Russian 5th-gen Su-57 fighter jet performs maiden flight with new engine:
PAK-FA / Sukhoi Su-57 - Updates and Discussions
- Thread starter Ashwin
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Yeah we aren't buying Su-57, cant trust the Ruskies anymore they are double crossing us.
Yup true, specially since they have compromised our MKI frequencies.
At first glance the engine looks to be shorter then AL-41 with crossed nozzles.Might take a bit to load up.
Media debut of Sukhoi Su-57 AAM launch from Internal Weapon Bay... High quality animation or reality, still unconfirmed...
Истребитель пятого поколения получил название Су-57
Истребитель пятого поколения получил название Су-57
I guess they used one new engine and one old engine. THough I doubt its possible but there are clear visible differences in engine.
Yes, could be an illusion because of the angle but they did promised the engine to be 15-20 percent lighter.
Of course. A new engine is always tested on a twin engine aircraft where one of the engines is already proven. So, if the new engine fails, the proven engine will bring the aircraft back safely.
Item-30 will be of similar weight compared to Al-41 but it designed to deliver 15-20% more power output than Al-41 powerplant...
Some details from a Russian Forum. Translator doesnt work very well but it has lot of information there.
… I would say that the Изделие 30 engine has larger diameter than the Saturn AL-41F-1 (изд. 117) engine (fan diameter: 932,0 mm / 36,7 in). I remember I have read somewhere that the Saturn AL-41F (Изделие 20) engine, once designed for the propulsion of MiG 1.44 MFI, had the fan diameter of 1.000,0 mm / 39,4 in. I don´t know if that was correct data and if Izdeliye 30 engine has the same one, but there is no doubt that the temperatures, pressures, N1 and N2 speeds, mass flow of the Изделие 30 engine were heavily increased compared to the Saturn AL-41F-1 (изд. 117) engine. Namely, there are two basic ways to increase core´s power: hot route - increasing HP turbine´s inlet temperature (TIT) and cold route - increasing core´s mass flow. The larger fan´s airflow requires (I guess Izdeliye 30 engine has larger BPR than the Izdeliye 117 engine) more power from the engine´s core. This can be achieved by raising the OPR (combustor inlet pressure / intake delivery pressure) just to induce more airflow into the core and by increasing TIT. Together, these parameters tend to increase core´s thermal efficiency and improve engine´s SFC.
The desired OPR (overall pressure ratio) for the engine´s Brayton cycle is usually achieved by the multiple axial stages on the core´s compression section. Let me explain in just a few words how that functions at civil airliners´engines. F.e. Rolls-Royce tend to split the core´s compression with an IP compressor supercharging the HP compressor, both units being driven by the turbines with a single stage on both, HPT and IPT turbines (the latest Trent XWB engine has two stages on IPT turbine – architecture 1F-8IPC=6HPC〨1HPT=2IPT-6LPT), mounted on separate HP and IP shafts. Consequently to this, the HPC needs only to develop a modest pressure ratio (~4.5:1). On the other side, U.S.´ civil turbofan engines - Pratt & Whitney and General Electric, those aimed for the civil aircraft, use much higher HP compressor pressure ratios (~23:1 on the GEnx-1B - Boeing 787 and GEnx-2B - Boeing 747-8 engines) and are driven by a two-stage HP turbine (GEnx-1B/-2B, GE90-123" fan /128" fan, GE9X and Engine Alliance GP7200 engines do all have 2 stages on HPT). Even so, there are usually a few IP axial stages mounted on the LP shaft, just behind the fan, to further supercharge the core´s compression system. Civil engines have multi-stages LP turbines, the number of stages being determined by the bypass ratio, the amount of IP compression on the LP shaft and the LPT blade´s circumferential speed. And, if the BPR´s ratio increases, the mean radius ratio of the fan and LPT increases. Consequently, if the fan is to rotate at its optimum blade speed, the LPT blading will spin slowly, so additional LPT stages will be required, to extract sufficient energy to drive the fan. That is the reason why the gearbox turbofan engines enter the big door at the civil airliners market. Currently, the most famous is Pratt & Whitney Pure Power PW1000G engine family, aimed for the wide spectar of the civil aircraft: Airbus A320neo aircraft family - A319neo, A320neo and A321neo, Bombardier CSeries - CS100 and CS300, Mitsubishi MRJ - MRJ70E and MRJ90E and Embraer E2family- E175-E2, E190-E2 and E195-E2…)
And while the commercial turbofan engine manufacturers have focused on developing very high BPR (bypass ratio) and OPR (overall pressure ratio) systems, a supersonic (mostly military) engines require a comparatively low BPR and CPR (compressor pressure ratio). Too large fan diameter creates too much frontal area drag, but also the massive volume of cool exhaust flow, and as such is not able to move fast enough to push the aircraft to supersonic speeds.
