Overall Mass Properties

Table 5-4 provides overall vehicle mass properties for the ISS crewed variant of the CEV CM. The mass properties reporting standard is outlined in JSC-23303, Design Mass Properties. A detailed mass statement is provided in Appendix 5A, CEV Detailed Mass Breakdowns. Vehicle Mass Properties for the ISS Crewed Variant of the CEV CM Vehicle Mass Properties for the ISS Crewed Variant of the CEV CM

CEV Overview and Recommendations

One of the keys to enable a successful human space exploration program is the development and implementation of a vehicle capable of transporting and housing crew on Low Earth Orbit (LEO), lunar and Mars missions. A major portion of the Exploration Systems Architecture Study (ESAS) effort focused on the definition and design of the Crew Exploration Vehicle (CEV), the the fundamental element by which NASA plans to accomplish these mission objectives. This section provides a summary of the...

Launch Eorlor Architecture

The combination of advanced propulsion technology on the LSAM and CEV and additional ascent and injection mass performance with an upgraded CLV introduced another architecture variant in Design Cycle 2. This variant, known as 1.5-launch EOR-LOR, is so named due to the large difference in size and capability of the LVs used in the architecture. Whereas the previous architectures have used one heavy-lift Cargo Launch Vehicle (CaLV) to launch cargo elements and another heavy-lift Crew Launch...

LOXCH versus Storable Propellant Trades

Many NASA studies have evaluated propellant combinations of on-orbit propulsion for spacecraft. These include various combinations of Oxygen (O2), Hydrogen Peroxide (H2O2), Nitrous Oxide (N2O), Nitrogen Tetroxide (NTO), and Chlorine Pentafloride (ClF5), together with fuels such as Hydrazine (N2H4), Monomethyl Hydrazine (MMH), Hydrogen, Ethanol (EtOH), Methane (CH4), Propane (C2H6), and Kerosene (RP1). The propellants exhibiting the best overall characteristics from these studies are LO2 LH2,...

EORLOR with CEVto Surface Architecture

The fourth and final Analysis Cycle 1 architecture is a hybrid between the previous two options. It combines the LOR aspect of the EOR-LOR architecture and single crew volume of the EOR-direct return architecture. Rather than leaving the CEV CM and SM behind in LLO, this architecture separates the two elements, leaves only the SM behind, and uses the CM to operate on the lunar surface. However, as in the EOR-direct return architecture, the combined mass of the CEV and LSAM exceeded the...

Approach

The ESAS effort was organized around each of the four major points of the charter CEV definition, Launch Vehicle (LV) definition, lunar architecture definition, and technology plan definition. Additional key analysis support areas included cost, requirements, ground operations, mission operations, human systems, reliability, and safety. The ESAS team took on the task of developing new CEV requirements and a preferred configuration to meet those requirements. The CEV requirements developed by...

Load Magnitudes Analysis

During all phases of flight, it is mandatory that accelerations be kept within the crew load limits set forth by the NASA-STD-3000, Volume VIII, Human-Systems Integration Standards document. An example of these limit curves, which are a function of the duration of the load as well as the direction taken in the human body, is shown in Figure 5-15. Three limit-curves exist for each of the three human body axis directions. The highest limit-curve is intended for use in abort situations. It...

LV and EDS Performance System Trades

The options for LVs have become increasingly complex as technical strides are made in materials and systems design. The broad trade space currently available for ETO transportation for crew and cargo is shown in Figure 6-13. Figure 6-13. Possible Range of Launch Trade Study Rocket sled Electromagnetic Catapult Towed Non-Propellant Tanking During ascent Non-Propellant Tanking During ascent In order to arrive at a set of manageable trade options, an objective evaluation must consider the external...

Maximum Allowed Lailer Time In Lunar Orbit tor Earth Return days

The final part of the trade involved assessing the cost of implementing global access and anytime return. For LOR missions, the propulsive (or mass) cost is affected by adding propulsive maneuvers and fixing surface sorties at 7 days. For a fixed TLI of 3,150 m sec, a maximum three-impulse LOI of 1,390 m sec is required to achieve global access, which includes a worst-case nodal plane change of up to 90 deg. Departure is from a 407-km circular LEO parking orbit at 28.7 deg inclination. Earth...

