The intake
The J-20 has a Diverterless Supersonic Inlet, also known as DSI, it has the advantages of simpler manufacturing process, less need for larger amounts of Radar absorbing material than conventional inlets, no need for bleed systems, a diverter cavitiy and mechanical Variation, however all of this limits too the speed to Speeds below Mach 2.
The high pressure Gradient generated by the bump becomes a wall for the lower pressure boundary layer, but by the bump having no mechanical variation it limits the Mach number it is designed to operate.
The Bump can not change its size and position with respect the inlet cowl, thus limiting the speed range it can operate.
Pressure recovery tends to decrease at supersonic speed at all conditions. This is due to the fact that the shock waves at the inlet in supersonic condition causes additional pressure loss and hence it results in lower pressure recovery as compared (Goldsmith and Seddon 1993, Mattingly 2002). This phenomena is quite similar in fixed intakes (Ibrahim, Ng et al. 2011).
The results revealed that Boundary Layer Diverter Intake (BLD intake) configuration is more effective in subsonic regime as compared to DSI configuration, whereas at supersonic speeds DSI configurations gave superior performance.
Comparative Flow Field Analysis of Boundary Layer Diverter Intake and Diverterless Supersonic Intake Configuration I. Arif† , S. Salamat, M. Ahmed, F. Qureshi and S. Shah
System-level trade studies were performed to quantify the weight, cost, and benefits of the DSI, compared to more conventional inlets (e.g., F-22 and F/A-18E/F caret inlet systems). In these studies, a 30-percent inlet weight reduction was estimated for the DSI, relative to the reference caret inlet. The largest contributing factor was the elimination of the bleed and bypass systems. Studies performed by other ACIS contractors [25] indicated similar savings for diverter-less/bleed-less systems.
https://www.lockheedmartin.com/content/dam/lockheed-martin/eo/documents/webt/F-35_Air_Vehicle_Technology_Overview.pdf
The Diverterless Supersonic Inlet was first designed for an F-16 variant and first tested in December 1996, China copied the design to apply it to the J-10, JF-17 and J-20.
By the mid 1950s intake research in the USA included the basic concept for the DSI intake, it is Antonio Ferri's compression bump with forward swept scoop or "Ferri"intake.
It is defined as:
An object of this invention is to provide a novel construction of a scoop-type supersonic air inlet for the efficient conversion of dynamic pressure to static pressure. A further object of this invention is to provide a novel construction to facilitate starting of a Scoop-type Supersonic air inlet. Increases in static pressure, or compression, in a supersonic inlet can be accomplished by the physical turning of the air which results in the generation of compressive shock waves. A further object of the invention comprises the provision of a scoop-type air inlet having a novel aerodynamic compression surface (hereinafter termed a compression bump) for more efficiently producing compression in the inlet. In accordance with the invention, the compression bump is formed on the inlet surface adjacent to the aircraft body member from which the inlet projects so that only a part of the physical turning and resulting air compression required is produced by the inlet outer wall formed by the scoop, the rest of the required physical turning and resulting air compression being effected by the bump surface. The compression bump preferably is integral with the surface of said aircraft body member and its upstream end may be disposed upstream or downstream of the leading edge of said outer wall of the inlet. With the leading edge of said bump disposed upstream of the leading edge of said inlet outer wall, the presence of the bump also provides precompression of the air before the air enters the inlet thereby reducing the Mach number of the supersonic airstream before the air enters the inlet so that less air spillage is required to start the inlet. In addition, in the case of a three-dimensional bump, the bump imparts to the air flow a pressure gradient which is normal to the direction of the entering air thereby assisting in removal of the boundary layer of air adjacent to said body member and bump.
US2990142A - Scoop-type supersonic inlet with precompression surface - Google Patents
It is the position of the bump relative to the intake that is the major difference and this shows how important the positioning is. It indicates that it is advantageous to place the maximum amplitude of the bump close to the cowl lips of the intake, so that they coincide with the shock from the bump surface.