On the front end of the engine, the FPR (fan pressure ratio) affects the specific thrust (thrust divided by the inlet airflow) and indirectly the speed of the air through the engine. And while the low-bypass engines tend to have very high specific thrust values, those large high-bypass turbofans have a very low specific thrust. Civilian turbofans usually use one large fan whereas high performance military turbofans typically use 3 or more fan stages for this exact reason (Izdeliye 117 engine has four of them while Izdeliye 30 engine has three). Retaining still for a moment on the engine´s propulsion efficiency, the exhaust nozzle has to be mentioned as a highly important part of the whole story. In ideal conditions, a jet engine exhausts the flow at the ambient pressure so it could produce a stable area of thrust. However, a given engine can push the air out at a higher pressure than is the ambient one, but this flow will simply over-expand, collapse-in at its LP core and possibly re-expand. This phenomenon causes inefficiency and could be dangerous to the aircraft´s operation. To allow the higher than the ambient pressure flow to expand under control, so the energy is translated more axially rather than radially, a divergent section of nozzle is required. Each angle designed considering the convergent and divergent sections has a specific Mach number and pressure ratio associated with it. Knowing that fact, it is not hard to conclude that the aircraft would have maximum efficiency across a wide range of Mach numbers with the variable convergent-divergent nozzle at its exhaust. But, such a type of nozzle is very complicated to build and requires a system to activate it, just like hydraulic one or bleed air. A fixed nozzle has much lower efficiency, but also and much lower cost of construction...
Considering the thermal efficiency of the engine; all until the TIT is kept constant, specific thrust expound a maximum in its variation with CPR, because as the compressor pressure ratio is increased, the combustor´s inlet temperature is also increased. This means that the fuel to air ratio must be decreased to avoid overheating the turbine, and if the CPR was large enough, the maximum allowable temperature would be achieved at the compressor outlet and any addition of the fuel would overheat the turbine. In this way, a turbojet engine with high CPR can´t produce thrust at high Mach numbers without exceeding the maximum allowable TIT. No matter of the fact the thermal efficiency is increased with the higher CPR, the attendant decrease in specific thrust, at the higher Mach numbers, makes high-CPR turbojet engines impractical for supersonic flight. The optimum CPR reduces quickly with the increasing Mach number, in supersonic flight. On the other side; for the subsonic flight, high(er) CPR is welcome just to attain better engine´s thermal efficiency and lower specific thrust. However, for the supersonic flights, lower CPRs are typically used to accomplish higher specific thrust …
The development of the material technology, to a large degree, conditions the efficiency of the construction of the modern turbofan engine, and to search for more power and thrust out of the existing or completely new engine´s constructions, implies the existence of the higher thermal and mechanical stresses of the engine´s construction, and all because of the increasing of their operating parameters (pressures, temperatures, mass flow, rotational speeds…). In the same way as the increased mechanical stresses may affect the fractures and tearing-offs of the materials, the increased long-term temperature stresses could easily cause a change in the structure of the materials leading to their disintegration. Besides, to have a thermally efficient engine does not mean that it is, in the same time, a propulsive efficient engine just for every purpose…
No matter of the engine´s limitations, conditioned by its architecture and the core´s geometry, there are still many opportunities for the engine´s improvements: in its aerodynamics, higher rotational speeds, improved combustors, higher working parameters (temperatures, pressures - FPR, CPR and OPR, TIT, mass flow...), stronger cooling of the thermally most loaded engine´s sections, using of the new sophisticated materials… Very interesting and wide area …
… I would say that the Изделие 30 engine has larger diameter than the Saturn AL-41F-1 (изд. 117) engine (fan diameter: 932,0 mm / 36,7 in). I remember I have read somewhere that the Saturn AL-41F (Изделие 20) engine, once designed for the propulsion of MiG 1.44 MFI, had the fan diameter of 1.000,0 mm / 39,4 in. I don´t know if that was correct data and if Izdeliye 30 engine has the same one, but there is no doubt that the temperatures, pressures, N1 and N2 speeds, mass flow of the Изделие 30 engine were heavily increased compared to the Saturn AL-41F-1 (изд. 117) engine. Namely, there are two basic ways to increase core´s power: hot route - increasing HP turbine´s inlet temperature (TIT) and cold route - increasing core´s mass flow. The larger fan´s airflow requires (I guess Izdeliye 30 engine has larger BPR than the Izdeliye 117 engine) more power from the engine´s core. This can be achieved by raising the OPR (combustor inlet pressure / intake delivery pressure) just to induce more airflow into the core and by increasing TIT. Together, these parameters tend to increase core´s thermal efficiency and improve engine´s SFC.
