Daniel J. H. Levack

Cycles for Earth-to-Orbit Propulsion

The reduction of Earth-to-orbit launch costs in conjunction with an increase in launcher reliability and operational efe ciency are the key demands on future space transportation systems. Results of various system analyses indicate that these demands can be met with future single-stage-to-orbit (SSTO) vehicles using advanced technologies for both structure and propulsion systems. This paper will provide a classie cation and description of all turbopump feed liquid rocket engine cycles, followed by a combined vehicle/propulsion analysis of the design parameters for the propulsion systems and their associated thermodynamic cycles for the launch of future SSTO vehicles. Existing and projected rocket engine cycles capable of SSTO missions will be presented.

Propulsion System Choices and Their Implications

In defining a space vehicle architecture, the propulsion system and related subsystem choices will have a major influence on achieving the goals and objectives desired. There are many alternatives and the choices made must produce a system that meets the performance requirements, but at the same time also provide the greatest opportunity of reaching all of the required objectives. Recognizing the above, the SPST Functional Requirements subteam has drawn on the knowledge, expertise, and experience of its members to develop insight that will effectively aid the architectural concept developer in making the appropriate choices consistent with the architecture goals. This data not only identifies many selected choices, but also, more importantly, presents the collective assessment of this subteam on the “pros” and the “cons” of these choices. The propulsion system choices with their pros and cons are presented in five major groups. A. System Integration Approach. Focused on the requirement for safety, reliability, dependability, maintainability, and low cost.

Tripropellant engine option comparison for SSTO

Recent evaluations of main engine options for the SSTO/RLV mission have considered the use of a tripropellant fueled engine to take advantage of its improved fuel density. A wide range of options have been addressed including configuration, technology level, and design practices. A study was conducted to evaluate engine configurations on a consistent basis of technology level, design practice, and design groundrules. Engine weight and performance of several cycles and cycle variants were determined and then compared on a vehicle dry weight basis to determine the merit of each. A series of bipropellant configurations were also studied under identical groundrules to evaluate the inherent differences due to propellant selection only. The results showed that the three major variants considered, a single-chamber tripropellant, an annular bell tripropellant and a bipropellant, were nearly identical in terms of overall vehicle performance. W

Determining Mars Mission NTP Thrust Size and Architecture Impact for Mission Options

Future exploration missions across the Solar System need technologies that reduce the time of flight, provide efficient payload capability, and reduce the size and the number of launch systems in order to reduce mission risk and cost. NTP (Nuclear Thermal Propulsion) is the proven, high Technology Readiness Level technology, which provides the performance to enable rapid transit and can minimize the number of SLS launches due to the higher ISP (specific impulse). The future of human Mars exploration will see substantial benefit in terms of lower mission mass and faster trip times when NTP is employed. NTP has been proven scientifically and many of the engineering challenges have been addressed in past ground testing of reactor cores for a range of power levels needed for NTP systems. The challenge today is to create an affordable, highly capable in-space propulsion system. Aerojet Rocketdyne (AR) believes that this could be achieved based on using smaller reactors in the NTP designs (e.g., < 500 MWt) that spring-board off the knowledge gained from past research and development and applies new technologies to improve the life and provide eventually reusability. Also, mission architectures that have the local planetary exploration elements pre-deployed ahead of the human crew can have a significant impact on the design of the human NTP spacecraft, NTP power level (thrust size) and eventual NTP system reusability. The desire would be to have the human crewed vehicle as small as technically feasible, optimize the thrust size and optimize the number of engines based on mission need. This could drive the NTP design to have a smaller reactor, present a more affordable development plan, and lower cost operational footprint for future human Mars exploration transportation systems. Approaches to use a small NTP reactor core could provide a development cost benefit with less uranium content and be more easily tested with a smaller facility foot-print. The smaller facility and lower exhaust flow rate provides for less effluent to clean and manage, which, in turn, reduces the development cost due to environmental safety and nuclear material security concerns. Fundamentally a lower power NTP reactor core (< 500 MWt) can reduce the development, procurement, and operational costs making it a more affordable NTP system for a nuclear cryogenic propulsion stage. AR has been working on several NTP system designs that have a wide range of thrust (core) sizes and the scalability for any exploration mission. AR has performed various architecture and design studies from 2011 through 2015 that have identified NTP approaches for using the capability of smaller size NTP systems for robotic and human Solar System missions. AR has used our multidisciplinary design and architecture analysis capability to analyze a split cargo and crew approach where low to medium power (i.e., 100-150 KWe) Solar Electric Propulsion (SEP) pre-position mission cargo (e.g., long duration habitats, transfer stages, and lander systems) and the crew vehicle uses NTP for rapid transfer to Mars. This paper will focus on results of the study for the NTP stage Mars mission architecture element and the examinations of the thrust size and number of NTP engine systems and how they impact a human Mars mission. 1 Fellow, Mission Architecture, PO Box 109680, M/S 712-67, and AIAA Associate Fellow. 2 Program Manager, Advanced Programs, P.O. Box 7922 / MS RFB19, AIAA Member. 3 Sr. Engineer, Mission Architecture, 555 Discovery Dr.