Direct Entry Versus Skip Entry Comparison

This section will compare a lunar return direct-entry flight to a skip-entry flight. The vehicle used in this comparison will be an Apollo-style capsule with a ballistic number of 106 psf and a L D ratio of 0.3. The drag coefficient is 1.29. The entry speed will be 36,309 fps at an EI altitude of 400,000 ft. The flight-path angle for the direct-entry flight is -6.65 deg and -6.0 deg for the skip-entry. The difference in nominal flight-path angle at EI is the most distinct difference in the...

Docking MechanismISS Docking Module Trades

Docking Berthing Mechanism

As indicated in the President's Vision for Space Exploration, the completion of the ISS is a high priority for the Agency and the U.S. aerospace community. As such, CEV access to the ISS is of primary importance, and the mechanism and operations required for mating to the ISS must be factored into the CEV design and operations concept. Also, as stated in the Vision for Space Exploration, there is a need to develop systems and infrastructure that are enabling and allow for an affordable and...

Aerodynamic Stability Analysis

Another key analysis in the shape trade study involved assessing the inherent aerodynamic stability in the design of the CEV CM as it relates to vehicle shape and CG location. In the presence of an active control system, the natural behavior of a vehicle can be augmented. Still, it is important to design a vehicle that can operate in a passively stable configuration for worst-case situations. An understanding of the stability characteristics of a vehicle cannot be obtained from a single...

Return from Lunar Missions CEV Entry Trajectories Landing Mode Skip Entry Technique Description

The skip-entry lunar return technique provides an approach for returning crew to a single CONUS landing site anytime during a lunar month. This is opposed to the Apollo-style entry technique that would require water or land recovery over a wide range of latitudes, as explained in the following sections. This section will discuss the top-level details of this technique, as well as the major technological and vehicle system impacts. The skip-entry trajectory approach is not a new concept. The...

Axisymmetric Capsule Shape Variations and Effects

The basic capsule shapes shown in Figure 5-26 were analyzed using a modified Newtonian aerodynamics code. Various shape parameters, such as after-body cone angles, base radii, corner radii, heights, and others, were parametrically changed and evaluated in the aerodynamics generator to assess the effects of these parameters on the desirable criteria. Of primary interest were the sidewall angle theta , the corner radius Rc , and the base radius Rb . The data quickly indicated the desired path to...

Axisymmetric Capsule Shape Variations

One way to achieve the required L D is to use a nonaxisymmetric shape similar to the AFE shape mentioned previously. A computer-generated shape optimization approach was pursued to attempt to optimize an OML that exhibited some of the desirable characteristics without necessarily being axissymmetric. The investigation of various optimized shapes used the optimization capabilities of the CBAERO computer code. These optimized shapes held the aft-body shape fixed, while the heat shield shape was...

Blunt Bodies Versus Slender Bodies Trade

Biconic Capsule

The shape study trade was initiated between major vehicle classes. The primary classes considered were capsules blunt bodies , slender bodies, lifting bodies, and winged vehicles. Winged bodies and lifting bodies such as X-38, X-24, HL-10, etc. were eliminated at the outset due to several factors, including 1 the extreme heating especially on empennages these would encounter on lunar return entries, 2 the additional development time required due to multiple control surfaces, and 3 the increased...

Heat Transfer Analysis and TPS Sizing

The TPS sizing analysis was conducted using a transient 1-D Plug model. The required TPS insulation thicknesses were computed by a TPS Sizer using the Systems Improved Numerical Differencing Analyzer SINDA Fluid Integrator FLUINT software solver for the reusable concepts and the FIAT software code for the ablative TPS materials. For the aft portion of the capsule, a full soak-out condition was imposed for TPS insulation sizing. Because the heat shield for all capsule configurations was assumed...

Esas Lsam Configuration

Based on the results of the LSAM configuration trade studies, a combined CM design was chosen as a POD for future lander design studies. This concept was chosen to both provide the required airlock function and simplify ascent and descent stage interfaces. The description and mass property breakouts presented in this section are the result of additional lander analysis and refined subsystem mass estimation using the Envision sizing tool. It was recognized, however, that returning this large...