The desired OPR (overall pressure ratio) for the engine´s Brayton cycle is usually achieved by the multiple axial stages on the core´s compression section. Let me explain in just a few words how that functions at civil airliners´engines. F.e. Rolls-Royce tend to split the core´s compression with an IP compressor supercharging the HP compressor, both units being driven by the turbines with a single stage on both, HPT and IPT turbines (the latest Trent XWB engine has two stages on IPT turbine – architecture 1F-8IPC=6HPC〨1HPT=2IPT-6LPT), mounted on separate HP and IP shafts. Consequently to this, the HPC needs only to develop a modest pressure ratio (~4.5:1). On the other side, U.S.´ civil turbofan engines - Pratt & Whitney and General Electric, those aimed for the civil aircraft, use much higher HP compressor pressure ratios (~23:1 on the GEnx-1B - Boeing 787 and GEnx-2B - Boeing 747-8 engines) and are driven by a two-stage HP turbine (GEnx-1B/-2B, GE90-123" fan /128" fan, GE9X and Engine Alliance GP7200 engines do all have 2 stages on HPT). Even so, there are usually a few IP axial stages mounted on the LP shaft, just behind the fan, to further supercharge the core´s compression system. Civil engines have multi-stages LP turbines, the number of stages being determined by the bypass ratio, the amount of IP compression on the LP shaft and the LPT blade´s circumferential speed. And, if the BPR´s ratio increases, the mean radius ratio of the fan and LPT increases. Consequently, if the fan is to rotate at its optimum blade speed, the LPT blading will spin slowly, so additional LPT stages will be required, to extract sufficient energy to drive the fan. That is the reason why the gearbox turbofan engines enter the big door at the civil airliners market. Currently, the most famous is Pratt & Whitney Pure Power PW1000G engine family, aimed for the wide spectar of the civil aircraft: Airbus A320neo aircraft family - A319neo, A320neo and A321neo, Bombardier CSeries - CS100 and CS300, Mitsubishi MRJ - MRJ70E and MRJ90E and Embraer E2family- E175-E2, E190-E2 and E195-E2…)
And while the commercial turbofan engine manufacturers have focused on developing very high BPR (bypass ratio) and OPR (overall pressure ratio) systems, a supersonic (mostly military) engines require a comparatively low BPR and CPR (compressor pressure ratio). Too large fan diameter creates too much frontal area drag, but also the massive volume of cool exhaust flow, and as such is not able to move fast enough to push the aircraft to supersonic speeds.
On the front end of the engine, the FPR (fan pressure ratio) affects the specific thrust (thrust divided by the inlet airflow) and indirectly the speed of the air through the engine. And while the low-bypass engines tend to have very high specific thrust values, those large high-bypass turbofans have a very low specific thrust. Civilian turbofans usually use one large fan whereas high performance military turbofans typically use 3 or more fan stages for this exact reason (Izdeliye 117 engine has four of them while Izdeliye 30 engine has three). Retaining still for a moment on the engine´s propulsion efficiency, the exhaust nozzle has to be mentioned as a highly important part of the whole story. In ideal conditions, a jet engine exhausts the flow at the ambient pressure so it could produce a stable area of thrust. However, a given engine can push the air out at a higher pressure than is the ambient one, but this flow will simply over-expand, collapse-in at its LP core and possibly re-expand. This phenomenon causes inefficiency and could be dangerous to the aircraft´s operation. To allow the higher than the ambient pressure flow to expand under control, so the energy is translated more axially rather than radially, a divergent section of nozzle is required. Each angle designed considering the convergent and divergent sections has a specific Mach number and pressure ratio associated with it. Knowing that fact, it is not hard to conclude that the aircraft would have maximum efficiency across a wide range of Mach numbers with the variable convergent-divergent nozzle at its exhaust. But, such a type of nozzle is very complicated to build and requires a system to activate it, just like hydraulic one or bleed air. A fixed nozzle has much lower efficiency, but also and much lower cost of construction...
Considering the thermal efficiency of the engine; all until the TIT is kept constant, specific thrust expound a maximum in its variation with CPR, because as the compressor pressure ratio is increased, the combustor´s inlet temperature is also increased. This means that the fuel to air ratio must be decreased to avoid overheating the turbine, and if the CPR was large enough, the maximum allowable temperature would be achieved at the compressor outlet and any addition of the fuel would overheat the turbine. In this way, a turbojet engine with high CPR can´t produce thrust at high Mach numbers without exceeding the maximum allowable TIT. No matter of the fact the thermal efficiency is increased with the higher CPR, the attendant decrease in specific thrust, at the higher Mach numbers, makes high-CPR turbojet engines impractical for supersonic flight. The optimum CPR reduces quickly with the increasing Mach number, in supersonic flight. On the other side; for the subsonic flight, high(er) CPR is welcome just to attain better engine´s thermal efficiency and lower specific thrust. However, for the supersonic flights, lower CPRs are typically used to accomplish higher specific thrust …
The development of the material technology, to a large degree, conditions the efficiency of the construction of the modern turbofan engine, and to search for more power and thrust out of the existing or completely new engine´s constructions, implies the existence of the higher thermal and mechanical stresses of the engine´s construction, and all because of the increasing of their operating parameters (pressures, temperatures, mass flow, rotational speeds…). In the same way as the increased mechanical stresses may affect the fractures and tearing-offs of the materials, the increased long-term temperature stresses could easily cause a change in the structure of the materials leading to their disintegration. Besides, to have a thermally efficient engine does not mean that it is, in the same time, a propulsive efficient engine just for every purpose…
No matter of the engine´s limitations, conditioned by its architecture and the core´s geometry, there are still many opportunities for the engine´s improvements: in its aerodynamics, higher rotational speeds, improved combustors, higher working parameters (temperatures, pressures - FPR, CPR and OPR, TIT, mass flow...), stronger cooling of the thermally most loaded engine´s sections, using of the new sophisticated materials… Very interesting and wide area …
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