In a previous report, scramjet engine characteristics of different propellant-fed cycles were compared and engine performances were discussed. In this study, the cooling requirement of a hydrogen-fueled airframe-integrated scramjet engine as well as an airframe was examined, and effects of various parameters including flight Mach number, flight dynamic pressure, engine wall temperature, and engine scale, on the engine characteristics were analyzed. The coolant required for the airframe was about 20% of the total coolant. Simple equations that correlate coolant flow rate with those parameters were derived. A B b CD Cp 7sp t M m P Q q

Parametric study of airframe-integrated scramjet cooling requirement

Rocket-based combined-cycle (RBCC) engines are known as the most promising type of engine with the potential to realize ‘single stage to orbit’ (SSTO). This will dramatically reduce the cost of space round trips and improve the physical experience of astronauts. However, sixty years after this concept was first proposed, no RBCC engine has completed a real flight. The challenges in RBCC development include ejector, scramjet, mode transition, and thermal protection technologies. However, great progress has been made in recent years, suggesting a bright future for these engines. In this paper, we review worldwide progress in the overall layout of RBCC engines. The working process of RBCC engines is introduced to show their distinctiveness among traditional engines. RBCC engines are classified as rectangular section or axisymmetric configuration engines and the development of both types in different countries is reviewed. The engine-airframe integration design and mission planning of RBCC powered aircraft systems are analyzed separately. Even though RBCC powered aircraft and their missions remain conceptual, the design and planning processes are important for RBCC development and space round trips in the future. RBCC study is a typical multi-disciplinary design process. Research addressing the problems encountered by RBCC studies will promote the development of a range of disciplines relevant to aerospace science.

The overall layout of rocket-based combined-cycle engines: a review

A rocket-ramjet combined cycle engine model, embedding twin rocket chamber on top wall side of a scramjet flow-pass, was tested in its ejector-mode operation under sea-level static conditions. The rocket chamber was driven with gaseous hydrogen and oxygen at nominal operation condition of 3 MPa in chamber pressure and 6.5~7.5 in mixture ratio. Gaseous hydrogen was also injected through secondary injector orifices to pressurize the ramjet combustor. Mixing between the hot rocket plume and cold airflow as well as combustion of residual fuel within the plume with the airflow caused entropy and static pressure increases in the constant-area mixing duct in our original flow-pass design, resulting in a high back-pressure to the incoming airflow and a limited airflow rate. Thus, the mixing duct was re-designed to have divergence from its onset to compensate the pressure-rise. With this modified flow-pass configuration, the airflow rate was increased by 40%. However, this flow-pass geometry resulted in generation of low speed area, through which pressure-rise due to secondary combustion (and flow-pass exit contraction to simulate secondary combustion) penetrate the mixing duct and reduced the incoming airflow rate. A contraction on the incoming airflow enabled choking condition of the incoming airflow to sustaining the airflow rate, while the rate itself was reduced. Contraction at the exit of the engine enhanced mixing, however, choking condition was not attained due to the higher pressure level associated with the high exit contraction. Balancing these factors and additional mixing enhancement are required. The engine performances were summarized.

Performance of a Rocket-Ramjet Combined-Cycle Engine Model in Ejector Mode Operation

An expander cycle for an airframe-integrated hydrogen-fueled scramjet is analyzed to study regenerative cooling characteristics and overall specific impulse. Below Mach 10, the specific impulse and thrust coincide with the reference values. At Mach numbers above 10, a reduction of the specific impulse occurs due to the coolant flow rate requirement, which is accompanied by an increase of thrust. It is shown that the thrust may be increased by injecting excess fuel into the combustor to compensate for the decrease of the specific impulse. 9 refs.

Propellant feed system of a regeneratively cooled scramjet

A combustion-capable combined cycle engine model which was constructed based on the rocket and ramjet technology was tested in Mach 4 flight condition. At this speed, engine is designed to shift its operation mode from an ejector-jet to a ramjet. Both modes were simulated by changing the rocket combustion pressure. Even with full rocket exhaust, no effect to the air flow could be observed. The injection point of the secondary fuel affected thrust performance. In the ramjet mode, the pressure rise due to the fuel combustion traveled to the entrance of the combustor, but it stayed near the injection point in the ejector-jet mode.