Advanced Low-Cost O2/H2 Engines for the SSTO Application

The recent NASA Access to Space study examined future Earth to orbit (ETO) transportation needs and fleets out to 2030. The baseline in the option 3 assessment was a single stage to orbit (SSTO) vehicle. A study was conducted to assess the use of new advanced low cost O2/H2 engines for this SSTO application. The study defined baseline configurations and ground rules and defined six engine cycles to explore engine performance. The cycles included an open cycle, and a series of closed cycles with varying abilities to extract energy from the propellants to power he turbomachinery. The cycles thus varied in the maximum chamber pressure they could reach and in their weights at any given chamber pressure. The weight of each cycle was calculated for two technology levels versus chamber pressure up to the power limit of the cycle. The performance in the SSTO mission was then modeled using the resulting engine weights and specific impulse performance using the Access to Space option 3 vehicle. The results showed that new O2/H2 engines are viable and competitive candidates for the SSTO application using chamber pressures of 4,000 psi.

Solar Electric Propulsion Architecture for Mars Cargo for Affordable Exploration and Sustained Permanence

Future exploration missions throughout the Solar System will require high efficiency propulsion technologies using moderate spacecraft power that are extensible to multiple mission areas and customers. Solar Electric Propulsion (SEP) has been proven at various power levels and propulsion unit design types and should be part of such a technology portfolio. Many of the SEP thrusters used for recent missions have established the high Technology Readiness Level for this technology. SEP’s high specific impulse (ISP) has a dramatic impact on reducing the required spacecraft propellant while increasing launch window flexibility, making the propulsion technology an ideal choice for delivering large unmanned cargo payloads for sustaining human Mars exploration. The future of human Mars exploration will see substantial benefit in terms of lower mission mass when SEP systems are employed in the architecture. It must be remembered that key design parameters (e.g., power, ISP, specific mass) need to be optimized in order to effectively deliver the highest payload mass possible within a given program. Architectures that have the local planetary exploration elements pre-deployed ahead of the human crew can have a significant impact on the design of the human exploration mission in terms of total mass available at Mars, the size of the crew spacecraft, and the number of total systems employed to create low risk transportation. This paper will discuss a recent Aerojet Rocketdyne (AR) study that examines the impact of SEP power levels, ISP, and stage size required for large (20+mt) cargo prepositioning missions to Mars orbit. The study also examines the impact on the crew vehicle size when different combinations of architecture elements are pre-positioned. The AR analysis has focused on finding the combination of right size launch capability and SEP vehicle power level that delivers large cargo but maintains reasonable trip times for pre-positioning. For near-term missions, we show that 50kWe SEP modules can be used either singly or in combination, enabling a path to lower cost human exploration missions. An approach that uses 40-50kWe SEP modules creates mission extensibility where these modules can also be used for cis-lunar missions and large science missions such as the Asteroid Redirect mission and geosynchronous Earth orbit commercial and military satellite delivery missions launched on existing expendable launch vehicles ensuring that the SEP modules have multiple applications. We show that this modular approach enables a dramatic reduction in total development and production costs for high power SEP and provides for a more gradual phasing of system development to fit within available budgets. 1 Fellow, Space Systems and Mission Analysis, PO Box 109680, M/S 712-67, AIAA Associate Fellow. 2 Executive Director, Space, 1300 Wilson Blvd., AIAA Associate Fellow. 3 Sr. Engineer, Mission Architecture, 555 Discovery Dr., AIAA Member. 4 Sr. Manager, Advanced Space and Launch, P.O. Box 7922 / MS RFB19, AIAA Member. 5 Executive Director, Advanced Space and Launch, 11411 139 Place, AIAA Fellow.