Crew Launch Vehicle

An array of options was assessed to determine their individual abilities to meet the stated requirements for the CLV. Those that most closely support the necessary demands are provided here. The remaining CLV options that were not evaluated further are discussed in more detail in Appendix 6A, Launch Vehicle Summary. Table 6-3. Shuttle-Derived CLV Options Assessed in Detail

LSAM Reference Design

The ESAS team examined the unique architecture of the lunar lander, or LSAM. Other architecture element designs and trade studies were also accomplished by the team. The reference LSAM concept, shown in Figure 1-16, for the ESAS 1.5-launch EOR-LOR architecture is a two-stage, single-cabin lander similar in form and function to the Apollo LM. The LSAM ascent stage, in conjunction with the descent stage, is capable of supporting four crew members for 7 days on the lunar surface and transporting...

Lunar Mission Mode

The lunar mission mode is the fundamental lunar architecture decision that defines where space flight elements come together and what functions each of these elements perform. Mission mode analysis had its genesis early in the design of the Apollo Program, with notable NASA engineers and managers such as Wernher Von Braun, John Houbolt, Joe Shea, and Robert Seamans contributing to the decision to use LOR as the Apollo mission mode. This study built on the foundation of the Apollo decision but...

Ascent Stage Description

The reference LSAM concept for the ESAS 1.5-launch EOR-LOR architecture is a two-stage, single-cabin lander similar in form and function to the Apollo LM. The LSAM ascent stage, in conjunction with the descent stage, is capable of supporting four crew members for 7 days on the lunar surface and transporting the crew from the surface to lunar orbit. The ascent stage assumes an integrated pressure-fed oxygen methane propulsion system, similar to the CEV SM, to perform coplanar ascent to a 100-km...

Lunar Cev Sm Vehicle Description

Electromagnetic Docking Mechanism

The Lunar CEV SM is included in the ESAS exploration architecture to provide major trans-lational maneuvering capability, power generation, and heat rejection for the CEV CM. The SM assumes an integrated pressure-fed oxygen methane service propulsion system and RCS to perform rendezvous and docking with the LSAM in Earth orbit, any contingency plane changes needed prior to lunar ascent, TEI, and self-disposal following separation from the CM. One 66.7 kN 15,000 lbf service propulsion system and...

Overview

Eelv Payload Comparison

A safe, reliable means of human access to space is required after the Space Shuttle is retired in 2010. As early as the mid-2010s, a heavy-lift cargo capability will be required, in addition to the crew launch capability to support manned lunar missions and follow-on missions to Mars. It is anticipated that robotic exploration beyond Earth orbit will have an annual manifest of five to eight spacecraft. The ESAS team was chartered to develop and assess viable launch system configurations for a...

CM Net Habitable Volume Trades

In the history of human spacecraft design, the volume allocated for crew operations and habit-ability has typically been the remaining excess after all of the LV constraints and vehicle design, weight, CG, and systems requirements were met. As a result, crew operability has often been compromised as crew sizes are increased, mission needs changed, and new program requirements implemented. CM habitability considerations have often been relegated to a second level behind engineering convenience...

Entry Trajectory for Cev Cm Returning from ISS

An evaluation of the CEV returning from the ISS was conducted as part of the ESAS. A simplified CEV vehicle model was used in the 4-DOF Simulation and Optimization of Rocket Trajectories SORT . The vehicle model consisted of an L D of 0.4 which included constant lift-and-drag coefficients as well as a constant ballistic number throughout the entry. A complete list of the simplified CEV model can be seen in Table 5-20. All entry scenarios were flown assuming two entry techniques, guided and...

CEV Design Evolution

The design and shape of the CEV CM evolved in four design cycles throughout the study, beginning with an Apollo-derivative configuration 5 m in diameter and a sidewall angle of 30 deg. This configuration provided an OML volume of 36.5 m3 and a pressurized volume of 22.3 m3. The CM also included 5 g cm2 of supplemental radiation protection on the cabin walls for the crew's protection. Layouts for a crew of six and the associated equipment and stowage were very constrained and left very little...

Lunar return heating is extremely high heavy heat shield Blunt Bodies versus Slender Bodies Conclusions

To summarize the results, it appeared that the capsule configurations have more desirable features and fewer technical difficulties or uncertainties than the slender body class of vehicles. Because one of the primary drivers for the selection was the minimal time frame desired to produce and fly a vehicle, the blunt bodies had a definite advantage. All the human and robotic experience NASA has had with blunt bodies has led to a wealth of knowledge about how to design, build, and fly these...