Ram and Ejector-Jet Modes Experiments of the Combined Cycle Engine in Mach 4 Flight Conditions

Design of Sub-Scale Rocket-Ramjet Combined Cycle Engine Model

The expander cycle or the coolant bleed cycle LOX/LH2 engine is considered suitable for use with the orbit transfer vehicle and the upper stage of a conventional rocket because of their simplicity of engine system. However, these engine cycles generally require a thrust chamber with a high expansion area ratio in order to obtain a high performance and necessarily require a High Altitude Test Facility even in the initial phase of development. Here, the authors propose a new engine cycle called the "Partial Bleed Expander(PBE) cycle". In the PBE cycle, only the turbopump of one propellant(usually LH2) is driven in the expander cycle and a portion of fuel is bled to drive the turbopump of the of the other propellant (usually LOX). The bled fuel may be utilized for dump cooling or film cooling. From the calculation of the PBE cycle, it was made clear that the PBE cycle can perform at a level close to that of the expander cycle with the bleeding rate resulting in minimum Isp loss.

Partial bleed expander cycle for low thrust lox/lh2 rocket engine

The high frequency instability of a JP-4 fuel and liquid oxygen rocket combustor, especially the effects of this instability on heat transfer augmentation to the chamber wall, was studied. The FOF triplet type injector element and two chambers with different lengths were used in the experiments. Both chambers delivered a 10kN thrust at a 3.5MPa chamber pressure. The pressure oscillation generated spontaneously was identified as being a longitudinal first mode. The ratio of rms pressure amplitude to average chamber pressure observed in unstable combustion was 0.23 at most. In order to obtain stable combustion or weak unstable combustion, a damping device based on Crocco’s theory of longitudinal high frequency instability was designed and used successfully. The heat flux to the wall increased six to seven times at most due to the instability. It was shown that the increase of the bulk heat transfer coefficient was proportional to the square root of rms pressure amplitude when the effective temperature, based on the frequency of oscillation, was used. Various considerations, including the effect of solid carbon particles, suggests that Dent’s model of convective heat transfer augmentation is more reasonably consistent with the experimental results than is Hanby’s model. The time average static pressure along the chamber axis increased downstream after it had first decreased. Theoretical analysis of the periodically oscillating one-dimensional flow showed reasonable agreement with the experimental results.

High Frequency Instability in a Lox-Kerosene Rocket Combustor (Effects on Heat Transfer Augmentation)

The injection and combustion characteristics of unlike-impinging triplet injector for JP4 fuel and liquid-oxygen rocket combustor were studied. The injector element types considered were the OFO type, which consists of two outer oxidizer jets and one central fuel jet, and the FOF type, which consists of two fuel jets and one oxidizer jet. The spray characteristics of typical single elements of the two types were investigated through cold flow tests. Based on these tests, three multiple-element injectors (20 and 40 elements, 10 kN thrust at 2 MPa chamber pressure) were designed and fabricated, and were evaluated by means of cold flow tests and firing tests with combustion chambers of various lengths. The results were compared with a theoretical analysis based on a vaporization model. The conclusions of the study are as follows: 1)The design correlation of orifice area ratio for maximum mixing efficiency was confirmed for both OFO and FOF type elements. 2)The OFO element showed wider spray angles and relatively more uniform mixture ratio distributions than the FOF element. The spray characteristics of the single elements were sensitive to a change in mixture ratio. However, multiple element injectors were insensitive to such a change and showed substantially similar mixing measures, suggesting a marked averaging effect due to mixing with adjacent elements. 3)Since the statistical frequency of mixture ratio for multiple element injectors is approximately expressed by the logarithmic normal distribution function, the measure of mixing can be evaluated in terms of the logarithmic standard deviation or the dispersion, and Rupe’s mixing factor, Em, can be correlated with the mixing C efficiency. 4)Initial droplet diameters of the injectors at several mixture ratios were estimated from the comparison of the experimental performance with the theoretical analysis. The estimated diameters indicate that the vaporization of fuel is the rate controlling process. 5)OFO type injectors showed higher performances than FOF types; however, the former gave higher heat flux to the chamber wall. 6)Although direct ignition by a spark plug was used in the study, an injection of a small amount of nitrogen gas into the fuel line was remarkably effective in avoiding hard starts.

義男若松

Papers

Design of Sub-Scale Rocket-Ramjet Combined Cycle Engine Model

Partial bleed expander cycle for low thrust lox/lh2 rocket engine

High Frequency Instability in a Lox-Kerosene Rocket Combustor (Effects on Heat Transfer Augmentation)

Injection and Combustion Characteristics of Unlike-Impinging Triplet Injector for LOX-Kerosene Rocket Combustor

義男 若松

Papers

Design of Sub-Scale Rocket-Ramjet Combined Cycle Engine Model

Partial bleed expander cycle for low thrust lox/lh2 rocket engine

High Frequency Instability in a Lox-Kerosene Rocket Combustor (Effects on Heat Transfer Augmentation)

Injection and Combustion Characteristics of Unlike-Impinging Triplet Injector for LOX-Kerosene Rocket Combustor

義男若松