials, such as cubic boron nitride, titania, zirconia or chromium carbide, for the preferred aluminum oxide, omission of either the fluorocarbon powder or binder for less demanding applications, or the use of mixtures of abrasive materials or grit sizes. Accordingly, the scope of our invention is to be limited only by the following claims.

* * * * *
Inventors
Brown, Lawrence E.
Clingman, David L.
Barber, Michael J.
Cross, Kenneth R.

Assignee
General Motors Corporation

Application
No. 941618 filed on 09/08/1992

US Classes:
51/309, Metal or metal oxide29/889.2, Turbomachine making51/295, IMPREGNATING OR COATING AN ABRASIVE TOOL51/298, WITH SYNTHETIC RESIN415/173.6, Between blade supported radial tip ring and static part416/241BCeramic material

Field of Search
51/293, MISCELLANEOUS51/295, IMPREGNATING OR COATING AN ABRASIVE TOOL51/298, WITH SYNTHETIC RESIN51/309, Metal or metal oxide416/223A, Turbo machine416/241R, Coating, specific composition or characteristic416/241BCeramic material

Examiners
Primary: Bell, Mark L.
Assistant: Thompson, Willie J.

Attorney, Agent or Firm
Brooks; Cary W., Hartman; Domenica N. S.

US Patent References
3850590, 4142872, Metal bonded abrasive tools
Issued on: 03/06/1979
Inventor: Conradi4249913, Alumina coated silicon carbide abrasive
Issued on: 02/10/1981
Inventor: Johnson , et al.4378975, Abrasive product
Issued on: 04/05/1983
Inventor: Tomlinson , et al.4591364, Abrasive materials
Issued on: 05/27/1986
Inventor: Phaal4610698, Abrasive surface coating process for superalloys Issued on: 09/09/1986
Inventor: Eaton , et al.4741973, Silicon carbide abrasive particles having multilayered coating
Issued on: 05/03/1988
Inventor: Condit , et al.4802828, Turbine blade having a fused metal-ceramic tip
Issued on: 02/07/1989
Inventor: Rutz , et al.4854196, Method of forming turbine blades with abradable tips
Issued on: 08/08/1989
Inventor: Mehan5096465Diamond metal composite cutter and method for making same
Issued on: 03/17/1992
Inventor: Chen, et al.
**********
I was suggesting the spray techinque over a RCC like blade (ofcourse not cement) but Carbon Carbon fibre with the spray

vera_k · Post by **vera_k** » 08 Aug 2009 09:31

John Snow wrote:Note this Patent expires in two years 2012, GTRE should wait round the clock and claim it

What's there to claim? They can use whatever info they find in the patent. It is not as if the patent owner can sue for royalties. Of course, the really important stuff is never patented.

Raj Malhotra · Post by **Raj Malhotra** » 08 Aug 2009 09:38

As a thumb rule, an engine requires twice the value of its cost in spare parts over its lifetime. Hence 600 million dollar order is equivalent to US$ 2 Billion order.

N3-why don't you write an article/post on what to do NOW! for Indian engine tech developement to cover things like MCA, UCAV, AJT, MTA, MLH, RTA, Naval & Indutrial which is equivalent to US$ 3 Billion imports per annum etc. Pls give a timeline-or say a direction per you, in english please.

John Snow · Post by **John Snow** » 08 Aug 2009 09:45

BHEL already make Industrial Gas turbines here it is.

BHEL - the largest Gas Turbine manufacturer in India, with the state-of-art facilities in all areas of Gas Turbine manufacture provide complete engineering in-house for meeting specific customer requirement. With over 100 machines and cumulative fired hours of over four million hours, BHEL has supplied gas turbines for variety of applications in India and abroad. BHEL also has the world's largest experience of firing highly volatile naphtha fuel on heavy duty gas turbines.

BHEL is one of the few business associates of M/s. GE, USA and under its comprehensive Technical Collaboration Agreement from 1986, (license to manufacture rotors and hot gas path components) offers complete power plant engineering solutions. BHEL also manufactures Gas Turbine under its ongoing Technical Collaboration Agreement with M/s Siemens, Germany. Some specific features:

1. Capability to fire a wide range of gaseous and liquid fuels and a mix of such fuels ranging from clean fuels like Natural Gas, Distillate Oil, Naphtha, LNG to heavy fuels like LSHS, crudes, blends, etc. Fuel economy over a wide range of ambient temperatures and loads.

2. Facilities like Black start, fast start and emergency start.

3. Suitable for power generation and mechanical drive application. Models below 100 MW suitable for 50 Hz and 60 Hz.

4. All matching equipment like generators, compressors, etc. manufactured in-house. Design of combustion systems as per international emission norms. Machines designed as per major international codes like API, etc.