Low Cost Small LOX/HC Launch Vehicle Enabled by Affordable Propulsion

The commercialization of space depends on having affordable access to space. This paper presents an affordable launch vehicle for delivering a small payload to low earth orbit, specifically a launch vehicle to place up to 1,000 lbm in low earth orbit (100 NM, circular orbit, due east, 28.5°). The enabling element for the launch vehicle is affordable propulsion, both in development and in production.

Commercial and Civil In-Space Applications of the Peacekeeper Stage IV (RS-34)

The Peacekeeper Stage IV propulsion and structure hardware will enable low-risk, cost effective in-space stages to provide the basic spacecraft bus propulsion and structure capability to support rapid hardware development schedules for commercial and civil in-space applications. The Peacekeeper was a four-stage vehicle consisting of three solid rocket motors and the liquid-fueled Stage IV. The fourth stage utilized the Stage IV propulsion, guidance and control systems to deploy vehicles in sequence with very precise state vectors and attitudes. Stage IV used Nitrogentetroxide and Monomethylhydrazine storable hypergolic propellants. The system is a simple, reliable, regulated pressure-fed system, with a high pressure helium tank, two propellant tanks with integrated propellant management devices, one axial thruster for orbit transfer maneuvers, and eight vernier thrusters for precise attitude control of the stage for inorbit operations. The propellant tanks were developed with surface tension propellant management devices to enable infinite operational and maneuvering flexibility. A total of 141 Stage IV propulsion systems were integrated, tested and delivered by Aerojet Rocketdyne to support development, qualification, fielding and sustainment operations. The Peacekeeper missile system was decommissioned and the program closed-out in 2006. The remaining stages were separated from their missiles and placed in controlled storage at Hill Air Force Base. About twenty-five stages are currently available. Aerojet Rocketdyne was the prime contractor responsible for development, production, and support of the Peacekeeper Stage IV. The Peacekeeper Stage IV is also referred to as the RS-34. Aerojet Rocketdyne recently successfully concluded support of the Ares 1-X flight test program where two Peacekeeper Stage IV propulsion systems were reconfigured into the Roll Control System (RoCS) modules for the successful Ares 1-X flight test vehicle for NASA MSFC. This successful program demonstrates the feasibility of reconfiguring Stage IV propulsion systems for civil space applications. This paper presents study results showing the capabilities of the Peacekeeper Stage IV for the orbital debris removal, resupply to the ISS, and general space bus applications.

Using Technical Performance Measures

All programs have requirements. For these requirements to be met, there must be a means of measurement. A Technical Performance Measure (TPM) is defined to produce a measured quantity that can be compared to the requirement. In practice, the TPM is often expressed as a maximum or minimum and a goal. Example TPMs for a rocket program are: vacuum or sea level specific impulse (lsp), weight, reliability (often expressed as a failure rate), schedule, operability (turn-around time), design and development cost, production cost, and operating cost. Program status is evaluated by comparing the TPMs against specified values of the requirements. During the program many design decisions are made and most of them affect some or all of the TPMs. Often, the same design decision changes some TPMs favorably while affecting other TPMs unfavorably. The problem then becomes how to compare the effects of a design decision on different TPMs. How much failure rate is one second of specific impulse worth? How many days of schedule is one pound of weight worth? In other words, how to compare dissimilar quantities in order to trade and manage the TPMs to meet all requirements. One method that has been used successfully and has a mathematical basis is Utility Analysis. Utility Analysis enables quantitative comparison among dissimilar attributes. It uses a mathematical model that maps decision maker preferences over the tradeable range of each attribute. It is capable of modeling both independent and dependent attributes. Utility Analysis is well supported in the literature on Decision Theory. It has been used at Pratt & Whitney Rocketdyne for internal programs and for contracted work such as the J-2X rocket engine program. This paper describes the construction of TPMs and describes Utility Analysis. It then discusses the use of TPMs in design trades and to manage margin during a program using Utility Analysis.

Very high thrust-to-weight rocket engines

High delta-V earth-to-orbit missions have put a premium on high performance booster rocket engines. While significant improvements to specific impulse are unlikely, high thrust-to-weight design provides a promising avenue for improving mission and vehicle capabilities and margins. Several approaches can contribute to achieving such engine designs, including proper design optimization, simplification, geometry, propellant selection, and the application of advanced materials. Incorporation of the first four approaches can yield factors of about two improvements in current liquid engine designs. The utilization of emerging material capabilities could yield another factor of two improvement with the possibility of even larger gains with far-term materials and designs.