Gases trapped in the propellant feed lines of space-based rocket engines due to cryogenic propellant boil-off or pressurant ingestion can result in poor combustion efficiencies, combustion instabilities, or long startup transients. To assist NASA in the use of the high performing liquid oxygen propellant combinations in space engines, IN Space proposes to investigate the feasibility of an innovative swirl injector design for liquid oxygen and hydrocarbon propellants to achieve high combustion efficiencies, stable operation, and short and smooth startup transients despite potential two-phase oxidizer flow. Additionally anticipated benefits of the injector include low inert mass and low manufacturing costs. IN Space plans to carry out the feasibility assessment of the injector design by conducting broad parametric test fire evaluations of a notional LOX/hydrocarbon workhorse thruster based on present NASA needs to assess the effects of several design considerations on the combustion efficiency, static combustion stability, and startup transient duration performance merits. A preliminary flightweight injector design will also be generated in order to compare the estimated injector mass with similar injector designs.