5. Suitable for IGCC applications.
6. Suitable for indoor or outdoor locations.

7. Use of water or steam injection for abatement of NOX emissions and power augmentation.

BHEL equipped with precision and sophisticated machine - tools like CNC Broaching Machine, 5-Axis Milling Machine and over speed Vacuum Balancing Tunnel offers Conversion, Modification and Up-gradation services - through joint venture with GE for all existing Gas turbines. Services are also offered for all Field support, Retrofits and repairs, inspections and Technical Consultancy on "Operation & Maintenance of Gas Turbine Based Power Plants

John Snow · Post by **John Snow** » 08 Aug 2009 10:20

Here one more

http://www.azom.com/details.asp?ArticleID=90

Thermal Barrier Coatings
Thermal barrier coatings have been used for some years on static parts, initially using magnesium zirconate but more recently yttria-stabilised zirconia. On rotating parts, the possibility of ceramic spalling is particularly dangerous, and strain‑tolerant coatings are employed with an effective bond coat system to ensure mechanical reliability.

Ceramic Matrix Composites
Further increases in temperature are likely to require the development of ceramic matrix composites. A number of simply shaped static components for military and civil applications are in the engine development phase and guide vanes have been manufactured to demonstrate process capability, such techniques involve advanced textile handling and chemical vapour infiltration.

However, it is the composite. ceramic rotor blade that provides the ultimate challenge. It will eventually appear because the rewards are so high, but it will take much longer to bring it to a satisfactory standard than was anticipated in the 1980’s. Research work has concentrated for some years on fibre reinforced ceramics for this application, as opposed to monolithic materials which possess adequate strength at high temperatures but the handicap of poor impact resistance.

Today's commercially available ceramic composites employ silicon carbide fibres in a ceramic matrix such as silicon carbide or alumina. These materials are capable of uncooled operation at temperatures up to 1200°C, barely beyond the capability of the current best coated nickel alloy systems. Uncooled turbine applications will require an all oxide ceramic material system, to ensure the long term stability at the very highest temperatures in an oxidising atmosphere. An early example of such a system is alumina fibres in an alumina matrix. To realise the ultimate load carrying capabilities at high temperatures, single crystal oxide fibres may be used. Operating temperatures of 1400°C are thought possible with these systems.

quote]http://www.tms.org/Meetings/Annual-97/P ... 2C.XI.html[/quote]

I think BARC should be put to task also for not getting this SCB going.

asprinzl · Post by **asprinzl** » 08 Aug 2009 11:41

Narayanan,
On the example dialog between the prof's as they listened to your presentation....wow their attitude is depressing. If Israel's leaders had such attitude or mentality, we would have perished preety soon after 1947. Amazingly depressing.

As per the discussion on GTRE and single crystal blade technology.....AQK was able to walk away with the blue prints and technology of uraniun enrichment.....a technology that was more sensitively guarded and protected.........I am sure it would not be impossible for India to obtain the single crystal technology and its manufacturing blue prints.
Avram

vasu_ray · Post by **vasu_ray** » 08 Aug 2009 20:55

instead of changing PSU culture, why not the GoI make infrastructure available to private organizations, research institutions, third parties and universities?

similar to IITs time sharing the CDAC made PARAM computers, many folks can share these national infra facilities for a fee, it provides a big boost to smaller companies and technology start ups increasing the number of players per opportunity

With funding barriers removed, professionals can find more opportunities in avenues other than IT

vina · Post by **vina** » 08 Aug 2009 22:32

instead of changing PSU culture, why not the GoI make infrastructure available to private organizations, research institutions, third parties and universities?

You post IT/Vity generation need to get a grip on how things were like 15 to 20 years ago. The private sector was pretty much something the cat brought in, with the fond hope being that in time it would wither away /die/get killed/whatever and we will have a glorious public sector in everything, literally (except for road side chai stalls or push cart vegetable vending) basically anything of consequence. In the industrial area everything was done to actively kill private initiative.

Take something as simple as someone running a high speed boat service between Bombay and Goa. When some private guy started a hydrofoil service between Bombay and Goa , babus did their damnest to try and kill it. Basically unleash the entire weight of red tape and license inspector raj (what finally killed it was I think Konkan rail and the subsidized train fare), but think of it, instead of flying, you get in at Bombay and in a couple of hours you are in Goa!.

While post 1990, economic reforms and IT/Vity has effectively killed such thinking, in defense and other still "protected" /"reserved" areas such thinking is deep rooted and is still the norm. No way in hell will any defense lab or anything built using public money be opened up to wide spread access to all of indian industry. The public sector guys will lose the only remaining source of advantage and control.

vasu_ray · Post by **vasu_ray** » 08 Aug 2009 23:07

Another example would be trying to run trains owned by private corps on national rail network, Indian Railways would enjoy the right of way to the detriment of of private ones meeting any higher standards, IR will not allow it anyways

as a counter argument, take civil aviation, airports built by the govt. are shared by private airlines, even the ones meant for IAF : the reason being public interest

RTI was against babu's interest, it went through after sometime : public interest, I don't know the legalese, but if it happens, it will definitely be a coup in the right direction

vina · Post by **vina** » 11 Aug 2009 15:53

Hmm. Lot of action sure in the skies over bangalore. In the mornings, I see a Jag doing regular flights. Just some 30 mins or so ago , I think I heard the distinctive noise of YellCeeYea.

Wonder what is going on. No hard news on this front for quite a while now.