It is interesting comparing the two best known first stages in the US that use kerosene and LOX. The Atlas 5 and the Falcon 9 use a similar fuel in their first stages and then diverge in the technical aspects. The Atlas 5 with the RD 180 engine has about 10% higher Isp at sea level while the Falcon 9 Merlin has nearly twice the thrust/weight ratio. The over all Falcon 9 first stage seems to have a much lower dry mass ratio which makes up the difference in engine performance and then some.
There are going to be new vehicles designed by the various companies eventually that would like to benefit from the competitive advantages of both vehicles. A high thrust/weight ratio engine with high Isp that also has low dry mass is a desirable target. The more these features can be designed in, the more mass is available for payload, reusability, or both.
One of the engine cycles that is discussed from time to time is the dual chamber concept. It is more or less a gas generator cycle with an exhaust pressure high enough to inject into a lower pressure thrust chamber to burn with fuel or oxidizer to get useful thrust. I suggest it might be possible to get very near RD 180 Isp with very near Merlin thrust/weight with a variation of the concept. A low stage dry mass being part of the goal, I add in a few features that may be unique.
The black boxes in the tanks are the electric inducer pumps from the previous post. They are to keep the propellants at high enough pressure to the main pumps to suppress cavitation as well as keeping required tank pressurization to a minimum.
The small blue tanks in the inter tank area are for the liquid hydrogen that serves multiple purposes. First the hydrogen feed hits a heat exchanger in the LOX tank to keep it cold enough to stay liquid and suppress cavitation even as tank pressure drops. Then it hits a heat exchanger in the RP tank for the same purpose. Then it is used to cool the turbine blades the same way that jet engines use air cooling. Finally it burns with the excess LOX from the gas generator to produce thrust.
With the pumps providing pressures to the main engine similar to that of the RD 180, the Isp of them should be similar. About 10% of the propellant goes to the gas generator driving the pumps with a residual pressure of 300 psi after the turbine. If the 300 psi engine was a normal kerosene engine, one would expect an Isp in the 250s from that portion of the thrust system. With the lean (LOX rich) gas generator driving a hydrogen cooled turbine at much higher than normal turbine inlet temperatures, the warm hydrogen mixes with the hot oxygen as it is used for film cooling of the blades and burns in the secondary chamber above the throat. The hydrogen/kerosene/LOX engine at 300 psi could approach the ISP of the main engines due to the higher performance of hydrogen. Hydrogen usage will be a fraction of a percent of the total propellant load.
The compensating nozzle of the low pressure engine in the center would allow reasonable Isp of that portion at sea level, especially with the hydrogen component. The higher expansion ratio made possible would allow much higher Isp at altitude, which, with the hydrogen component, could give vacuum Isp higher than the RD 180. I think the potential result is low hardware mass combined with high first stage performance.