A comparison between Intake & Mod 1 and Intake & Mod 2 show that high a amplitude of the bump is preferable to a low amplitude. This gives both higher pressure recovery as well as better boundary layer diversion
http://www.diva-portal.org/smash/get/diva2:221/FULLTEXT01.pdf
Extensive experiments were conducted on a body-integrated diverterless supersonic inlet (DSI). Diverterless supersonic inlets are designed and developed in order to provide both supersonic flow compression and boundary-layer diversion by using a three-dimensional bump in combination with a suitable cowl lip. The present experiments were performed at three different freestream Mach numbers of M∞=0.75M∞=0.75, 1.65 (the design Mach number), and 1.85, as well as at 0 deg angles of attack and angles of sideslip. To model the performance accurately, the intake was integrated with a typical forebody including a nose with an elliptical cross section. Wind-tunnel tests were conducted at critical, subcritical, and supercritical operating conditions. The results showed that the present DSI has acceptable performance in these operating conditions and is able to provide the required mass flow and static pressure ratios. For all conditions examined in this study, as a significant result, the fixed geometry of the designed DSI showed acceptable performance in the ranges of supersonic Mach numbers tested: M∞=1.65–1.85M∞=1.65–1.85; furthermore, its operation in the subsonic condition of M∞=0.75M∞=0.75 was satisfactory. It should be mentioned that there were no movable parts or an auxiliary flow control system for this intake
https://arc.aiaa.org/doi/abs/10.2514/1.C035328
Therefore, at supersonic speed higher amplitude of bump is preferred over smaller amplitude. In case of Config 1 bump, shock on lip phenomenon is met since its maximum amplitude is kept near the cowl lip. Pressure above the intake duct is almost same in all the cases since intake duct is same for all the cases so.
https://www.eares.org/siteadmin/upload/8484EAP5171002.pdf
I. Introduction The inlet is a duct before the engine. Its basic function is to capture a certain amount of air from the freestream and supply it to the engine. Most gas turbine engines require the Mach number at the engine face at a moderate subsonic speed, to be about Mach 0.4. Therefore, for supersonic aircraft with a gas turbine engine, the inlet will reduce the supersonic freestream to subsonic speed, and provide a matched air mass flow rate to the engine. The gas turbine engine requires a supply of uniform high total pressure recovery air for good performance and operation, thus the quality of the airflow at the engine face will significantly affect the performance of the engine, especially the total pressure loss which affects the engine thrust and consequently the fuel consumption. For 1% total pressure loss, the engine will suffer at least 1% thrust loss. Therefore, it is important to maximize the total pressure recovery at the engine face. The total pressure recovery is the ratio of the total pressure of the airflow at the engine face to that of the freestream.
http://citeseerx.ist.psu.edu/viewdoc/download?doi=10.1.1.559.484&rep=rep1&type=pdf
An aircraft inlet captures freestream air and reduces its velocity so the engine can process if in a stable and efficient manner. In order to minimize compressor work, inlet diffusion should be accomplished with a minimum of total pressure loss. The inlet should also deliver the working fluid with minimum distortion, all over a wide range of Mach number, angle-of-attack, angle-of-sideslip, and engine demand. The supersonic inlet for a tactical aircraft must also be sized to provide a maximum demand airflow which usually occurs at maneuver or acceleration -.- conditions. When the aircraft is at a subsonic cruise condition, however, the engine needs to process only a limited mass flow associated with 40-60 percent maximum dry thrust. The inlet, however, is still capable of processing larger mass flow closer to maximum demand.
The control of the shock wave position and prevention of shock induced flow separation in the inlet can be accomplished by bleeding boundary air from the inlet ramps, cowls, or sidewalls and dumping that flow overboard. This produces forces similar to the bypass flow which must be considered in supersonic inlet throttle dependent forces.
https://apps.dtic.mil/dtic/tr/fulltext/u2/a162939.pdf
A Ventral Diverterless High Offset S-shaped Inlet at Transonic Speeds Xie Wenzhong*, Guo Rongwei College of Energy and Power Engineering, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, China Received 13 September 2007; accepted 18 December 2007
The advent of supersonic aircraft powered by airbreathing engines opened up a new set of challenges for intake designers. A rule of thumb often used is that 1% pressure loss reduces thrust by 1%, but it became clear early on that the thrust loss caused by pressure losses in supersonic flight increases nonlinearly. For example, at a flight speed of Mach 2.2, a typical engine losing 8% of the freestream total pressure through the intake will suffer a reduction in thrust of 13% and a 5% increase in fuel consumption [9]. Since the mid-1950s, when this first became evident, a tremendous amount of research effort has gone into the study of supersonic intake pressure recovery and drag
Boundary-layer bleed in supersonic inlets is typically used to avoid boundary layer flow separation f_m adverse shock-wave/boundary-layer interactions and subsequent total pressure loss in the subsonic diffuser and to stabilize the normal shock
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19950013353.pdf
Turbojet installed thrust 6 • Uninstalled thrust is obtained from engine manufacturer, preliminary cycle analysis or a fudge factor approach. • Every 10 years: 25% less SFC, 30% less weight, 30% less length, • Installed thrust = uninstalled thrust – installation effects – drag contribution assigned to the propulsive system
http://www.ae.metu.edu.tr/~ae452/lecture1_propulsion.pdf
The present work is an extension of our earlier computational study [24] on an integrated Diverterless Supersonic Inlet (DSI) where complete external and internal (intake duct) flow features along with intake performance characteristics were presented. In this paper the computed flow and performance characteristics at low angle-of-attack (AOA) of a passive bleed system on the same DSI are presented near intake design mass flow rate. The passive bleed system comprises of a porous wall segment on the bump compression surface that is connected to a bleed chamber with a single outlet exposed to a specific pressure. The bleed mass flow rate is controlled through controlling the bleed chamber outlet pressure. The subsonic characteristics of this passive bleed system on the DSI are evaluated at M∞=0.8 while the supersonic characteristics are evaluated at M∞=1.7, which is near the design Mach number for the intake. The results of the present study indicates that although some low energy bleed chamber air seeps into the intake duct at M∞=0.8, however the passive bleed system on the whole results in some improvement in engine face total pressure recovery due to removal of part of bump compression surface residual boundary layer. At supersonic conditions (M∞=1.7) the boundary layer diversion mechanism behind the shock wave [24] prevents the low energy bleed chamber seepage air, along with upstream boundary layer from entering the intake duct. However, the passive bleed system results in significant deterioration of engine face total pressure recovery due to its affect on the inlet shock wave structure and related flow field.
Effect of Passive Bleed System on an Integrated Diverterless Supersonic Inlet
https://www.researchgate.net/publication/268565958_Effect_of_Passive_Bleed_System_on_an_Integrated_Diverterless_Supersonic_Inlet
An advanced diverterless supersonic inlet (bump) design and the wind test of the thunder/JF-17 aircraft are analyzed. The bump inlet has excellent aerodynamic performances: high total pressure recovery(σ is 0.02～0.04 higher than traditional fixed geometry inlet), low distortion and good compatibility. The Bump inlet eliminates many traditional complex features associated with boundary layer management: the diverter, bleed system and bypass system. By eliminating these features, the inlet aperture is integrated with the aircraft forebody, thus producing the best value configuration with low drag, low weight and high reliability
Design of Bump Inlet of Thunder/JF-17 Aircraft
Yang Yingkai 1,2 (1.College of Energy and Power Engineering, Nanjing University of Aeronautics & Astronautics, Nanjing, 210016, China; 2.Chengdu Aircraft Design and Research Institute, Chengdu,610041 ,China)
The requirements for the inlet: – A high pressure recovery (1% loss in inlet pressure recovery results in 1.3% loss in thrust).
• DSIs also crucially improve the aircraft's low-observable characteristics (by eliminating radar reflections between the diverter and the aircraft's skin). • Additionally, the "bump" surface reduces the engine's exposure to radar, significantly reducing a strong source of radar reflection because they provide an additional shielding of engine fans against radar waves. • However, a diverterless intake reacts considerably to oblique flow, which is a disadvantage in maneuvering flight especially for two-engine aircraft.
http://ae.metu.edu.tr/~ae452sc2/lecture1_intakes_nozzles.pdf
CONCLUSION
The DSI on J-20 is a fixed type intake, as such will limit the J-20 at a Max speed of Mach 1.6 to Mach 1.8, with an ideal pressure recovery at Mach 1.0, the intake potentially could operate at slightly higher speeds, perhaps Mach 2, however loses in pressure recovery will decrease the max yield of the engine, thus thrust will be reduced and fuel consumption increased.
Not even with WS-15 the aircraft will surpass fighters like Su-57 or F-15 at speeds higher than Mach 2.0 in terms of acceleration and thrust to weight ratio because these aircraft that use variable geometry intakes that can achieve much higher pressure recovery at Mach 2.0, thus J-20 will remain in the max speed range of J-10C or F-35.
So the Dragon will not surpass the Penguin in terms of Max speed, but perhaps the J-20 could have better acceleration than F-35.
There are studies for DSI intakes for speeds up to Mach 2.5 but do not correspond to the one seen on J-20
Design of the Bump The generic design of intake described here is a typical 2D wedge intake where the external bleed diverter has been replaced by a conical shock bump. It is designed to operate at M=2.5. This differs from the concept used for low supersonic aircraft, where the bump is intended to replace both diverter and the ramp.
The air intake works on-design if the conical shock issued from the bump intersects the middle of the cowl, that is:
tan β = hc+hp/LB+LC
The bump shape is designed to follow the streamlines created by the desired conical shock. Corresponding solutions are not unique since several bumps can generate the same conical shock. In order to obtain a solution, a variable has to be chosen (length or width).
http://www.garteur.org/Technical%20Reports/AD_AG-34_%20Phase%201_TP-129_OPEN.pdf
In 1971-72 Vought designed a lightweight fighter with Ferri`s intake, this was the V-1100; General Dynamics studied a Ferri`s intake too for their Model 401, however they decided to use a fixed intake with a splitter plate instead for their F-16. Both companies considered for for the LWF competition the Ferri`s intake due to its light weight and decent performance.
